Chapter 7.2 Composites

Question 1

composite material

Answer 1

consists of two or more materials producing together desirable properties that cannot be achieved by any of the constituents alone

Question 2

purpose of composites

Answer 2

higher stiffness, strength fatigue life, wear resistance and corrosion resistance, reduced weight, less thermal expansion

Question 3

ply

Answer 3

the smallest unit of laminate

homogeneous, continuous, obeys Hooke’s law

but not isotropic!

Question 4

laminate assumptions

Answer 4

1. laminate thickness is very small compared to its other dimensions
2. plies of the laminate are perfectly bonded
3. lines perpendicular to the surface of the laminate remain straight and perpendicular to the surface after deformation
4. ply and laminate are linear elastic
5. through-the-thickness stresses and strains are negligible (2D stress state)

Question 5

A16 and A26

Answer 5

equal zero when stack is balanced

Question 6

B matrix

Answer 6

couples bending stresses with normal force

equals zero when stack is symmetric

Question 7

D matrix

Answer 7

bending stiffness matrix relating the curvature to the bending moments

Question 8

von Mises

Answer 8

is not adequate for brittle materials!

Question 9

Puck’s failure criterion

Answer 9

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

Question 10

action plane / fracture plane

Answer 10

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

Question 11

resistances

Answer 11

R||t&raquo_space; R||c because of microbuckling in compression which reduces the strength

R|_ t is very small, that’s why we have 90deg layers

Question 12

fiber fracture FF

Answer 12

simultaneous breaking of thousands of filaments

final failure

primarily caused by stresses acting parallel to fibers

Question 13

inter-fiber fracture IFF

Answer 13

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

Question 14

structure sizing

Answer 14

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

Question 15

no more than […] plies oriented in the same direction o top of each oter

Answer 15

4

Question 16

composite safety approaches

Answer 16

1. first ply failure (FPF): complete laminate is considered t fail when the first matrix crack or fiber breakage occurs (most conservative)
2. 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)
3. first fiber failure (FFF): complete laminate is considered to fail when the first fiber breakage occurs (matrix can fail, fiber cannot) (usual approach)

Question 17

damage tolerance evaluation

Answer 17

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

Question 18

damage tolerance in metallic structures

Answer 18

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

Question 19

damage tolerance in composite structures

Answer 19

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

Question 20

curve of residual compressive strength

Answer 20

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

Question 21

barely visible impact damage (BVID)

Answer 21

structure must withstand the ultimate loading throughout the lifespan of the aircraft

< 0.5 mm indentation

Question 22

visible impact damage (VID)

Answer 22

structure must withstand only the limit loading, and the damage must be repaired as quickly as possible

< 2 mm indentation

Question 23

obviously detectable damage

Answer 23

structure must withstand the flight loads, and the damage must be repaired

sometimes needs clearance (plane repair on spot)

> 2 mm indentation

Question 24

compression after impact (CAI) test

Answer 24

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

Question 25

knock-down factors

Answer 25

B-Value, impact and notch sensitivity, environment (hot/wet), required safety factor decrease the allowed strain level to around 33% of (reference) material values

Question 26

composite column buckling

Answer 26

differs from metallic column buckling due to no plasticity (no yield)

two cases for flange crippling: one-edge-free / no-edge-free

Question 27

composite panel buckling

Answer 27

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

Question 28

increasing critical buckling stress of a composite panel

Answer 28

1. decrease the width of the panel b
2. decrease the aspect ratio alfa
3. increase thickness t
4. “improve” stacking sequence