(unisim(behas)-introduction to aerospace)eas105 -lab3

Mohd Ashraf Mohd Ismail

Laboratory Experiment 3

Name : Mohammed Ashraf Bin Mohammed Ismail

Student No: N0806406

Contact No: 98225529

Date Submitted:

Lab. : Wind Tunnel Experiment

Course Instructor: Mr Roger Chua

2

Table of Contents

ABSTRACT .................................................................................................................. 3

INTRODUCTION ......................................................................................................... 4

OBJECTIVES................................................................................................................ 6

EXPIREMENT PROCEDURE ..................................................................................... 7

EXPIREMENT RESULT.............................................................................................. 9

Comparison between Theoretical and Experimental at 0° of flaps........................ 9

Comparison between 0° and 10° of flaps............................................................. 10

Comparison between 0°and 30° of flaps.............................................................. 11

DISCUSSION OF RESULT........................................................................................ 12

CONCLUSION............................................................................................................ 13

REFERENCE .............................................................................................................. 14

APPENDIX.................................................................................................................. 15

Introduction to Aerospace Engineering Lab 3

3

Abstract

In this experiment, we are out to show the relationship of the coefficient of lift and

drag in relation to the deployment of flaps (0, 10, 30). In addition to every change on

the angle of flaps, we also adjusted the angle of attack (AOA) (0,5,9,12,15,18) with

reference to the airflow. This show how it further affect the relationship how it affects

lift and drag. In this experiment we are using the aerofoil design of NACA 4412.

4

Introduction

Aircraft are supported in the air by an aerodynamic force called lift, which is

generated by the wings of the aircraft as air flows past the wings as a result of the

forward movement of the aircraft.

Many factor can affect the lift and drag component of the aerofoil which include

temperature, density, wing geometry, angle of attack(AOA) and angle of flaps

deployment and other factors. In this experiment we will concentrate on

Angle of Attack

Angle of Flaps Deployment - Flaps may be used to increase the maximum lift

coefficient, increase the wing area, or both. A change in the maximum lift coefficient

may be realized by a change in the shape of the airfoil section or by increased camber


5

The Lift Coefficient and the Drag Coefficient represent the changes in lift and drag as

the angle of attack changes. CL and CD are not expressed by any physical unit, they

are rather absolute numbers obtained from either wind tunnel tests or derived

mathematically.

Lift and Drag Formulas

6

Objectives

From the experiment we were able to :

I. Show the relationship of Coefficient of lift and drag with varying the angle of

flaps deployment.

II. Show the relationship of Coefficient of lift and drag with varying the angle of

attack( AOA) within the same flap deployment angle

III. Compare and calculate the difference between the experimental value and the

theoretical value for the coefficient of lift and drag (only at clean flaps)


7

Experimental Procedure

To observe, investigate and measure the lift and drag forces while varying the aerofoil

angle of attack, angle of flap and the test section velocity.

The type of aerofoil selected for the following experiment is NACA 4412 camber

aerofoil(Figure 1).

Figure 1

- Aerofoil Span b 300mm

- Aerofoil Chord length 100mm

- Average Sea level Temperature 23 ° c

- Average Sea level Pressure 1013mbar

- Velocity of airflow 16.9 m/s

- Air Density 1.19 kg/m³

8

Procedure:

1. Mount the aerofoil on the test section of the wind tunnel( figure 2).

2. Adjust the angle of flaps deployment first.

3. Adjust the Angle of attack (AOA) and then tighten he set screw with the Allen

wrench.

4. Monitor the Lift the Drag vales on the computer.

5. Repeat step 3 until the readings for all different angle of attack for that

particular flap deployment has been taken down.

6. Then repeat step 2 to adjust the angle of flaps deployment

7. Record and tabulate the result in a table form.

Figure 2


9

Experiment Result

Comparison between theoretical and experimental values at 0° flaps deployment

10

Comparison between 10° and 0° of flap deployment


11

Comparison between 30° and 0° of flap deployment

12

Discussion of Result

For the 1st lab comparison between the experimental value and the theoretical values(Both 0° flaps), the % error is quite big partly because NACA uses 54 pressure points and uses a 24’’ chord length to test on their aerofoil 4412 different from sample aerofoil. Another issue is that the environmental variable such as air pressure, air density, sea level temperature and velocity air speed are not the same. Other possible errors may be due to improper setup of experiment and the fluctuating airspeed readings

For the 2nd lab comparison between flaps deployment of 0° and 10°, the values are almost similar but at 10° flaps, the aerofoil starts to gain more lift at low AOA(0°-4°) but also loses lift faster at it’s stall angle(12°). For both value for drag the values are very similar

For the 3rd lab comparison between flaps deployment of 0° degrees and 30°, the difference are more distinct. At lower AOA airflow over the 30° flaps gain much more lift. The Cl values are in fact more that 2 times than clean flaps. The Cd is also slightly more than the clean flaps. At 30° flaps aerofoil start to stall at a much lower AOA compared to 0° flaps. The Cl starts to loses lift at 9° and the Cd starts to increase exponentially also at 9°.

I have a better understanding of lift and drag and finding the most optimum condition and applying it to the different phase in flying.

Take Off – Most aircraft would take off with Flaps 10° as it will give them the max lift and therefore BEST Climb Rate. It will also reduce the length of runway needed.

Cruising – At cruising you don’t need to climb but just to have the best something ratio. It you look properly at clearer research, the greatest difference between lift and drag is when aircraft is at AOA of 4° and no flaps. That’s why most aircraft wings are rigged at an angle of 4° (Angle of Incident). It has the least drag therefore aerodynamically it’s the most efficient condition to cruise.

Landing – At landing aircraft most aircraft want to descend gradually and land as slowly as possible (landing speed). Therefore we need more drag than lift but not till aircraft is stalled. Most of the time full flaps either 30 or 40° is being used for landing phase


13

Conclusion The experiment shows the relationship of lift and drag is affected by adjusting the angle of attack, angle of flaps deployment or even both. By completing the experiment

To sum it all up:

1) Different AOA for same flaps deployment:

a) Increase in AOA will result in higher lift than lower AOA but only before

stalling angle. Upon exceeding stalling angle, lift will decrease drastically

b) Drag will remain quite constant, only increasing slightly with increase in

AOA. Upon exceeding stalling angle, it will increase exponentially.

2) Different AOA at different flaps deployment:

a) Increase in lift initially with more deployment of angle of flaps. As same AOA

more lift will be generated with more deflection in flaps.

b) More angle of flaps will result in slightly increase in drag because of the

deflection in shape.

c) More angle of flaps will result the stalling angle to occur at lower AOA.

14

Reference 1. http://pilotsweb.com/principle/liftdrag.htm

2. http://classicairshows.com/Education/Aerodynamics/BernoulliAT1243

.htm

3. http://acam.ednet.ns.ca/curriculum/wing.htm

4. http://www.tpub.com/content/nasa1996/NASA-96-jcp-wka/NASA-96-

jcp-wka0009.htm

5. Theory of the Wing Section by IRA H.Abbot and Albert E. Von

Doenhoff. Published by Dover Publication. 1st publication in 1959

6. http://www.aerolab.com/

7.


15

Appendix I http://www.google.com/imgres?imgurl=http://mshades.free.fr/flapping/Cx4412.jpg&imgrefurl=http://

mshades.free.fr/flapping/selfincidentwingsection.html&h=557&w=324&sz=58&tbnid=DinGwPpi10U

J::&tbnh=133&tbnw=77&prev=/images%3Fq%3Dpicture%2Bof%2BNaca%2B4412&usg=__3dsax-

vRhkwdahNJ86899t8_3pE=&sa=X&oi=image_result&resnum=1&ct=image&cd=1


17

NACA 4412 (Stations and ordinates given in per cent of airfoil chord)

Upper Surface Lower Surface Station Ordinate Station Ordinate

0.0000 0.0000 0.0000 0.0000 0.5000 1.6549 0.5000 -0.8857 0.7500 1.9551 0.7500 -1.1096 1.2500 2.4478 1.2500 -1.4338 2.5000 3.3829 2.5000 -1.9484 5.0000 4.7302 5.0000 -2.4834 7.5000 5.7597 7.5000 -2.7429

10.0000 6.5986 10.0000 -2.8638 15.0000 7.8878 15.0000 -2.8791 20.0000 8.7963 20.0000 -2.7320 25.0000 9.4055 25.0000 -2.5089 30.0000 9.7589 30.0000 -2.2595 35.0000 9.8849 35.0000 -2.0162 40.0000 9.8030 40.0000 -1.8030 45.0000 9.5563 45.0000 -1.6058 50.0000 9.1916 50.0000 -1.3990 55.0000 8.7174 55.0000 -1.1930 60.0000 8.1404 60.0000 -0.9955 65.0000 7.4644 65.0000 -0.8124 70.0000 6.6958 70.0000 -0.6483 75.0000 5.8340 75.0000 -0.5054 80.0000 4.8817 80.0000 -0.3855 85.0000 3.8369 85.0000 -0.2888 90.0000 2.7001 90.0000 -0.2146 95.0000 1.4642 95.0000 -0.1609

100.0000 0.1302 100.0000 -0.1248 L.E. radius = 1.587 percent c slope of mean line at LE = 0.2000

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NACA 4412 (Stations and ordinates given in per cent of airfoil chord)

Upper Surface Lower Surface Station Ordinate Station Ordinate

0.0000 0.0000 0.0000 0.0000 0.2634 1.2975 0.7366 -1.0988 0.4641 1.6057 1.0359 -1.3085 0.8898 2.1054 1.6102 -1.6132 2.0181 3.0543 2.9819 -2.0856 4.3872 4.4390 5.6128 -2.5640 6.8264 5.5049 8.1736 -2.7862 9.3054 6.3810 10.6946 -2.8810

14.3370 7.7414 15.6630 -2.8664 19.4291 8.7091 20.5709 -2.7091 24.5557 9.3621 25.4443 -2.4871 29.7003 9.7442 30.2997 -2.2442 34.8513 9.8843 35.1487 -2.0093 40.0000 9.8030 40.0000 -1.8030 45.0620 9.5526 44.9380 -1.6082 50.1176 9.1816 49.8824 -1.4038 55.1650 8.6997 54.8350 -1.1997 60.2026 8.1144 59.7974 -1.0033 65.2292 7.4317 64.7708 -0.8206 70.2437 6.6558 69.7563 -0.6558 75.2451 5.7897 74.7549 -0.5119 80.2323 4.8350 79.7678 -0.3906 85.2042 3.7924 84.7958 -0.2924 90.1599 2.6611 89.8401 -0.2166 95.0979 1.4395 94.9021 -0.1617

100.0167 0.1249 99.9833 -0.1249 L.E. radius = 1.587 percent c slope of mean line at LE = 0.2000 http://www.pdas.com/sections45.htm#4412


19

Aluminum Alloy 5052

Available

Shapes

Typical

Chemistry

Characteristcs

Typical

Applications

Mechanical

Properties

Fabrication Guide

Available Shapes

5052 is available in Coil, Plate and Sheet.

- Top -

Typical Chemistry (% Maximum unless shown as a range)

Cu Si + Fe Mn Mg Zn Cr Al

0.10 0.45 0.10 2.2 / 2.8 0.10 0.15 / 0.35 Balance

- Top -

Characteristics

5052 is one of the higher strength non-heat-treatable alloys. It has a high fatigue strength and is a good choice for structures subjected to excessive vibration. The alloy has excellent corrosion resistance, particularly in marine atmospheres. The formability of the grade is excellent and in the annealed condition it offers higher strengths than 1100 or 3003 grades.

- Top -

Typical Applications

20

5052 is often used in high strength sheet metal work, marine components, appliances, fuel and oil tubing.

- Top -

Mechanical Properties

Tensile Strength Yield Strength

Elongation

Brinell Hardness

ksi MPa ksi MPa % in 2" (50mm)

5052-0 28.0 196 13.0 91 25 47

5052-H32 33.0 231 28.0 196 12 60

5052-H34 38.0 266 31.0 217 10 68

- Top -

Fabrication Guide

Weldability

Corrosion

Resistance

Formability

Machinability Mpa TIG Resist.

5052-0 A A D A A B

5052- A B C A A A


21

H14

5052-H18

A B C A A A

Aluminum 2024-O

Subcategory: 2000 Series Aluminum Alloy; Aluminum Alloy; Metal; Nonferrous

Metal

Close Analogs:

Composition Notes:

A Zr + Ti limit of 0.20 percent maximum may be used with this alloy designation for

extruded and forged products only, but only when the supplier or producer and the

purchaser have mutually so agreed. Agreement may be indicated, for example, by

reference to a standard, by letter, by order note, or other means which allow the Zr +

Ti limit.

Aluminum content reported is calculated as remainder.

Composition information provided by the Aluminum Association and is not for

design.

Key Words: Aluminium 2024-O; UNS A92024; ISO AlCu4Mg1; NF A-U4G1

(France); DIN AlCuMg2; AA2024-O, ASME SB211; CSA CG42 (Canada)

Component Wt. %

Al 90.7 - 94.7

Cr Max 0.1

Cu 3.8 - 4.9

Fe Max 0.5

Component Wt. %

22

Mg 1.2 - 1.8

Mn 0.3 - 0.9

Other, each Max 0.05

Other, total Max 0.15

Component Wt. %

Si Max 0.5

Ti Max 0.15

Zn Max 0.25

Material Notes:

General 2024 characteristics and uses (from Alcoa): Good machinability and surface

finish capabilities. A high strength material of adequate workability. Has largely

superceded 2017 for structural applications. Use of 2024-O not recommended unless

subsequently heat treated.

Uses: Aircraft fittings, gears and shafts, bolts, clock parts, computer parts, couplings,

fuse parts, hydraulic valve bodies, missile parts, munitions, nuts, pistons, rectifier

parts, worm gears, fastening devices, veterinary and orthopedic equipment, structures.

Data points with the AA note have been provided by the Aluminum Association, Inc.

and are NOT FOR DESIGN.

Physical Properties Metric English Comments

Density 2.78 g/cc 0.1 lb/in³ AA; Typical


Hardness, Brinell 47 47 AA; Typical; 500 g load; 10 mm ball

Ultimate Tensile Strength 186 MPa 27000 psi AA; Typical

Tensile Yield Strength 75.8 MPa 11000 psi AA; Typical

Elongation at Break 20 % 20 % AA; Typical; 1/16 in. (1.6 mm) Thickness

Elongation at Break 22 % 22 % AA; Typical; 1/2 in. (12.7 mm) Diameter


23

Modulus of Elasticity 73.1 GPa 10600 ksi AA; Typical; Average of tension

and compression. Compression modulus is about 2% greater than tensile modulus.

Ultimate Bearing Strength 345 MPa 50000 psi Edge distance/pin

diameter = 2.0

Bearing Yield Strength 131 MPa 19000 psi Edge distance/pin

diameter = 2.0

Poisson's Ratio 0.33 0.33

Fatigue Strength 89.6 MPa 13000 psi AA; 500,000,000 cycles

completely reversed stress; RR Moore machine/specimen

Machinability 30 % 30 % 0-100 Scale of Aluminum Alloys

Shear Modulus 28 GPa 4060 ksi

Shear Strength 124 MPa 18000 psi AA; Typical

Electrical Properties

Electrical Resistivity 3.49e-006 ohm-cm 3.49e-006 ohm-cm AA; Typical at

68°F

Thermal Properties

CTE, linear 68°F 23.2 µm/m-°C 12.9 µin/in-°F AA; Typical; Average over 68-

212°F range.

CTE, linear 250°C 24.7 µm/m-°C 13.7 µin/in-°F Average over the range 20-

300ºC

Specific Heat Capacity 0.875 J/g-°C 0.209 BTU/lb-°F

Thermal Conductivity 193 W/m-K 1340 BTU-in/hr-ft²-°F AA; Typical at

77°F

Melting Point 502 - 638 °C 935 - 1180 °F AA; Typical range based on typical

composition for wrought products 1/4 inch thickness or greater. Eutectic melting is

not eliminated by homogenization.

Solidus 502 °C 935 °F AA; Typical

Liquidus 638 °C 1180 °F AA; Typical

24

Processing Properties

Annealing Temperature 413 °C 775 °F

Solution Temperature 256 °C 493 °F

Aluminum 5052-O


Metal

Close Analogs:

Composition Notes:



design.

Key Words: UNS A95052; ISO AlMg2.5; Aluminium 5052-O; AA5052-O

Component Wt. %

Al 95.7 - 97.7

Cr 0.15 - 0.35

Cu Max 0.1

Fe Max 0.4

Component Wt. %

Mg 2.2 - 2.8

Mn Max 0.1


Component Wt. %


25


Si Max 0.25

Zn Max 0.1

Material Notes:








Tensile Yield Strength 89.6 MPa 13000 psi AA; Typical





Ultimate Bearing Strength 345 MPa 50000 psi Edge distance/pin

diameter = 2.0

Bearing Yield Strength 131 MPa 19000 psi Edge distance/pin

diameter = 2.0


Fatigue Strength 110 MPa 16000 psi AA; 500,000,000 cycles

completely reversed stress; RR Moore machine/specimen

Machinability 30 % 30 % 0-100 Scale of Aluminum Alloys

Shear Modulus 25.9 GPa 3760 ksi



26

Electrical Resistivity 4.99e-006 ohm-cm 4.99e-006 ohm-cm AA; Typical at

68°F

Thermal Properties


212°F range.

CTE, linear 250°C 25.7 µm/m-°C 14.3 µin/in-°F Average over the range 20-

300ºC

Specific Heat Capacity 0.88 J/g-°C 0.21 BTU/lb-°F Estimated from

trends in similar Al alloys.

Thermal Conductivity 138 W/m-K 960 BTU-in/hr-ft²-°F AA; Typical at 77°F

Melting Point 607 - 649 °C 1125 - 1200 °F AA; Typical range based on

typical composition for wrought products 1/4 inch thickness or greater




Annealing Temperature 343 °C 650 °F holding at temperature not required

Hot-Working Temperature 260 - 510 °C 500 - 950 °F


27

Aluminum 7075-O


Metal

Close Analogs:

Composition Notes:

A Zr + Ti limit of 0.25 percent maximum may be used with this alloy designation for

extruded and forged products only, but only when the supplier or producer and the

purchaser have mutually so agreed. Agreement may be indicated, for example, by

reference to a standard, by letter, by order note, or other means which allow the Zr +

Ti limit.



design.

Key Words: UNS A97075; ISO AlZn5.5MgCu(A); Aluminium 7075-O; AA7075-O

Component Wt. %

Al 87.1 - 91.4

Cr 0.18 - 0.28

Cu 1.2 - 2

Fe Max 0.5

Component Wt. %

Mg 2.1 - 2.9

Mn Max 0.3



28

Component Wt. %

Si Max 0.4

Ti Max 0.2

Zn 5.1 - 6.1

Material Notes:

General 7075 characteristics and uses (from Alcoa): Very high strength material used

for highly stressed structural parts. The T7351 temper offers improved stress-

corrosion cracking resistance.

Uses: Aircraft fittings, gears and shafts, fuse parts, meter shafts and gears, missile

parts, regulating valve parts, worm gears, keys, aircraft, aerospace and defense

applications.







Hardness, Knoop 80 80 Converted from Brinell Hardness Value

Hardness, Vickers 68 68 Converted from Brinell Hardness Value


Tensile Yield Strength 103 MPa 15000 psi AA; Typical






Shear Modulus 26.9 GPa 3900 ksi


29



Electrical Resistivity 3.8e-006 ohm-cm 3.8e-006 ohm-cm

Thermal Properties


212°F range.

CTE, linear 250°C 25.2 µm/m-°C 14 µin/in-°F Average over the range 20-

300ºC

Specific Heat Capacity 0.96 J/g-°C 0.229 BTU/lb-°F

Thermal Conductivity 173 W/m-K 1200 BTU-in/hr-ft²-°F

Melting Point 477 - 635 °C 890 - 1175 °F AA; Typical range based on typical

composition for wrought products 1/4 inch thickness or greater. Homogenization may

raise eutectic melting temperature 20-40°F but usually does not eliminate eutectic

melting.




Annealing Temperature 413 °C 775 °F

Solution Temperature 466 - 482 °C 870 - 900 °F

(unisim(behas)-introduction to aerospace)eas105 -lab3

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