1) The document analyzes the structural response of an aircraft fuselage stiffened panel through finite element modeling and analysis. 2) Key aspects analyzed include modeling different stiffener configurations, determining stresses and deformations under various loads, and identifying critical locations. 3) Results show the panel with I-stringers has the least deformation. Buckling analysis identified a critical buckling stress. Stress near mouse holes increases with their height. Maximum stresses at rivet holes were below the material yield strength.

1. Finite Element Modeling and Analysis of
Practical Airframe Stiffened Panel
A Shiva Rama Krishna (07BME208)
Final Review
Internal Guide External Guide
Prof. Akash Mohanty K E Girish
Assistant Professor (S.G) (Director,BAIL)

2. INTRODUCTION
• A stiffened panel is a generic representative structural
element of an airframe structure.
• Fuselage panels are commonly
fabricated as skin–stringer
constructions, which are permitted
to locally buckle under normal flight
loads.
• Basic stiffened panel of fuselage has:
-Bulk Head
-Stringer
-Skin
-Rivets

4. PROBLEM DEFINITION
• Analyzing the structural response of a stiffened panel in
various kinds of loading conditions.
• Identifying the finite element modeling techniques for
representing necessary structural details of stiffened
panel.
• Structural response of stiffened panel with various
stiffeners.
• Determining the stress values at
rivet location.
• Identification of critical locations.

5. LITERATURE REVIEW
C. Lynch presented a finite element modelling procedure for the post buckling
analysis of conventional riveted fuselage panels. He compared results from
different modeling approaches with test results.
Finite Element Modeling techniques:

6. STIFFENED PANEL ANALYSIS:
Most of the literature present on stiffened panel has focus towards buckling
and post buckling behavior.
Xiaozhi WANG studied the buckling and ultimate strength of aluminum plates
and stiffened panels.
Eduard Riks analysed the Buckling and post-buckling behaviour of stiffened
panels in wing box structures using finite strip method.

7. Gary A. Fleming carried out the modal analysis of stiffened panel using finite
element analysis and experimental techniques.
Dr. Edmond Chow compared the nonlinear finite
element analysis of wing box structure using NASTRAN with the actual test
results.
William L. Ko performed buckling analysis on a hat-stiffened panel subjected
to shear loading.

8. IDEALIZED STIFFENED PANEL
The curved
fuselage panel
is idealized as
a flat panel.

9. METHODOLOGY
• To study and calculate different kinds of
loads that act on the stiffened panel.
• To obtain the critical size of mouse hole in
bulkhead by analyzing the bulkhead web.
• Selecting appropriate finite element
modeling techniques for representing
necessary structural details.
• To simulate structural response for various
stiffeners using PATRAN/NASTRAN.
• Determining the stress values at rivet
location using the free body diagram of

11. LOADS ACTING ON THE FUSELAGE
Stresses generated in fuselage due to cabin
pressurization:
- Hoop stress
- Axial stress.
(source:ref.2)

12. •Stress due to bending of fuselage:
(source:ref.2)

13. MATERIAL PROPERTIES
 The Material being used is Aluminum 2024 alloy.
 Its composition is :
 Al-94%,Cu-4.4%,Mg-1.5%,Mn-0.6%
 Its properties are:
Young’s Modulus : 72 -75.69GPa
Tensile strength : 359 -510MPa
Elastic limit : 248 – 372MPa
(source:CES Edupack2005)

14. Calculations
2
Where,
P = cabin pressure = 6 psi = 0.00422 kg/mm2
r = radius of the fuselage = 59 inches = 1498.6 mm
t = thickness of the Stiffened plate = 1.75mm
Hoop stress = (0.00422 x 1498.6)/1.75 = 3.6137 kg/mm2
We know,
Force on the plate (F) = (Hoop stress x Area of cross-section)
F =3.6137 x (1500 x 1.75) = 9486.138kg
Load per unit length = 9486.138/1500 = 6.324kg/mm
Hoop stress = P * r/t

15. 2
Longitudinal stress = 3.6137/2 = 1.80685kg/mm2
Load on the plate = (Longitudinal stress x Area of cross-section)
Area of cross-section = (900 x 1.75) mm
Load on the plate = 1.80685 x 900 x 1.75 = 2845.788 kg/mm
Load per unit length = 2845.788/900 = 3.16198 kg/mm
Longitudinal stress= P * r/2t

16. •Buckling stress
1
Where,
Kc = Buckling Co-efficient which is obtained from graph shown below = 6.5
E = Modulus of Elasticity = 72GPa = 72 x 103N/mm2 (Al 2024 alloy)
µ = Poisson’s ratio =0.33
b = short dimension of the plate or loaded edge = 150mm
t = plate thickness = 1.75
σbuckling= (π2 kc E t2)/(12(1-µ2)b2)

17. 1 Graph : (b/t) v/s Kc

18. σbuckling
= (3.142 x 72 x 103 x 1.752 x 6.5)/(12(1-(0.33)2 x 1502)
= 58.734 N/mm2
= 5.8734 kg/mm2
Force per unit length due to compression = 10.27845kg/mm

19. GEOMETRY
 Stiffened Panel.PDF

20. ASSUMPTIONS FOR ANALYSIS OF FUSELAGE
STRUCTURE
 Skin: All skin panels are in a state of plane
stress(membrane behavior).
 Bulkheads: They work in orthogonal plane.
 Stiffeners: They resist body bending in
tension and compression.
(source: ref.4)

21. RESULTS
Von Mises stress of L-Stringer Stiffened Panel

22. Biaxial loading for C stringer stiffened panel

23. Von Mises stress of I-Stringer Stiffened Panel

24. Deformation of L-Stringer Stiffened Panel

25. Deformation of C-Stringer Stiffened panel

26. Deformation of I-Stringer Stiffened Panel

27.  Bi-axial loading

28. The above graphs compares the maximum deformation for stiffened panels with L,
C, I stringers for which the loads are applied as calculated. Significant difference in
the maximum deformation is noted where as the maximum stress value is observed
to be same for all the stiffened panel as the stringers do not contribute in taking the
hoop stress.

29. Stress distribution at Mouse Hole for L stringer

30. •Buckling of stiffened panel
The buckling stress for flat stiffened panel as proposed by
Bruhn1 is calculated and the structural response in
buckling is analysed.
Buckled shape for stiffened panel

31. Deformation due to buckling for L stringer stiffened panel

32. •Maximum stress induced in bulkhead
Maximum stress induced in the bulkhead is obtained and
graph is plotted to find the critical mouse hole size.

33. 0
5
10
15
20
25
30
35
23.6 25.8 27.6 31.1 34.7 39.1
Mouse hole height Vs Maximum stress

34. •Maximum Stress at Rivet Locations
•The Maximum stress value at the rivet location is found to be less than the yield strength
of the material.
Maximum Stress at Rivet hole

35. CONCLUSION
 Maximum deformation for stiffened panels with L, C, I stringers have been
determined. Significant difference in the maximum deformation is noted.
 The maximum stress value is observed to be same for all the stiffened panels as
the stringers do not contribute in taking the hoop stress.
 It is observed that the stiffened panel with I stringer has least maximum
deformation than other types of stringers considered.
 Buckling analysis has been performed for critical buckling stress as proposed by
Bruhn[1].

36. • It is also inferred that stress induced near Mouse Holes increases with increase
its height. As stress increases there is a point where it will be more than the
yield strength of the material and the critical size of the mouse hole is the size at
which the maximum stress is equal to the yield strength of material.
•The maximum stress at the rivet hole has been found to be less than the yield
strength of the material from the local analysis of rivet joint.

37. [1] Bruhn, Analysis and Design of Flight Vehicles Structures.
[2] Robert L. Mott, Applied Strength of Materials, 2002, Prentice Hall of India.
[3] C. Lynch, A. Murphy, M. Price, A. Gibson. The computational post buckling
analysis of fuselage stiffened panels loaded in compression. Thin-Walled
Structures, 42 (2004) 1445–1464.
[4] BOEING DESIGN MANUAL FE Modelling Guide for Aircraft Structural
Analysis. 1989.
[5] Wilhelm Rust, M.Kracht. RECENT EXPERIENCES IN LOAD ANALYSIS
OF AIRCRAFT FUSELAGE PANELS
REFERENCES

38. •Partial amount of this work has been communicated for a National Seminar on
Aerospace structures(NASAS) in IIT Kanpur