Transcript Design of a Composite Wing with Leading Edge Discontinuity Daniel Hult AerE 423 Project December 12, 2009

Design of a Composite Wing
with Leading Edge Discontinuity
Daniel Hult
AerE 423 Project
December 12, 2009
Overview
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Background
Project Goals
Design
Computational Analysis
Fabrication
Testing
Results & Conclusions
Future Work
Background
• Purpose
– Discontinuity causes vortex to form, keeping flow
attached to outer wing and ailerons
– Improved stability and performance at high α
– Spin prevention
Cirrus Aircraft Company
Project Goals
• Determine the structural feasibility of a
composite, single-piece wing with a
discontinuous leading edge.
• Design, build and structurally test a singlepiece composite wing.
Design
• Phases:
– Aerodynamic Analysis
– Structural Design
• Purpose of project is structural
• Aerodynamics only to get accurate loads
Aerodynamics
• XFLR5 Analysis
– Open source aerodynamics for R/C gliders
– Uses Vortex Lattice Method
– Allows low Reynolds Number analysis of any wing
Structural Design
• Laminate Study
– Analysis of laminate geometry with comp_core
– Varied combinations of 0/90 plies and ±45 plies
– Loading
• Tension and Bending
• Compression and Bending
– Laminates with more ±45 plies performed better
in bending
Structural Design
• Final Laminate
– 6 plies of 0.002 in. thick bi-weave fiberglass
– 4 plies at 0 and 90 degrees
– 2 plies at +45 and -45 degrees
Computational Analysis
• ANSYS used for Finite Element Analysis
• Three cases tested
– Isotropic material (aluminum)
– Graphite-Epoxy composite
– Fiberglass-Epoxy composite
• 300 N distributed load at tip
– Loading from XFLR5
– Depicted test to be performed
Computational Analysis
• Fiberglass
– Max Stress= 587 Mpa
– Max disp = 1.25 cm
Fabrication
• Mold
– Airfoil sections cut out of
particle board
– Used as stencils to hotwire blue
foam
– 2 sections joined and handle
added to root
Fabrication
Fabrication
• Lay-up
– Hand lay-up around mold
– Wrapped and cured with vacuum assistance.
Testing
• A Successful test would clearly accomplish
project goals
• Wing anchored at root with load applied at tip
• Load added to tip until failure
Testing
Screw Method
Clamp Method
Testing
• Wood mount failed along screws (20 lb)
• Fiberglass failed along clamped shims (40 lb)
Results & Conclusions
• Wing failed at clamped root at small load
• ANSYS predicted stress concentrations and
therefore failure at discontinuity
• The results were inconclusive, necessitating
further testing
Future Work
• Better fabrication techniques and materials
– Two-piece wing
– Carbon Fiber
– VARTM or Pre-Preg
• Better testing and mounting methods
– Metal or composite mounting plate and insert
– Metal or composite tip insert for loading
References
•
Abbott, Ira H. and Albert E. von Doenhoff. Theory of Wing Sections: Including a Summary of Airfoil Data.
New York: Dover Publications, Inc. c1959.
•
“CAPS™ and Stall/Spin.” Cirrus Aircraft Company. Accessed 12 October 2009.
<http://www.whycirrus.com/engineering/stall-spin.aspx>.
•
Deperrois, André. “About XFLR5 calculations and experimental measurements” August 2008.
<http://xflr5.sourceforge.net/xflr5.htm#_Help>.
•
Deperrois, André. “Guidelines for XFLR5 V4.16.” April 2009. <http://xflr5.sourceforge.net>.
•
Goyer, Robert. “Airplane on a Mission: Created for use in the humanitarian field, the Quest Kodiak
delivers raw utility at a great price.” Flying Magazine. February 2009
<http://www.flyingmag.com/turbine/1344/quest-kodiak-airplane-on-a-mission.htmlQuest Kodiak>.
•
“Kodiak Features.” Quest Aircraft Company. Accessed 29 September 2009.
http://www.questaircraft.com/index.php?filename=features.php
•
Meschia, Francesco. “Model analysis with XFLR5.” RC Soaring Digest. February 2008: p27-51.
•
NASA Langley Research Center. “Spin Resistance” Updated 17 October 2003.
<http://oea.larc.nasa.gov/PAIS/Concept2Reality/spin_resistance.html>.
Acknowledgments
• Dr. Vinay Dayal, Professor
• Chunbai Wang & Peter Hodgell, TA’s
• AerE 462 group, especially Robert Grandin for ideas and
support
• Iowa State University, Department of Aerospace Engineering
Questions?
Background
Airliners.net
• Uses
– Messerschmitt Bf-109
Airliners.net
– Large commercial jets
– NASA Spin Prevention Tests
– Cirrus SR20
– Quest Kodiak
Cirrus
Airliners.net
NASA Spin Prevention
Figures from NASA Langley report
Aerodynamics
• Airfoil Design
– NACA 2412 chosen for basis
– Common, well-known low-speed airfoil
– Discontinuity created by extending NACA 2412
Aerodynamics
• Wing Design
– Basic wing designed to be fabricated and tested
structurally
– NACA 2412 inner section (0.3 m)
– Modified airfoil outer section (0.2 m)
– b/2=0.5 m
– cr=0.25 m
Table A1: Test Laminates
Laminate Study
Test
T1
T2
T3
T4
A1
A2
A3
A4
Test
T1
T2
T3
T4
A1
A2
A3
A4
Ex (Gpa)
95.99
88.51
79.84
69.68
73.83
68.34
62.02
54.69
Material
T300/5208 Graphite-Epoxy
T300/5208 Graphite-Epoxy
T300/5208 Graphite-Epoxy
T300/5208 Graphite-Epoxy
AS/3501 Graphite-Epoxy
AS/3501 Graphite-Epoxy
AS/3501 Graphite-Epoxy
AS/3501 Graphite-Epoxy
Planar
Ey (Gpa) Gxy (Gpa)
95.99
7.17
88.51
13.74
79.84
20.31
69.68
26.88
73.83
6.9
68.34
11.68
62.02
16.47
54.69
21.25
Poisson
0.0302
0.1058
0.1934
0.296
0.0367
0.1085
0.1909
0.2865
Laminate
{[0,90];6}s
{[0.90];2,[45,-45];1,[0.90];3}s
{[0,45,90,0,-45,90];2}s
{[0,90,45,-45];3}s
{[0,90];6}s
{[0.90];2,[45,-45];1,[0.90];3}s
{[0,45,90,0,-45,90];2}s
{[0,90,45,-45];3}s
Ex (Gpa)
106.7
96.74
88.69
83.25
81.95
74.58
71.95
64.69
Bending
Ey (Gpa) Gxy (Gpa)
85.28
7.17
79.68
13.9
70.31
19.97
71.33
21.85
65.73
6.9
61.67
11.8
51.05
15.89
55.66
17.59
Poisson
0.0339
0.1187
0.2103
0.2333
0.0413
0.1214
0.2116
0.2286
Laminate Study
Test
T1
T2
T3
T4
A1
A2
A3
A4
Tensile
X (Mpa) Y (Mpa)
-7.59
39.9
-57.9
39.4
-95.1
38.7
-132
38.5
-9.75
48.1
-59.7
47
-95.2
45.6
-130
45.3
Z (MPa)
1.22E-08
-1.89
-6.33
-2.3
1.25E-08
-2.22
-7.68
-2.82
Compressive
X (Mpa) Y (Mpa) Z (Mpa)
-8.23
39.9
1.22E-08
-58.6
39.4
-1.89
-95.7
38.7
-6.33
-132
38.5
-2.3
-10.4
48.1
1.25E-08
-60.4
47
-2.21
-95.8
45.6
-7.67
-131
45.2
-2.82
Computational Analysis
• Isotropic
– Max Stress= 654 Mpa
– Max disp = 1.34 cm
Computational Analysis
• Carbon Fiber
– Max Stress= 600 Mpa
– Max disp = 1.25 cm
Mold
Finished Laminate
Testing