Complete Design Review with proper analysis through the use of Solidworks.

1. OPTIMAL WING DESIGN OF
NACA 2412 SERIES
GROUP MEMBERS:
VINAY KHANDELWAL (16M7061)
RUCHIT MALIK (16M7042)
KANISHKA SHRIVASTAVA (16M7023)
SHIVAM SINGH BAGRI (16M7050)
GUIDED BY:
MR. OM PRAKASH SONDHIYA

2. OBJECTIVES
• STUDY OF NACA 2412 SERIES WING DESIGN.
• DESIGN NACA SERIES 2412 with all the specific
parameters.
• To successfully carry out static, dynamic and flow
stimulation.
• To find our own optimal design for the NACA 2412
SERIES. (We are presenting our own design that is
inspired from the same.)
• Analysis of structure, materials and geometry and
also the mathematical equations involved.
• After stimulations practical approach of design in the
real world.

3. READINGS GENERATED
• FIRST COLUMN SHOW X COORDINATE VALUE
• SECOND COLUMN SHOW Y COORDINATE
VALUE
• Z COORDINATE VALUE ASSUMED ZERO
• TAKEN DURIND EXTRUSION
REFERENCE - MEAN CAMBER LINE

4. DESIGNING OF PART IN
SOLIDWORKS (THE INITIAL
DESIGN)
BASIC DESIGN
PARAMETERS-
• LOAD – 20 KN
• PRESSURE – 2000 Mpa
• TORQUE – 20 KNm
DESIGN BASED ON THE
COORDINATE SYSTEM
USED IN SOLIDWORKS.
CALULATIONS DONE BY
USING THE EQUATIONS
PRESCRIBED BY NASA
( McGhee and Beasley
1973 )

5. MATERIAL SELECTION
MATERIAL SELECTED –
ALUMINIUM ALLOY 7079
• LIGHT WEIGHT
• EXTENSIVELY USED FOR
LIGHT AS WELL AS HEAVY
PAYLOAD
• PROPERTIES ARE SHOWN IN
TABLE

6. MATERIAL SELECTION
• The design and materials needs to satisfy two main
requirements:
1. Spar cross-section is not to be changed along the
wing.
2. First the wing handles a max point load of 445 N.
• Two main objectives in this selection are:
1. To keep the cost of the design as low as possible.
2. To make sure the best quality is being achieved .
• Main feature in this material selection are:
1.Long service time and long life.
2.Provide enough lift in all weather conditions

7. Material Density(g/cm^3) Young Modulus
(Gpa)
Yield Stress
(Mpa)
Cost(USD/KG)
6061 Al 2.70 69 55 2.53
CF Composite 1.60 134 131 55
Comparing to most suitable materials for
designing:
6061 Aluminum vs CF Composite
Young's modulus is important criteria in selection but not necessary one.
As of here, Cf Composite has higher Young’s Modulus but Cost is much higher
Hence by designing airfoil to bear suffering stress is desired.

8. Main question is whether the Aluminum can be able to handle a load of 445 N while having
deflection of less than 5%?
• Solution : The calculations are as follows:
• Length of the spar: L=1.8 (m)
• Width of cross-section: b=0.4 * 0.07= 0.028 (m)
• Height of cross-section: h=0.4*0.12 = 0.048 (m)
• Moment of Inertia: I== 1/12*0.028*(0.048)^3=2.58*10^-7 (m^4)
• Maximum Point Load: P=445 (N)
• Based on the following formulas for a cantilevered beam, the values for Slope, Deflection,
Moment and Shear force also can be calculated
• Distributed Load: w=P/L=445/1.8= 247.2 (N/m)
• Maximum deflection of Aluminum takes place at 0.018 (m) which is 1 percent of 1.8 (m)
(the length of spar)
• This shows satisfactory design requirements
• factor of safety 5 times more than the requested design.
• This way with6061 Aluminum $52.47 is saved in addition

9. STATIC SIMULATION ( VON MISES
STRESS READINGS )

10. RESULTS OF STATIC SIMULATION
• I
BY THE GRAPH
SHOWN WE INFER
THAT
1. STRESS EXCEED
THE PRESCRIBED
VALUE AT THE
TAIL PORTION
2. THESE EFFECTS
ARE GENERALLY
NEGLECTED
BECAUSE THAT
PORTION OF
WING IS
ASSEMBLED
WITH MAIN BODY
AND CAN BEAR
THE GIVEN
STRESS
3. SCOPE FOR
IMPROVEMENT
IN DESIGN

11. STATIC SIMULATION PART 2-
DISPLACEMENT READINGS
THE OBSERVED
DISPLACEMENT IS
UNDER THE
PRESCRIBED
VALUE UNDER
THE EFFECT OF
STATIC LOAD
APPROX
( 1 – 1.5 mm)

12. STATIC SIMULATION PART 3-
STRAIN READINGS
OBSERVATIONS-
AS SHOWN IN THE
FIGURE THE BLUE
REGION IMPLIES THE
STRAIN IS UNDER
CONTROLLABLE
LIMIT
( 7.713 * 10-9 –
1.722 * 10-7 )

13. THERMAL SIMULATION

14. OBSERVATIONS OF THERMAL
SIMULATION
• WORKING TEMPRATURE RANGE –
300 K – 500 K ( DUE TO STAGATION POINTS)
• DUE TO TEMPRATURE RISE THE STRESS
GENERATED ARE IN THE RANGE-( N/m2 )
( 4.654*102 - 1.628 *104 )
• MEAN TEMPRATURE RISE IN STUDY(FIG4) - 350 K
• THE STRESS GENERATED AT THE TAIL END NEED
TO BE REDUCED

15. Work Breakdown
• OBJECTIVES COVERED-
 STUDY OF GEOMETRY OF NACA2412 SERIES
 BASIC DESIGN FORMULATION ON SOLIDWORKS
 STATIC , THERMAL SIMULATION ON BASIC DESIGN DONE
 FLOW AND DYNAMIC SIMULATION ON WING
 ALTERNATE WING SELECTION
DECEMBER-JANUARY
 MODIFICATION OF GEOMETRY TO INCORPORATE THE
EXCEEDING STRESS VALUES
 PRATICAL DESIGN APPLICATION THEORY
 COST ANALYSIS IF THE DESIGNED PART IS MANUFACTURED

16. EXPECTED OUTCOMES
• Feasible Design that is more efficient and
effective.
• Optimal design of NACA Series 2412.
• Cost efficient model of NACA2412.

17. REFERENCES
•
• [1]- "GENERAL ALUMINUM INFORMATION from Aircraft Spruce Canada." GENERAL ALUMINUM
INFORMATION from Aircraft Spruce Canada. Aircraft Spruce and Specialty Co, 1995. Web. 18 Mar.
2016.
• [2]- Brandt, Steven A. "Introduction to Aeronautics." Google Books. N.p., n.d. Web. 18 Mar. 2016.
<https://books.google.ca/books?id=NJE3wTL6IDYC&pg=PA333&lpg=PA333&dq=cross%2Bsection%
2Bshapes%2Bof%2Bspar&source=bl&ots=Q-mz5I_qGf&sig=qj6Iaj1SUnGzjgqE2j-
tzT7oAxs&hl=en&sa=X&ved=0ahUKEwi-
w4m_9cjLAhVF1h4KHY_tAfUQ6AEIITAB#v=onepage&q=cross%20section%20shapes%20of%20spar
&f=false
• [3]- "Cantilever Beams." Cantilever Beams. N.p., n.d. Web. 17 Mar. 2016.
<http://www.engineeringtoolbox.com/cantilever-beams-d_1848.html>.
• [4]- Advanced Mechanical Engineering Solutions. (n.d.). Retrieved March 18, 2016, from
http://www.amesweb.info/StructuralBeamDeflection/CantileverBeamStressDeflectionCalculator.as
px>.
•
• [5]- Aluminum wing spar strength. (n.d.). Retrieved March 18, 2016, from
http://www.homebuiltairplanes.com/forums/aircraft-design-aerodynamics-new-technology/5472-
aluminum-wing-spar-strength-2.html
•