Chapter 7.2 Composites Flashcards
composite material
consists of two or more materials producing together desirable properties that cannot be achieved by any of the constituents alone
purpose of composites
higher stiffness, strength fatigue life, wear resistance and corrosion resistance, reduced weight, less thermal expansion
ply
the smallest unit of laminate
homogeneous, continuous, obeys Hooke’s law
but not isotropic!
laminate assumptions
- laminate thickness is very small compared to its other dimensions
- plies of the laminate are perfectly bonded
- lines perpendicular to the surface of the laminate remain straight and perpendicular to the surface after deformation
- ply and laminate are linear elastic
- through-the-thickness stresses and strains are negligible (2D stress state)
A16 and A26
equal zero when stack is balanced
B matrix
couples bending stresses with normal force
equals zero when stack is symmetric
D matrix
bending stiffness matrix relating the curvature to the bending moments
von Mises
is not adequate for brittle materials!
Puck’s failure criterion
2 failure modes:
1. fiber fracture FF: simultaneous breaking of thousands of filaments in a ply (final failure)
2. inter-fiber fracture IFF (resin failure): macroscopic crack which runs parallel to the fibers and separates an isolated UD-layer into two pieces (not necessarily final failure)
action plane and fracture plane
action plane / fracture plane
action plane: plane with maximal loading with respect to a certain stress component
fracture plane: plane in which the fracture occurs
the fracture plane where IFF occurs is the action plane where the critical stress combination is applied
resistances
R||t»_space; R||c because of microbuckling in compression which reduces the strength
R|_ t is very small, that’s why we have 90deg layers
fiber fracture FF
simultaneous breaking of thousands of filaments
final failure
primarily caused by stresses acting parallel to fibers
inter-fiber fracture IFF
the interaction between normal and shear stresses
mode A: transverse tensile stress and/or longitudinal shear stress fracture
mode B: longitudinal shear stress fracture, while transverse stress is compressive
mode C: most dangerous, forbidden to tolerate, implies risk of delamination
structure sizing
tensile/compressive FF: increase thickness (nr of plies) of the corresponding ply direction
tensile IFF (mode A): increase thickness of perpendicular oriented plies
compressive IFF (mode B): increase thickness of +45/-45 oriented plies with respect to failed ply
compressive IFF (mode C): increase thickness of perpendicular oriented plies
no more than […] plies oriented in the same direction o top of each oter
4
composite safety approaches
- first ply failure (FPF): complete laminate is considered t fail when the first matrix crack or fiber breakage occurs (most conservative)
- last ply failure (LPF): complete laminate is considered to fail only when the very last ply fails (something fails, you degrade data, run the analysis again until last ply fails)
- first fiber failure (FFF): complete laminate is considered to fail when the first fiber breakage occurs (matrix can fail, fiber cannot) (usual approach)
damage tolerance evaluation
is intended to ensure that should serious fatigue, corrosion or accidental damage occur within the operational life of the airplane, the remaining structure can withstand reasonable loads without failure or excessive structural deformation until the damage is detected
damage tolerance in metallic structures
damage propagation within metallic structures is relatively slow and well controlled
it must be demonstrated that the structure’s residual strength will always be higher than limit loading
damage tolerance in composite structures
composite materials are almost insensitive to fatigue, but are very sensitive to impact
inspection intervals are based on environmental deterioration and accidental damage ratings
the loss of strength due to impact might reach 50-70% and it is hard to detect
the impact damage is often visible from the unimpacted site (which is often unaccessible)
compression strength is most affected by the impact damage, because of the delaminations
curve of residual compressive strength
defines the loads that a structure needs to withstand
three sizing areas:
1. barely visible damage (BVD)
2. visible impact damage (VIB)
3. obviously detectable damage
barely visible impact damage (BVID)
structure must withstand the ultimate loading throughout the lifespan of the aircraft
< 0.5 mm indentation
visible impact damage (VID)
structure must withstand only the limit loading, and the damage must be repaired as quickly as possible
< 2 mm indentation
obviously detectable damage
structure must withstand the flight loads, and the damage must be repaired
sometimes needs clearance (plane repair on spot)
> 2 mm indentation
compression after impact (CAI) test
test conducted in order to determine the residual strength
for a simple laminate, after the CAI test, we observe two types of matrix cracking:
1. vertical matrix cracks in the lower part of the laminate due to high transverse stress sig_t
2. matrix cracks at 45deg at the centre of the laminate under the impactor due to high out-of-plane shear stress
two damage phenomena leading to fracture are observed:
1. buckling of sub-laminates delaminated during impact (very bad)
2. propagation of compressive fibre fractures of plies oriented along the loading direction
test outcomes: longitudinal residual strain and shear residual stress
knock-down factors
B-Value, impact and notch sensitivity, environment (hot/wet), required safety factor decrease the allowed strain level to around 33% of (reference) material values
composite column buckling
differs from metallic column buckling due to no plasticity (no yield)
two cases for flange crippling: one-edge-free / no-edge-free
composite panel buckling
special case: D16 and D26 are negligible (special orthotropic)
boundary conditions:
1. biaxial loading, all edges simply supported
2. uniaxial loading, 3 edges SS and 1 edge free
3. shear loading, all edges simpliy supported
increasing critical buckling stress of a composite panel
- decrease the width of the panel b
- decrease the aspect ratio alfa
- increase thickness t
- “improve” stacking sequence
