Design and Fatigue Analysis of a Typical Aircraft Wing fuselage Lug attachmen...
Damage Tolerance Evaluation of Wing Bottom Skin Panel by Analytical Approach
1. “DAMAGE TOLERANCE
EVALUATION OF WING BOTTOM
SKIN”
MOHAMMED ABDUL MEHBOOB (3KB06AE002)
ABDUL RAHIMAN GAZZALI (3KB08AE001)
MD. ZUNAID ALAM (3KB08AE018)
SHETTY SHISHIR SITARAM (3KB08AE013)
2. Aircraft are members and transverse frames to enable it
to resist bending, compressive and vehicles which are
able to fly by being supported by the air, or in general,
the atmosphere of a planet. An aircraft counters the
force of gravity by using either static lift or by using the
dynamic lift of an airfoil, or in a few cases the downward
thrust from jet engines. An aircraft is a complex
structure, but a very efficient man-made flying
machine.
4. Fuselage
The main body structure is the fuselage to which all other components are
attached. The fuselage contains the cockpit or flight deck, passenger
compartment and cargo compartment.
5. Wings
The wings are airfoils attached to each side of the
fuselage and are the main lifting surfaces that support
the airplane in flight. Wings vary in design depending
upon the aircraft type and its purpose.
6. The Empennage
The empennage is most commonly referred to as the tail of
the aircraft. It consists of two primary structures, the
vertical stabilizer and the horizontal stabilizer.
7. Landing gear
The landing gear is the principle support of the airplane when parked, taxiing,
taking off, or when landing. The most common type of landing gear consists of
wheels, but airplanes can also be equipped with floats for water operations, or
skis for landing on snow.
8. Control Surfaces
Following are the various Control Surfaces
The Flaps on the wings control the drag and lift on the
structure
The Rudder is used to change pitch (side-to-side) movement.
The Elevator is used to change pitch (up-down) movement.
The Aileron is uses to change lift, drag and roll.
The Slats are used to change lift.
Following are the various Control Surfaces:
The Flaps on the wings control the drag and lift on the
structure
The Rudder is used to change pitch (side-to-side) movement.
The Elevator is used to change pitch (up-down) movement.
The Aileron is uses to change lift, drag and roll. The Slats are
used to change life.
15. Damage Tolerance
Damage tolerance is a property of a structure
relating to its ability to sustain defects safely until
repair can be affected. The approach to
engineering design to account for damage
tolerance is based on the assumption that flaws
can exist in any structure and such flaws
propagate with usage.
Slow Crack Growth Design
Fail-Safe Design
18. Loads Acting on the panel.
The all-up weight of the aircraft (30-seater) is 15000 kg
Weight of the aircraft = 15000 kg.
Design load factor considered= 3g.
Total load acting on the aircraft = 15,000×3
= 45,000 kg
Factor of safety considered = 1.5
The design Ultimate load = 45,000×1.5
= 67,500 kg
Lift load experienced by both fuselage and wing.
Lift load on the wing = 80% of total load
= 0.8×67,500
= 54,000 kg
Load acting on each wing = 0.5×54,000
= 27,000 kg
20. Total span of the wing = 12000mm
Bending moment at the section A-A = 27,000×300
= 81×105 kg-mm
Depth of the wing at section A-A = 450 mm
The axial load in the bottom skin= 81×105/450
= 18,000 kg
22. Axial load acting on the edge of the panel = 18,000×1000/1500
= 12000 kgs
This 12000 kg load is converted into uniformly distributed load
(UDL) and applied at the one side of wing bottom skin cutout
panel.
UDL = 12000/width of bottom skin
= 12000/1000
UDL= 12 kg/mm
23. Material specification
The testing done on the materials are the linear elastic
ones. In material specification the kin d has to be specified.
The material may be hollow or a non hollow cross sectional
kind. In the thickness along with the aspect ratio has to be
specified.
24. Young’s Modulus = Modulus of elasticity = Stress /
Strain
For ALUMINIUM,
Young’s Modulus = 7000 kg/ mm².
The unit can be expressed as ‘ kg/ mm² ’.
Poisson’s Ratio = Lateral Strain/ Longitudinal Strain.
Poisson’s Ratio is dimensionless.
Poisson’s Ratio = 0.3 .
26. Steps involved in FEA
1. Discretization of the geometry into nodes and
elements.
2. Stiffness matrix is calculated for each element.
3. Assembly of stiffness matrix
4. For the equation F=K Q applying the boundary
conditions.
5. Solving the set of simultaneous equations using
different mathematical techniques.
6. Displacements are calculated first
7. Using displacements strains and stresses are
calculated
27. Finite Element Analysis software programs
Calculate element stresses and reaction forces from the displacement results.
Solve the matrix equation {F} = [K].{u} for displacements.
Apply loads to the model (Forces, moments, pressure, etc.)
Apply boundary conditions to constrain model.
Assemble all element stiffness matrices into a global stiffness matrix.
Formulate element stiffness matrices from element properties, geometry, and material properties .
Represent a continuous structure as a collection of grid points connected by discrete elements.
NASTRAN Solution Flow Chart
33. LOADS AND BOUNDARY CONDITIONS.
The load applied= 5000 kgs.
The span= 1000 mm.
Load/unit length = 5000/1000= 5 kg/mm
In this case while applying boundary conditions.
Stress (linear) = F/A.
= ( kg *9.81)/(mm^2).
34. STRESS ANALYSIS
The main steps employed in this are as follows in detail:
Geometry
Element Size
Material Specifications
Material Properties
Analysis
42. MODIFIED VIRTUAL CRACK CLOSURE INTEGRAL
In this paper, attention is focussed on a new predictor-
corrector method that results in an incremental
curved approximation of the crack path on the basis of
quantities which the straight forward MODIFIED
VIRTUAL CRACK CLOSURE INTEGRAL Method can
provide. In order to show significance of the proposed
simulation technique computational results are
compared with the findings from experimental
investigation with the aid of specially designed
specimens under proportional loading conditions, in
particular under combined proportional bending and
shear loading.
49. CONCLUDING REMARKS
Damage tolerance design philosophy is generally used in the
aircraft structural design to reduce the weight of the structure.
Stiffened panel is a generic structural element of the fuselage
structure. Therefore it is considered for the current study.
A FEM approach is followed for the stress analysis of the
stiffened panel.
The internal pressure is one of the main loads that the fuselage
needs to hold.
Stress analysis is carried out to identify the maximum tensile
stress location in the stiffened panel.
50. REFERENCES
T. Swift “Damage tolerance capability”, international journal of
fatigue, Volume 16, Issue 1, January 1994, Pages 75-94.
F. Erdogan and M. Ratwani, International journal of fracture
mechanics, Vol. 6,
No.4, December 1970.
H. Vlieger, 1973, “The residual strength characteristics of
stiffened panels containing fatigue crakes”, engineering
fracture mechanics, Vol. 5pp447-477, Pergamon press.
H. Vlieger, 1979, “Application of fracture mechanics of built up
structures”, NLR MP79044U.
