Transcript Finite Element Analysis of Bulging Factors of Aircraft

Foam Reinforced Aircraft
Fuselage Study
Narasimha Harindra Vedala, Tarek Lazghab, Amit Datye, K.H. Wu
Mechanical And Materials Engineering Department
Florida International University
Miami, Florida
Overview

Propose a strategy to reduce the effect of fatigue
failure

Simulate the failure behavior of fuselages due fatigue
using finite element analysis
Determine the effect of reinforcement on fatigue life due
to the new proposed strategy

Fatigue Failure
It is defined as the failure of a metal structure
due to cyclic loading.
Fatigue Failure occurs at load amplitudes
much lower than the breaking load of the
material
Fatigue Failure in Aircrafts
 Aging aircrafts fatigue analysis

Pressurization and Depressurisation of the
fuselage results in cyclic stresses
Aloha incident in 1988 caused due to fatigue failure
How to reduce the effect of fatigue failure?
Using composite materials for manufacturing
fuselages.
Using Reinforcement to the fuselage
Pantherskin
Research on polyisocyanurate foam at FIU since
1988. Nicknamed as Pantherskin
Why Pantherskin?
Inexpensive foam
 Light weight :Density =2 to 4 lb/ft3
 High fire resistance

Fracture in metals
Modes of Cracking in flat bodies
(based on Linear Elastic Fracture Mechanics)
Mode I
Mode II
Mode III
Objective
Study the effect of adding foam layer to stress
distribution in the fuselage
 Generate a Finite Element model to conduct
fatigue analysis on fuselage with foam
reinforcement
 Compare with results from previous methods

Fracture parameters
Stress Intensity Factors
It is defined a quantity that characterizes the
severity of the crack situation
For flat plate with infinite width

KI    a
Strain Energy Release Rate
K
G
E
2
E = Elastic modulus
2a
Verification study
Procedure Verification
Flat plate with central crack
Results were compared with
Rybicki’s (1977) model
2a
a
Percentage Error Vs Element size near the crack tip
8
Percentage error (%)
7
6
5
4
3
2
FEM G value compared to Theory
1
FEM G value to Rybickis FEM value
0
0
0 .1
0 .2
0 .3
0 .4
0 .5
0 .6
Element Size (in)
Monda y, J une 16, 2003
Finite Element Analysis
Model Definition
Aircraft fuselage can be considered as a cylinder
with equal spaced cracks
Considering
symmetric
geometry
a section of the
fuselage is used for
finite element analysis
Verification study : Unstiffened
curved panel
Procedure
Verification
Compared
with Young’s model
using STAGS FE
package
Longitudinal crack configuration
a
Bulging of crack in the fuselage
FEA of Aircraft Fuselage
Boeing B737
obtained
Configuration
R=74.018
t= 0.036
Aluminum
EA=10.5 Msi
Foam
EF =3000 psi
vA = 0.33
vF = 0.3
Components of fuselage
Finite Element Model
Boundary Conditions and
loading
p
Von Mises Stress
Stress distribution in the panel
when 2 foam is added
Pct Hoop Strain Reduction
30%
y = 0.0002x 3 - 0.0156x 2 + 0.1156x
Pct Hoop Strain Reduction
25%
20%
15%
10%
PctStr
Fit
5%
0%
0.00
0.50
1.00
1.50
2.00
2.50
Foam Thickness (in)
3.00
3.50
4.00
4.50
Percent Stress reduction vs. Foam Thickness
30%
y = 0.0003x 3 - 0.0156x 2 + 0.1159x
Pct Stress reduction
25%
20%
15%
10%
PctSig
Fit
5%
0%
0.00
0.50
1.00
1.50
2.00
2.50
Foam Thickness (in)
3.00
3.50
4.00
4.50
Fatigue Analysis
Coffin-Manson equation

2


b  0.071
 'f  0.18
  'f

 E

  0.008632

c  0.645
 'f
2N 

f
E
b
Strain range
 
'
  2N
f
f
c
Nf
Cycles until failure.
Slope of elastic strain amplitude vs. fatigue life.
One reversal intercept of plastic strain vs. life line.
On reversal intercept of elastic strain vs. life line.
Slope of plastic strain amplitude vs. fatigue life
Fatigue Analysis
Fuselage panel with Longitudinal Crack 4.5
Global model and Submodel
A submodel is generated to study the stress
distribution near the crack
Results for the global model are used for the nodes
at the boundary of the submodel
Location of submodel in the global fuselage panel
Von mises stress Distribution in
Submodel
Fatigue Life increase Vs Foam Thickness
y = 0.0446x4 - 0.3775x3 + 0.8904x2 + 0.3025x + 1
4
3.5
3
Fold Increase
2.5
2
1.5
1
Fold Increase in fatigue life
Poly. (Fold Increase in fatigue life)
0.5
0
0
0.5
1
1.5
2
2.5
Foam Thickness(in)
3
3.5
4
4.5
Conclusion
Reinforcement of the fuselage skin with foam
help to reduce the stresses distribution
18% for 2 " foam layer
25% for 4 " foam layer
 A maximum of 32% reduction in stress intensity
factor is possible for 4" foam addition
 A 3.7 fold increase in life of the aircraft is
achievable if 4" foam is used
 A parametric is developed which can be used for
different fuselage configurations

References
•Chen, D.,(1991), “Bulging of Fatigue Cracks in a Pressurized Aircraft Fuselage,” Ph. D.
Dissertation, Department of Aerospace Engineering. Delft, The Netherlands: Delft
University of Technology
•Rybicki, E.F., Kannien, M.F.,(1977) “A Finite Element calculation of stress intensity
factors by a modified crack closure integral”. Engineering Fracture Mechanics, Vol. 9, pp.
931-938
•Ahmed, A.A, Backukas, J., Jr. and Paul W. Tan, Awerbuch J., (2002) “Initiation and
distribution of multiple-site damage (MSD) in fuselage lap joint curved panel”, Sixth Joint
FAA/DoD/NASA Conference of Aging Aircrafts, Albuquerque, San Francisco, CA.
•Torres, M.J. and Wu, K.H., (1993), “A New Approach to solve aging airplane problems
using Polyisocyanurate”, Journal of Cellular Plastics, Vol. 29, N4, p.p 380.
•Rahman, A., Backukas J.G., Jr., Paul W. Tan, and Catherine A. Bigelow,(2002), “Bulging
Effects on longitudinal cracks in lap joints of pressurized aircraft fuselage”, Sixth Joint
FAA/DoD/NASA Conference of Aging Aircrafts, Albuquerque, San Francisco, CA.
Future Research
Several modeling issues were encountered
 Numerical formulation for defining fracture
parameters for composite materials is not yet
available
 Results have to be verified by experimental tests

Questions