51 Structure

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SINGLE AISLE TECHNICAL TRAINING MANUAL

MAINTENANCE COURSE - T1 (V2500-A5/ME) STRUCTURE



This document must be used for training purposes only

Under no circumstances should this document be used as a reference

It will not be updated.

All rights reserved

No part of this manual may be reproduced in any form,

by photostat, microfilm, retrieval system, or any other means,

without the prior written permission of AIRBUS S.A.S.



STRUCTUREStructure General (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

Doors D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18

Fuselage D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34

A318 Fuselage D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68Pylons/Nacelles D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100

Stabilizers D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 126

A318 Stabilizers D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160

Windows D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 194

Wings D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208

Structure Damage Identification D/O (3) . . . . . . . . . . . . . . . . . . . . . 256

Window Damage Identification D/O (3) . . . . . . . . . . . . . . . . . . . . . 288

SRM D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 298

Window Damage Assessment D/O (3) . . . . . . . . . . . . . . . . . . . . . . . 438

Damage Assessment Example 1 D/O (3) . . . . . . . . . . . . . . . . . . . . . 454

Damage Assessment Example 2 D/O (3) . . . . . . . . . . . . . . . . . . . . . 518

A318 Damage Assessment Example 3 D/O (3) . . . . . . . . . . . . . . . . 582

Structure Protections & Awareness D/O (3) . . . . . . . . . . . . . . . . . . . 654

Damage Assessment Ex. 1 Operational Scenario (3) . . . . . . . . . . . . 686

Damage Assessment Ex. 2 Operational Scenario (3) . . . . . . . . . . . . 694

A318 Damage Assessment Ex. 3 Operat. Scenario(3) . . . . . . . . . . . 702

MAINTENANCE COURSE - T1 (V2500-A5/ME)

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TABLE OF CONTENTS May 11, 2006

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STRUCTURE GENERAL (3)

AIRCRAFT MATERIALS

METALLIC MATERIALS

The basic A/C structure is made of aluminum alloys with stainless

steel and titanium alloys in specific areas.

COMPOSITE MATERIALS

Composite materials are used for primary and secondary structure.

Composite materials represent about 15% of the A/C structural weight.

Carbon Fiber Reinforced Plastic (CFRP) is mainly used for primary

structures, whilst Aramid Fiber Reinforced Plastic (AFRP) and Glass

Fiber Reinforced Plastic (GFRP) are only used for secondary

structures.


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AIRCRAFT MATERIALS - METALLIC MATERIALS & COMPOSITE MATERIALS


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STRUCTURE PROTECTION

AIRCRAFT DRAINAGE

Wings and fuselage have different types of drains. Holes and gaps are

meant to be used for a natural drainage of the fluid collection points.

Drain holes are drilled before application of pretreatments. Remote

drains are used when natural drainage is not possible.

SURFACE PRETREATMENT

The protection of the structure against corrosion is achieved by means

of appropriate surface pretreatment of the metallic parts.

Aluminum alloys: the primary protection is generally a pure aluminum

cladding. The main pretreatment used is the unsealed chromic acid

anodizing.Titanium alloys: surface interfaying with aluminum alloy parts are

zinc sprayed. The other titanium alloy surfaces are left bare. Titanium

fasteners are either sulphuric acid anodized or aluminum coated.

Composite materials are left bare.


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STRUCTURE PROTECTION - AIRCRAFT DRAINAGE & SURFACE PRETREATMENT


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STRUCTURE PROTECTION (continued)

PAINT SYSTEM

Before the final paint system, all aluminum parts are primed. The

paint system used includes polyurethane primers and paint on the

external surfaces, and epoxy primers and polyurethane paint on the

internal surfaces. Anti-slip paint is the overwing escape zones.


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STRUCTURE PROTECTION - PAINT SYSTEM


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STRUCTURE PROTECTION (continued)

NO STEP AREAS

Protective mats are required on the horizontal stabilizer as it is a carbon

fiber structure.

JACKING POINTS

Three jacking points are provided, one below each wing outboard of the

pylon and one in front of the NLG bay.


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STRUCTURE PROTECTION & JACKING POINTS


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A318 STRUCTURE DIFFERENCES

The Main structure differences between the A319/A320/A321 and the

A318 are due to the reduced length of the fuselage. There are several

general structure changes, laser beam welded structures and the vertical

stabilizer fin tip extension.

GENERAL STRUCTURE CHANGES

The main general structure changes are:

- on section 17, due to reduced length of the fuselage, the longitudinal

beams, the seat rails and the Z-profiles are replaced by new ones. The

crossbeams at FRame 52, FR53 and FR54 are removed. New

crossbeams are installed between FR55 to FR64,

- due to its location in the non-cylindrical part of the fuselage, a new

cargo sill box replaces the A319/A320/A321 one's, in section 17,

- on section 15, the A319/A320/A321 skin panels have been modified.

For weight reduction the A318 skin panels are thinner than the

A319/A320/A321 one's,

- the aft part of the belly fairing is modified due to an overlap with

non-cylindrical part of the fuselage. To avoid interference with cargo

compartment door, the A318 belly fairing is two panels shorter than

the A319/A320/A321 one's.


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A318 STRUCTURE DIFFERENCES - GENERAL STRUCTURE CHANGES


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A318 STRUCTURE DIFFERENCES (continued)

VERTICAL STABILIZER

Compared with A319/A320/A321 A/Cs, the A318 vertical stabilizer

fin tip is 750 mm (29,5 in.) longer.

The new developed tip is completely made of GFRP. There is an

additional fin leading edge panel. There is a new spar and a new CFRP

adaptor box, between the fin base and the fin tip.

The metallic rudder tip is longer by 100 mm in vertical direction. The

rudder trailing edge is increased in width by 50 mm.


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A318 STRUCTURE DIFFERENCES - VERTICAL STABILIZER


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LASER BEAM WELDING

The technology used for the A319/A320/A321 A/Cs is riveted

skin/stringer. On the A318, the skin/stringer connections are welded.

The new laser beam welded skin panels are installed in:

- the sections 13/14, FR24 to FR35, stringers 18 to 32,

- the sections 16/17, FR47/54 to FR64, stringers 32 to 41.

The skin panels are made thicker where the stringers are welded onto

them.


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A318 STRUCTURE DIFFERENCES - LASER BEAM WELDING


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CARGO DOORS

The A318 forward and aft cargo doors are smaller. The new cargo

door width is reduced from 1.82 m (71.5 in) to 1.28 m (50.5 in). The

under-floor cargo offers a usable volume of 21.21 m3. There is no

containerized cargo system option.


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A318 STRUCTURE DIFFERENCES - CARGO DOORS


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DOORS D/O (3)

GENERAL

The fuselage has:

- 4 passenger/crew doors,- 2 or 4, emergency exits depending on the A/C type,

- 2 cargo compartment doors,

- 1 bulk cargo compartment door (A320 & A321 only),

- landing gear bay doors and access doors for servicing and maintenance.


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GENERAL


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DOORS D/O (3)

PASSENGER COMPARTMENT DOORS

PASSENGER/CREW DOOR

The aircraft has four type C passenger doors, located on each side of

the fuselage at frame (Fr) 16/20 and 66/68.

Normal operation of the door is possible from the inside and the

outside of the aircraft. Arming of the emergency operation is only

possible from inside.

The doors are of fail-safe, plug-type construction. The structure is of

conventional design, composed of an outer skin, frame segments and

beams. Edgemembers built a surrouding frame on which hinge fittings

and locking mechanisms are installed. The loads resulting from cabin

pressure are transferred by stop fittings located on each side of the

door and the frame.All the doors include an evacuation system. The escape slides or slide

/ rafts are stowed at the lower part of the passenger/crew door.


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PASSENGER COMPARTMENT DOORS - PASSENGER/CREW DOOR


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DOORS D/O (3)

PASSENGER COMPARTMENT DOORS (continued)

EMERGENCY EXIT DOORS

On A318 and A319 aircraft are two Type III overwing emergency

exits installed, one on each side of the fuselage.

The A320 aircraft has four Type III overwing emergency exit doors,

two on each side of the fuselage.

In an emergency, these exits can be opened manually.

These emergency exits are of conventional plug type construction and

contain a standard size passenger cabin window.

The A321 aircraft has four Type "C" emergency exits, one on each

side of the fuselage sections 14A and 16A, between Fr 35.1 and 35.3A

and between Fr 47.2A and 47.4. The structural design and operation

of these plug-type exits is similar to the passenger doors.In an emergency, these exits can be opened manually; they are operated

like the passenger doors.

These emergency exits are of conventional plug-type construction.

A slide (or slide/raft) is installed in a compartment below each door.


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PASSENGER COMPARTMENT DOORS - EMERGENCY EXIT DOORS


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DOORS D/O (3)

CARGO COMPARTMENT DOORS

FWD & AFT CARGO DOORS

Two doors in the lower RH side of the fuselage provide access to the

main cargo compartments.

These doors are designed to carry loads from differential pressure and

circumferential loads of the frames from the fuselage. With this

consideration, they are of conventional design and have:

- an outer and inner skins,

- an internal structure of drop-forged machined circumferential frames.

The upper ends of these frames are connected to the hinges for the

door, and the lower ends are attachment for the locking hooks. The

A318 cargo doors cutout is reduced by 534 mm (one frame pitch).


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CARGO COMPARTMENT DOORS - FWD & AFT CARGO DOORS


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DOORS D/O (3)

CARGO COMPARTMENT DOORS (continued)

BULK CARGO DOOR (A320 & A321 ONLY)

The bulk cargo compartment, at the rear, has a conventional plug-type

door, located between Fr 60 and 62.

The door is operated, locked and unlocked manually and can be opened

from the outside.

It is opened by pushing inward and upward and is locked in the open

position onto the ceiling of the compartment. (In this compartment,

nets are provided to maintain the clearance for the door opening). The

weight of the door is compensated by a torsion bar. The door is

connected to the door locking warning system.


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CARGO COMPARTMENT DOORS - BULK CARGO DOOR (A320 & A321 ONLY)


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DOORS D/O (3)

ACCESS & SERVICE DOORS

The access doors are installed in the aircraft for inspection of the structure

and to give access to maintenance. Service doors are installed in thefuselage to give access to the servicing of systems.

All access and service doors are opened and closed manually.

Access and service doors are illustrated as follows:

- Avionics compartment door: there are four avionic compartment doors

like the one illustrated. This avionics compartment access door is installed

in the lower shell of the fuselage between Fr 3 and Fr 5 in a pressurized

area. The door can be opened from the inside or the outside.

- APU doors: The APU access doors are installed in the fuselage tail cone

in Zone 310. These doors are located in the lower part of the fuselage

between Fr 80A and Fr 84A. The doors give you access to the APU for

maintenance.

There are also access and service doors - not-illustrated: These doors are

located in the fuselage and belly fairing for water, waste, external power

and maintenance.


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ACCESS & SERVICE DOORS


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DOORS D/O (3)

LANDING GEAR DOORS

NOSE LANDING GEAR (NLG) DOORS

Landing gear doors give protection to the landing gear when the

aircraft is in flight.

The nose and auxiliary landing gear doors have five parts:

- two forward doors, hydraulically actuated, which can be closed with

the gear in the extended or retracted position. These doors are made

from CFRP (Carbon Fiber Reinforced Plastic) sandwich materials

with a honeycomb core. They are hinged to the landing gear bay

longitudinal edges.

- two aft doors, linked to the gear by a rotating rod, which are made

from CFRP sandwich materials with an honeycomb core. The purpose

of these doors hinged to the landing gear bay rear lateral edge, is toallow the forward doors to be retracted when the gear is extended.

- one small door (fixed door) attached to the landing gear leg is made

from aluminum alloy.


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LANDING GEAR DOORS - NOSE LANDING GEAR (NLG) DOORS


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DOORS D/O (3)

LANDING GEAR DOORS (continued)

MAIN LANDING GEAR (MLG) DOORS

The main landing gear doors are made from CFRP sandwich materials

with a honeycomb core, and have three parts:

- a main door, hydraulically actuated, which is hinged to the fuselage

keel beam parallel to the aircraft center line and can be closed with

the gear in the extended or retracted position,

- a fairing attached to the gear leg (fixed fairing door),

- a small door hinged to the wing structure in the neighborhood of the

upper end of the main leg (hinged fairing door).

All doors are part of the fuselage belly fairing and wing lower surface

in closed position.


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LANDING GEAR DOORS - MAIN LANDING GEAR (MLG) DOORS


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FUSELAGE D/O (3)

GENERAL

FUSELAGE LAYOUT

The fuselage is divided into five main parts:

- the nose forward fuselage (ATA 53-10-00),

- the forward fuselage (ATA 53-20-00),

- the center fuselage (ATA 53-30-00),

- the rear fuselage (ATA 53-40-00),

- and the cone/rear fuselage (ATA 53-50-00).


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GENERAL - FUSELAGE LAYOUT


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FUSELAGE D/O (3)

GENERAL (continued)

FUSELAGE BREAKDOWN

Compared with the A320, the A321 forward fuselage is eight framebays longer (additional section 14A, extending between frames (Fr)

35 and 35.8).

The A321 rear fuselage is five frame bays longer (additional section

16A, extending between Fr 47 and Fr 47.5.

Compared with the A320, the A319 forward fuselage (section 13/14)

and the rear fuselage (section 16/17) are respectively three frame bays

and four frame bays shorter.


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GENERAL - FUSELAGE BREAKDOWN


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FUSELAGE D/O (3)

NOSE FORWARD FUSELAGE

GENERAL ARRANGEMENT

The nose forward fuselage includes section 11, between Fr 1 and Fr12, and section 12, from Fr 12 to Fr 24.

The pressurized zone extends from Fr 1 to Fr 24.

The unpressurized zones are the radome, forward of Fr 1 and the nose

landing gear bay.

The structure of the nose forward fuselage has three parts:

- the forward upper structure, between Fr 1 and 11, which makes the

flight deck,

- the aft upper structure, between Fr 12 and 24, which makes the

forward part of the passenger cabin,

- the lower structure between Fr 1 and 24.


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NOSE FORWARD FUSELAGE - GENERAL ARRANGEMENT


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FUSELAGE D/O (3)

NOSE FORWARD FUSELAGE (continued)

FORWARD & AFT UPPER STRUCTURES

The forward upper structure between Fr 1 and Fr 12 includes:- closed frames,

- opened frames at level of openings for windshield and side windows,

- the forward pressure bulkhead,

- the flight deck floor support structure including two lateral boxes,

- the skin panels and the windshield frames,

The skin panels just above and below the windshield are made of

titanium alloy for bird impact requirements.

The aft upper structure, between Fr 12 and Fr 24, is the forward

passenger compartment and contains:

- the forward passenger/crew door between Fr 16 and 20,- conventional assembly of skin, stringers and frames,

- the floor support structure.


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NOSE FORWARD FUSELAGE - FORWARD & AFT UPPER STRUCTURES


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FUSELAGE D/O (3)


LOWER STRUCTURE

This part of section 11/12 contains the nose landing gear bay, accessand service door cutouts.

The nose landing gear bay is shaped by three machined panels

reinforced by horizontal and vertical extruded sections attached to the

corresponding frames. The lower parts of Fr 9 and Fr 20 are the

forward and rear limits of the gear bay.

The lower fuselage comprises three skin panels. The central panel has

an opening for access between Fr 3 and 5 and the opening for the nose

landing gear bay between Fr 9 and 20.

The right hand side panel has two openings for access, between Fr 12

and 14 and Fr 21 and 23.


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NOSE FORWARD FUSELAGE - LOWER STRUCTURE


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FUSELAGE D/O (3)

FORWARD FUSELAGE

GENERAL ARRANGEMENT

This area of the fuselage lies between Fr 24 and Fr 35.It contains the front part of the passenger cabin and beneath the cabin

floor and the forward cargo compartment. The forward cargo door is

on the starboard side.

The A321 section 14A extends from Fr 35 to Fr 35.8.

Section 14A is of similar construction to section 13/14 but includes

the emergency exit cut-outs (one on each side of the fuselage) between

Fr 35.1 and Fr 35.2A.


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FORWARD FUSELAGE - GENERAL ARRANGEMENT


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FUSELAGE D/O (3)

FORWARD FUSELAGE (continued)

TYPICAL STRUCTURE

This section is of conventional construction composed primarily of chemically milled skin panels, frames and stringers made in sheet

metal.

The standard frames have a common Z-shaped section made from

formed sheet, which provides a continuous structural member attached

to the skin and stringers by sheet metal cleats.

The structure of the cabin floor has:

- cross beams,

- seat tracks,

- floor support struts,

- floor panels.


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FORWARD FUSELAGE - TYPICAL STRUCTURE


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FUSELAGE D/O (3)

CENTER FUSELAGE

GENERAL ARRANGEMENT

The fuselage center section (section 15) extends from Fr 35 to Fr 47for A320, from Fr 35.8 to Fr 47 for A321 and from Fr 35 to Fr 47/51

for A319.

The upper section includes part of the passenger compartment.

The passenger floor structure is made of longitudinal beams, seat and

support tracks, support struts and floor panels.

The lower section is non-pressurized and integrates:

- the center wing box which extends across the width of the fuselage.

The two main frames 36 and 42 are also part of the center wing box,

- the main landing gear bay between Fr 42 and Fr 46,

- the keel beam which keeps the longitudinal structural continuity of the lower fuselage,

- the belly fairing supporting structure, panels and doors.


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CENTER FUSELAGE - GENERAL ARRANGEMENT


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FUSELAGE D/O (3)

CENTER FUSELAGE (continued)

KEEL BEAM

The longitudinal structural continuity of the lower fuselage in thisarea is maintained by the keel beam.

This beam is an aluminum alloy box structure, including skins,

stringers and ribs, and provides attachments for the main landing gear

doors and door actuators.

In its center area, the keel beam side walls are connected to the

wing-box aft lower panel.


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CENTER FUSELAGE - KEEL BEAM


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FUSELAGE D/O (3)


BELLY FAIRING

The belly fairing includes a substructure made of aluminum alloyframes and webs which are attached to the fuselage via fittings and

rods.

This substructure supports the panels made of composite materials.

The belly fairing also includes the landing gear doors, external access

panels and access doors for maintenance.


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CENTER FUSELAGE - BELLY FAIRING


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FUSELAGE D/O (3)

REAR FUSELAGE - A319 & A320 GENERAL

ARRANGEMENT

The rear fuselage assembly is a pressurized area, which extends from Fr47 to Fr 70.

The A319 and A320 rear fuselage is divided into two sections (the A321

has an additional section 16A):

- section 16/17 between Fr 47 and Fr 64,

- section 18 between Fr 64 and Fr 70.

Section 16/17 is shorter by four frames than on the A320.

The upper part of the fuselage contains the aft section of the passenger

cabin and the aft passenger/crew doors located between Fr 66 and Fr 68.

The lower part contains the aft cargo compartment. The aft cargo

compartment door is installed between Fr 52A and Fr 56 (RH side); the

bulk cargo compartment door is installed between Fr 60 and Fr 62 (RHside).

The design of section 16/17 is similar to that of forward fuselage sections

(typical skin, stringer and frame arrangement).

Skin panels of the lower area have support attachment structures for the

belly fairing rear part.


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REAR FUSELAGE - A319 & A320 GENERAL ARRANGEMENT


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REAR FUSELAGE - A321 GENERAL ARRANGEMENT

The A321 rear fuselage assembly is a pressurized area, which extends

from Fr 47 to Fr 70.The A321 rear fuselage is divided into three sections:

- section 16/17 and 18 which are similar to the A320,

- section 16A,

The section 16A of the A321 fuselage extends from Fr 47 to Fr 47.5.

The section 16A includes the passenger cabin part in the upper section,

and beneath the cabin floor, the forward part of the rear cargo

compartment.

The section 16A is of similar construction to section 16/17 but includes

the emergency exit cut-outs (one on each side of the fuselage) between

Fr 47.2A and Fr 47.4.The slide/slide-raft is installed in a separate

compartment below each door.


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CONE/REAR FUSELAGE

GENERAL ARRANGEMENT

This section comprises the un-pressurized part of the rear fuselageextending from Fr 70 to Fr 87.

It includes:

- the mounting structures for the vertical and horizontal stabilizers,

- the rear pressure bulkhead,

- a jacking point,

- attachment structure for the tail cone, which houses the Auxiliary

Power Unit (APU).

It is divided into two main sections:

- section 19 between Fr 70 and Fr 77,

- section 19.1(tail cone) aft of Fr 77.

Section 19 is composed of chemically milled skins, riveted stringers

and frames.

The side skin panels include the horizontal stabilizer cut-out. The

lower panel has an access door for this section where a maintenance

floor is installed.


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CONE/REAR FUSELAGE (continued)

REAR PRESSURE BULKHEAD

The rear pressure bulkhead installed at Fr 70, divides the pressurizedrear fuselage from the cone/rear fuselage, which is not pressurized.

It is made of a spherical membrane, and four aluminum alloy sheet

segments joined together on the inner surface by means of four "I"

profile sections. Four additional "I" profile radial stiffeners are also

installed.


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FUSELAGE D/O (3)


VERTICAL STABILIZER ATTACHMENT FITTINGS

The vertical stabilizer spar box attachment fittings are located at Fr70, Fr 72 and Fr 74.

They have six fail safe yokes, which transmit the vertical stabilizer

loads into the fuselage frames via shear bolts.

The upper segments frames 70, 72 and 74 are machined from plates

while the lower segments are made from sheet metal.


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THS ATTACHMENT FITTINGS

The fuselage area between Fr 73 and Fr 77 houses the horizontalstabilizer.

There is a large cut-out between Fr 73 and Fr 77, which is surrounded

by machined beams. A system of diagonal struts is installed on the

horizontal plane in the upper and lower areas of the cutout to increase

the rigidity of this open section.

The machined frame 77 supports the tailplane hinge bearings and the

lateral load fittings. They introduce horizontal stabilizer loads into

the fuselage structure, via the central bracing structure and the upper

and lower bracing structures.

Frame 77 also includes four lugs for the attachment of the tail cone

unit.


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TAIL CONE

The tail cone unit is located aft of Fr 77 and houses the APU. Thissection is connected to section 19 by means of four lugs and one

spigot.


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A318 FUSELAGE D/O (3)

GENERAL

FUSELAGE LAYOUT

The fuselage is divided into five main parts: the nose forward fuselage(section 11/12), the forward fuselage (section 13/14), the center

fuselage (section 15), the rear fuselage (sections 16/17 and 18) and

the cone/rear fuselage (section 19/19.1).


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GENERAL (continued)

FRAME/SKIN/STRINGER ASSEMBLY

Standard frames have a common z-shape section made from formedsheet. These frames are continuous structural members attached to

the skin and stringers by sheet metal cleats.

A panel with laser beam welded stringers has been introduced:

- in section 13, between frames (Fr) 24 and 35, from stringer (Stgr)

18LH to Stgr 32LH,

- in section 16/17, between Fr 47/54 and 64, from Stgr 32LH to Stgr

41RH.


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NOSE FORWARD FUSELAGE

GENERAL ARRANGEMENT

The nose forward fuselage has section 11, from Fr 1 to Fr 12 andsection 12, from Fr 12 to Fr 24. The pressurized area extends from Fr

1 to Fr 24. The unpressurized areas are the radome, forward of Fr 1,

and the nose landing gear bay.


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UPPER STRUCTURE

The upper structure between Fr 1 and Fr 12 has closed frames andopened frames at level of openings for:

- the windshield and side windows,

- the forward pressure bulkhead,

- the flight deck floor support structure,

- skin panels and windshield frames.

The upper structure between Fr 12 and Fr 24 makes the forward

passenger compartment and contains the two forward passenger/crew

doors.


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LOWER STRUCTURE

This part of section 11/12 contains the nose landing gear bay, accessand service door cutouts. The nose landing gear bay is made of

machined flat panels stabilized laterally and longitudinally by struts.

The struts are attached respectively to frames and flight deck

crossbeams.


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FORWARD FUSELAGE

GENERAL ARRANGEMENT

This region of the fuselage lies between Fr 24 and 35. It contains thefront part of the passenger cabin and, beneath the cabin floor, the

forward cargo compartment. The forward cargo door is located

between Fr 24A and 28 on the RH side of the fuselage


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FORWARD FUSELAGE (continued)

TYPICAL STRUCTURE

This section is of conventional construction, having chemically milledskin panels, frames and stringers made from sheet metal. The standard

frames have a common Z-shaped section made from formed sheet.

They are continuous structural members attached to the skin and

stringers by sheet metal cleats. A skin panel with laser beam welded

stringer is installed between Fr 24A and 35, and between Stgr 18LH

and 32LH.


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CENTER FUSELAGE

GENERAL ARRANGEMENT

The fuselage center section extends from Fr 35 to Fr 47/54, andintegrates the center wing box. The upper section contains a part of

the passenger compartment, with two overwing emergency exit door

cutouts. The pressure boundary is delimited by the forward bulkhead

at Fr 35, the upper skin panel of the center wing box prolonged by a

pressure diaphragm up to frame 46 and ending by an inclined pressure

bulkhead. Beneath the cabin floor are the air conditioning, hydraulic

and main landing gears, in conjunction with a belly fairing.


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KEEL BEAM

In this area, the longitudinal structural continuity of the lower fuselageis maintained by a keel beam located between Fr 35.8 and 46. The

keel beam transmits the overall fuselage vertical bending loads. This

beam is a box structure having attachments for the main landing gear

doors and door actuators. In its center region, the keel beam side walls

are connected to the bottom skin panels of the center wing box.


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BELLY FAIRING

The belly fairing has a substructure made of aluminum alloy framesand webs, attached to the fuselage via fittings and rods. This

substructure supports the panels, made of sandwich construction. The

belly fairing also incorporates the landing gear doors, external access

panels and access doors for maintenance.


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REAR FUSELAGE - GENERAL ARRANGEMENT

The rear fuselage assembly is a pressurized area, which extends from Fr

47/54 to Fr 70. It is divided into two sections:

- section 16/17 between Fr 47/54 and 64,- section 18 between Fr 64 and 70.

The design of section 16/17 is similar to that of forward fuselage sections.

Skin panels of the lower region have support attachment structures for

the belly fairing rear part. The aft cargo door cutout is located between

Fr 57A and 60 on the RH side of the fuselage. Aft passenger door cutouts

are located between Fr 66 and 68.


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CONE/REAR FUSELAGE

GENERAL ARRANGEMENT

This section is the unpressurized part of the rear fuselage, aft of Fr

70. It has the mounting structure for vertical and horizontal stabilizers

and houses the Auxiliary Power Unit (APU). It is divided into two

main sections:

- section 19 between Fr 70 and 77,

- section 19.1 (tail cone) aft of Fr 77.

Section 19 has chemically milled skins, riveted stringers and frames.

Side skin panels have the horizontal stabilizer cutout. The lower panel

has a door, which gives access to this section where a maintenance

floor is installed.


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REAR PRESSURE BULKHEAD

The Fr 70 supports the rear pressure bulkhead, designed as a pressure

diaphragm. It is made of aluminum alloy. The bulkhead is attached

to the inside of the fuselage with a connecting strap, made of aluminum

alloy.


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VERTICAL STABILIZER ATTACHMENT FITTINGS

The vertical stabilizer spar box attachment fittings are three pairs of

fail safe yokes, made from forging aluminum alloy. They transmit the

fin loads into the fuselage and are located at Fr 70, 72 and 74. At those

locations, the upper frame segments are made of integrally machined

plates.


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THS ATTACHMENT FITTINGS

To house the Trimmable Horizontal Stabilizer (THS), there is a large

cutout in the fuselage between Fr 74 and 77. Frame 77 is made of

integrally machined plates and carries the THS bearing loads with the

vertical link fittings. The side loads are carried through an eye bolt,

linked to:

- the side load fitting on the rear spar of the THS,

- and oblique struts attached to the lower and upper areas of Fr 77.


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TAIL CONE

The tail cone unit is located aft of Fr 77 and houses the Auxiliary

Power Unit (APU). This section is connected to section 19 by means

of four lugs and one spigot.


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PYLONS/NACELLES D/O (3)





GENERAL

The function of the engine pylons installed under each wing is:

- to support the engine,

- to transmit the engine thrust to the aircraft,- to enable the routing and attachment of all the systems connected with

the engine (electrical wiring, hydraulic, bleed air and fuel lines).

The nacelle gives the engine an aerodynamic shape and supports the

thrust reverser system.

Information concerning structure of the nacelle can be found within the

nacelle manufacturer documentation.


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GENERAL


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PYLONS - GENERAL ARRANGEMENT

The pylon has:

- a primary structure attached to the wing and supporting the engine,

- a secondary structure, essentially fairings, housing most of the systems.


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PYLONS PRIMARY STRUCTURE - PYLON BOX

GENERAL ARRANGEMENT

The pylon box is the primary structure. It supports the engine by two

points and is attached to the wing at three points.


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PYLONS PRIMARY STRUCTURE - PYLON BOX (continued)

MAIN ASSEMBLY

The pylon box is composed of ribs, two upper spars and one lower

spar, and panels mainly made from steel and titanium alloys.


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PYLON TO WING ATTACHMENT

The forward pylon to wing attach fitting has a double lugged fork

attachments connected to the wing fitting by means of four shackles.

This fitting located at Rib 4 is made of titanium alloy and carries

vertical loads.

The aft pylon to wing attach fitting has a single fail safe lug connected

to the wing fitting by means of two shackles. This fitting located at

Rib 10 is made of titanium alloy and carries vertical and side loads.

Immediately behind the forward attach fitting a spherical bearing

transmits the thrust to a spigot bolted to the bottom wing skin panel.


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PYLON TO ENGINE ATTACHMENT

At The forward engine to pylon attach fitting there is a pyramid

attached to the rib and made of steel alloy.

This fitting transmits the engine thrust, side loads and vertical loads.

At The aft engine to pylon attach fitting there is an engine mount

located at Rib 3 for CFM 56-5 engine configuration or at Rib 4 for

IAE V2500 engine configuration. This fitting reacts to vertical loads,

side loads and roll movement.


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PYLONS SECONDARY STRUCTURE

GENERAL ARRANGEMENT

The secondary structure is composed of:

- the forward fairing,- the pylon to wing center fillets,

- the aft fairing,

- the lower fairing.


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( )

PYLONS SECONDARY STRUCTURE (continued)

FORWARD FAIRING

The forward fairing can be divided into two sections; the cantilever

structure between Rib 01 and Rib 05, and the structure between Rib05 and Rib 9.

The cantilever structure gives an aerodynamic contour between the

engine nose cowl and the pylon box structure. It routes all systems

and the bleed air from the engine to the fuselage.

The structure between Rib 05 and Rib 9 gives an aerodynamic contour

between the cantilever structure and the wing leading edge, and enables

the routing of various system lines and electrical wiring.

It includes in particular two pressure relief doors (made from titanium),

which are designed to open in case of hot bleed air duct bursting.

The structure is mainly made of stainless steel alloy.


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( )


PYLON TO WING CENTER FILLETS

The pylon to wing center fillets give an aerodynamic contour between

the pylon box and the wing bottom skin panel.The pylon-to-wing center fillets are made of aluminum alloy ribs.

These ribs support the panels made of hybrid Carbon Fiber Reinforced

Plastic (CFRP)/Aramid Fiber Reinforced Plastic (AFRP) sandwich

construction.


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AFT FAIRING

The aft fairing is a removable secondary structure composed of two

parts:- a fixed fairing located at the rear of the pylon box,

- a movable fairing underneath the flap.

The fixed fairing is attached by two points to the pylon box at Rib 10

and by one point to the wing box at the false rear spar.

The fixed fairing is assembled of ribs and skin panels made of

aluminum alloy, and includes a lower aft fairing made in AFRP

sandwich construction.

The movable fairing is hinged at Rib 14 and linked to the flap by a

rod attached to the fairing by a serrated plate system.

The internal structure of the movable fairing is mainly made of aluminum alloy. The side panels are made in CFRP or AFRP sandwich

construction.


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LOWER FAIRING

A fairing located under the pylon box (lower fairing) makes sure there

is a continuity of the aerodynamic profile between the pylon box andthe engine nozzle.

Its function is:

- to supply thermal protection to the pylon from the engine exhaust

gases,

- to smooth out protrusions with minimal aerodynamic drag changes.

The lower fairing is made of stainless steel alloy sheet except for the

bottom removable sole which is made of inconel 625 alloy.


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PYLON TO NACELLE JUNCTION

The pylon to nacelle junction has:

- Fan cowl door attachments.

The hinge fittings of the fan cowl doors are located at Rib 01, Rib 03 andRib 05. They are made of titanium and installed on the forward secondary

structure.

- Thrust reverser doors attachments

The hinge fittings of the thrust reverser doors are located at Rib 1 and

Rib 2. They are made of titanium and installed on the primary structure

(pylon box). An other hinge (tie-bar) goes through the secondary structure.


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NACELLES - GENERAL

The nacelle cowling includes the inlet cowl, the fan cowl, the thrust

reverser and the exhaust nozzle.

There are two types of engine: CFM and IAE.The IAE nacelle is installed with a Common Nozzle Assembly (CNA).

The nacelles are under the responsibility of engine manufacturers.


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STABILIZERS - GENERAL ARRANGEMENT

Stabilizers are composed of: Trimmable Horizontal Stabilizer (THS),

elevators, the vertical stabilizer and rudder.


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TRIMMABLE HORIZONTAL STABILIZER (THS)

GENERAL ARRANGEMENT

The THS main structure has:

- the spar boxes (Center, Left Hand (LH) and Right Hand (RH) sides),- the leading edge,

- the trailing edge,

- the attachment fittings.

The spar boxes are the primary structure of the horizontal stabilizer

and support all the other components.


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(continued)

SPAR BOXES

The complete spar box assembly has the LH and RH boxes and thecenter joint.

Each spar box includes top and bottom skin panels, a front spar, a rear

spar and thirteen ribs (from Rib 2 thru Rib 14).

The LH and RH spar boxes are laminated in Carbon Fiber Reinforced

Plastic (CFRP).

The center joint is made from titanium and connects the LH and RH

spar boxes to make one single unit.


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(continued)

MAIN SUPPORT FITTINGS

A hydromechanical actuator enables the adjustment of the angle of incidence of the THS. The actuator is connected to a dual fitting (front

spar fitting) at the forward end of Rib 1, by means of hinge arms.

The THS is attached to the cone rear fuselage structure at two pivot

points (rear support fittings). They are installed on each side of the

THS centerline at Rib 3. All fittings are made of CFRP.


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(continued)

ELEVATOR ATTACHMENT FITTINGS

Each rear spar bears six elevator hinge arms and two fittings for theattachment of the elevator servocontrol actuators.


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(continued)

LEADING EDGE

The leading edge has an aerodynamic shape at the front of the THS.On each side of the THS centerline, the THS leading edge includes:

- three leading edge primary ribs,

- one inboard leading edge section,

- one outboard leading edge section and,

- one leading edge intersection.

All components are laminated in CFRP.

The front part of the inboard and outboard leading edge sections has

a stainless steel protection, bonded to the leading edge.

The leading edge intersection is fitted to Rib 1 and to the spar box. A

rubber strip is fitted to the intersection. It seals the gap between the

fuselage skin and the leading edge intersection.


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(continued)

TIP

The tips of the THS are the LH and RH outer fairings. The tips aremade of aluminum alloy and include rib and skin panels. The tips are

attached to the leading edge rib 25 and to the upper and lower shells

of the spar box.


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(continued)

TRAILING EDGE

The trailing edge shapes an aerodynamic surface between the THSspar box and the elevator.

On each side of the THS centerline, the trailing edge panels are

supported by six intermediate ribs, and by the six hinge elevator arm

supports.

The access panels are laminated in CFRP bonded to a honeycomb

core.

On each side there are four panel assemblies on the top surface and

four access panels on the bottom surface. A rubber seal is installed

between the panel assemblies and the access panels along the trailing

edges to prevent dirtiness.


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ELEVATORS - STRUCTURE LAYOUT




ELEVATORS STRUCTURE LAYOUT

The structure of each elevator includes:

- a front spar,

- top and bottom skin panels,

- four ribs.

All components are laminated in CFRP; the top and bottom panels are

made in sandwich construction.

Rivets attach an aluminum profile to the trailing edge to make the trailing

edge stronger.

Six hinge fittings attach each elevator to the spar box of the THS. Two

fittings attach the servo control units. You can remove the leading edge

access panels, the tips and the inboard end caps.

Each elevator has three hoisting points and four static dischargers.


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VERTICAL STABILIZER




VERTICAL STABILIZER

GENERAL ARRANGEMENT

The vertical stabilizer is attached to the top of the rear fuselage. It

supports the rudder, which is operated by three servo control units.The High Frequency (HF) antenna and the Very high frequency

Omnibearing Range (VOR) antenna are also attached to it.

The main components of the vertical stabilizer are:

- the spar box,

- the leading edge,


- the tip,

- the attach fittings.


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VERTICAL STABILIZER (continued)




( )

SPAR BOX - STRUCTURE LAYOUT

The spar box is the primary structural component of the vertical

stabilizer.All the other components of the vertical stabilizer are attached to it.

The spar box has a front, a center and a rear spar, ribs and two side

panels with stiffeners, all laminated in CFRP.

Three pairs of main attach fittings made of CFRP attach the spar box

to the fuselage.


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SPAR BOX - STRUCTURE LAYOUT (CONT'D)

The seven rudder hinge arms and the three actuator hinge fittings are

made from aluminum alloy and are attached to the spar box rear spar.


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LEADING EDGE

The vertical stabilizer leading edge has three removable sections.

They are attached to the forward edge of the spar box side panels andto the leading edge ribs. The lower section gives access to the HF

antenna (see ATA 53 fuselage description for the lower section).

The three sections give an aerodynamic shape to the front of the

vertical stabilizer.

The three sections are laminated in Glass Fiber Reinforced Plastic

(GFRP) bonded to a honeycomb core. A stainless steel cover is bonded

to the inner surfaces of the sections and protects them against hail and

bird impact damage.


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TIP - STRUCTURE LAYOUT

The vertical stabilizer tip is laminated in GFRP bonded to a

honeycomb core. It is attached to the leading edge end rib and thestabilizer spar box. An aluminum lightning strike protection strap is

bonded along the top of the tip.


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TRAILING EDGE

The trailing edge is attached to the rear of the vertical stabilizer.

It has a basic aluminum alloy supporting structure made of sparsections and profiles, and four access panels on each side. The panels

give access to the rudder servo control actuators and the hinge arms.

The panels are laminated in CFRP and GFRP bonded to a honeycomb

core.


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RUDDER




GENERAL ARRANGEMENT

The rudder is one of the primary flight controls of the aircraft.

The components of the rudder are:- the main structure,

- the leading edge,

- the tip,

- the hinge and actuator fittings.


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RUDDER (continued)




STRUCTURE LAYOUT

The rudder main structure is the primary structural component of the

rudder.It is an assembly of two CFRP sandwich panels, CFRP laminates front

spar, top and bottom closing ribs.

An access panel, installed on the left hand side shell, gives access to

the No. 7 rudder hinge fittings. At the other locations, cutouts in the

side shells give access to the adjacent hinge fittings.

Three actuators and seven rudder hinge fittings are attached to the

forward face of the rudder main structure, and rivets attach them to

the spar and to the skin panels.


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A318 STABILIZERS D/O (3)

STABILIZERS




GENERAL ARRANGEMENT

Stabilizers are composed of the Trimmable Horizontal Stabilizer

(THS), the elevators, the vertical stabilizer and the rudder.


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GENERAL ARRANGEMENT

The THS main structure has the LH and RH side spar boxes, the

leading edge, the trailing edge, the THS tip and the attachment fittings.The spar boxes are the primary structure of the horizontal stabilizer

and support all the other components.


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(continued)




( )

SPAR BOXES

The complete spar box assembly has the LH and RH spar boxes joinedtogether with a center joint to make one single unit. Each spar box

includes top and bottom skin panels, a front spar, a rear spar and

thirteen ribs (from Rib 2 to Rib 14), all parts being laminated in Carbon

Fiber Reinforced Plastic (CFRP).

The center joint includes a web (Rib 1) made of CFRP and upper and

lower fittings made of titanium.


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(continued)




MAIN SUPPORT FITTINGS

The front spar joint at Rib 1 made of CFRP supports the trim actuatorhinge arms.

The THS is attached to the cone/rear fuselage at Rib 3.

On each spar box side, the attachment fittings include a THS rear

support fitting of fail safe design with a lower and upper support

fittings, and a side load fitting. All fittings are made of CFRP except

the side load fitting made of aluminum alloy.


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(continued)




ELEVATOR ATTACHMENT FITTINGS

On each THS box, the rear spar bears:- six hinge arms, made of CFRP, for the attachment of the elevator,

- two fittings, made of CFRP, for the attachment of the elevator servo

control actuators.


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(continued)




LEADING EDGE

At the front of the THS, the leading edge gives an aerodynamic shape.On each side of the THS centerline, the THS leading edge includes:

three leading edge primary ribs, one inboard leading edge section,

one outboard leading edge section and one leading edge intersection.


The front part of the inboard and outboard leading edge sections has

a stainless steel protection; it is bonded to the leading edge.

The leading edge intersection is attached to Rib 1 and to the spar box.

A rubber strip is installed at the intersection, it seals the gap between

the fuselage skin and the intersection.


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(continued)




TIP

The tips of the THS are the LH and RH outer fairings. The tips are

attached to the leading edge Rib 25 and to the top and bottom skin

panels of the spar box. The tips are made from aluminum alloy.


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(continued)




TRAILING EDGE

The trailing edge has an aerodynamic surface between the THS spar

box and the elevator.

On each side of the THS centerline, the trailing edge panels are

supported by six intermediate ribs, and by seven hinge arm supports.

The panels are laminated in CFRP bonded to a honeycomb core.

On each side there are four panel assemblies on the top surface and

four access panels on the bottom surface. A rubber seal is installed

between the panel assemblies and the access panels along the trailing

edges to prevent dirtiness.


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ELEVATORS

STRUCTURE LAYOUT




The structure of each elevator has a front spar, a top and a bottom

skin panel and four ribs.


Six hinge fittings attach each elevator to the spar box of the THS and

two fittings attach the servo control units. You can remove the leading

edge access panels, the tips and the inboard end caps. Each elevator

has three hoisting points and four static dischargers.


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VERTICAL STABILIZER

GENERAL ARRANGEMENT




The vertical stabilizer is attached to the top of the rear fuselage. It

supports the rudder, which is operated by three servo control units.

The High Frequency (HF) antenna and the Very high frequency

Omnibearing Range (VOR) antenna are also attached to it.

The main components of the vertical stabilizer are:

- the spar box,

- the leading edge,


- the tip,

- the attach fittings.


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STRUCTURE LAYOUT




The spar box is the primary structural component of the vertical

stabilizer.

All the other components of the vertical stabilizer are attached to it.

The spar box has: a front, a center and a rear spar, ribs and side panels

with integrated stiffeners, all laminated in CFRP.

Three pairs of primary attach fittings made of CFRP attach the spar

box to the fuselage.


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SPAR BOX - STRUCTURE LAYOUT (CONT'D)




The seven rudder hinge arms and the three actuator hinge fittings are

made from aluminum alloy and are attached to the spar box rear spar.


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VERTICAL STABILIZER - SPAR BOX - STRUCTURE LAYOUT (CONT'D)


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LEADING EDGE




The vertical stabilizer leading edge has four removable sections. They

are attached to the forward edge of the spar box side panels and to the

leading edge ribs. The lower section gives access to the High

Frequency (HF) antenna.

The four sections give an aerodynamic shape to the front of the vertical

stabilizer.

The four sections are laminated in Glass Fiber Reinforced Plastic

(GFRP) bonded to a honeycomb core. A protective foil is bonded to

the inner surfaces of the sections and protects them against hail and

bird impact damage.


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TIP - STRUCTURE LAYOUT




The vertical stabilizer tip is laminated in GFRP bonded to a

honeycomb core. It is attached to the leading edge end rib and the

stabilizer spar box. An aluminum strap is bonded along the top of the

tip for lightning strike protection.


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TRAILING EDGE

Th t ili d i tt h d t th f th ti l t bili




The trailing edge is attached to the rear of the vertical stabilizer.

It has a basic aluminum alloy supporting structure made of spar

sections and profiles and four access panels installed on each side.The panels give access to the rudder hydraulics, the servo controls,

the control rods and the hinge arms.

The access panels are made of CFRP and GFRP laminations bonded

to a honeycomb core.


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RUDDER

GENERAL ARRANGEMENT

The rudder is one of the primary flight controls of the aircraft




The rudder is one of the primary flight controls of the aircraft.

The main components of the rudder are:

- the main structure,- the leading edge,

- the tip,

- the hinge fittings and the actuator fittings.


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RUDDER (continued)

STRUCTURE LAYOUT






rudder.

It has an assembly of two side skin panels, a front spar, a bottomclosing rib and a top closing rib. All components of the rudder main

structure are laminated in CFRP and are attached to the rudder main

structure.

Seven rudder hinge fittings and three actuator fittings are installed on

the front spar of the rudder (all fittings are made from aluminum alloy).

An aluminum profile is installed on the trailing edge of the rudder for

lightning strike protection.


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WINDOWS D/O (3)

GENERAL

The windows are installed in:

- the cockpit,




- the cabin,

- the doors.


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GENERAL


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COCKPIT WINDOWS

GENERAL ARRANGEMENT

There are two types of windows:




w yp w w

- the fixed windows,

- the sliding windows.The fixed windows are described as follows:

There are four fixed windows installed in the cockpit.

- two windshields,

- two fixed side windows.

The left and right windows are symmetrical.

These windows are mounted in a frame and can be removed and

installed from the exterior.

The sliding windows are installed as follows:

- on a mobile frame with a mechanism controlled from the cockpit.


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COCKPIT WINDOWS (continued)

WINSHIELDS STRUCTURE

The windshield panels are made of several layers of different materials




p y

depending on the windows supplier (LUCAS-ACT, PPG, SPS), and

are interchangeable. They are held by three retainers bolted onto theouter surface of the frame. They are installed with an anti-icing and

defogging system.


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FIXED SIDE WINDOWS

The fixed side windows have of two layers of different materials





are interchangeable. They are held by retainers bolted onto the innersurface of the frame. They are installed with an integral anti - icing

and defogging system.


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SLIDING WINDOWS

The sliding windows have several layers of different materials





are interchangeable. Each panel has an anti-icing and defoggingsystem. The sliding windows are installed on a mobile frame, which

is controlled from the cockpit, and the crew can use them as emergency

exits.


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CABIN WINDOWS

GENERAL ARRANGEMENT

The windows are installed in frames and make a smooth surface with

h f l ki




the fuselage skin.

The cabin windows are installed and removed from the inside of theaircraft.

The windows have a circular seal, inner and outer panes made of

stretched acrylic resin held together by a retainer ring, and eye bolts.

A vent hole in the inner pane lets the cabin pressure maintained in the

window.


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DOOR WINDOWS

STRUCTURE LAYOUT

The passenger/crew doors and emergency exit doors have a circular

i d Th d f i ti d b ti i d t




window. They are used for inspection and observation in order to

check if the cabin is pressurized and the escape slide is armed.The windows have a circular seal, inner and outer panes made of

stretched acrylic, held together by a retainer ring. A vent hole in the

inner pane lets the cabin pressure maintained in the window.


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WINGS D/O (3)

GENERAL

The aircraft wing is in the continuity of the structure going through the

fuselage which is divided into three parts:

- the center wing box,




g ,

- the left outer wing and,- the right outer wing.


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GENERAL


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WINGS D/O (3)

CENTER WING BOX

GENERAL ARRANGEMENT

The center wing is installed in the center fuselage between the main

frames (Fr) 36 and 42 to make an integral fuel tank




frames (Fr) 36 and 42 to make an integral fuel tank.

The center wing box structure has:- the front and the rear spars respectively located at Fr 36 and 42,


- the two main frames 36 and 42,

- internal spanwise lattice ribs,

- the left rib 1 and the right rib 1.

The junction between the center wing box and the outer wings is done

at the left hand and right hand sides rib 1.

The access for maintenance to the center wing box is done through

two triangular openings in the rear spar.


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CENTER WING BOX (continued)

WING ROOT JOINT

An upper cruciform fitting and a lower triform fitting ensure the

junction between the center wing box and the outer wing box.




junction between the center wing box and the outer wing box.

The upper cruciform fitting makes the junction between the centerwing box top skin panels, the outer wing box top skin panels, fuselage

and Rib 1.

The lower triform fitting and a safety butt-strap fitting make the

junction between the center wing box bottom skin panels, the outer

wing box bottom skin panels and Rib 1.


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OUTER WING

GENERAL ARRANGEMENT

Each outer wing has:

- a main structure (outer wing box),




( g ),

- a wing tip,- a leading edge and leading edge devices,

- a trailing edge and trailing edge devices.

The trailing edge control surfaces are:

- the inboard flap,

- the outboard flap,

- the two ailerons,

- the six spoilers.


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OUTER WING BOX

GENERAL ARRANGEMENT

The outer wing box tapers from Rib 1 (the wing root) to Rib 27 hold:

- the wing spars (front and rear),




- the ribs,- the top and bottom skin panels,

- the top and bottom stringers,

- the wing-root joint.


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OUTER WING BOX (continued)

SKIN PANELS

The top and the bottom surfaces of the outer wing box are made of

skin panels machined from aluminum alloy.




There are three panels on each surface. The skin panels are stiffenedby stringers machined in aluminum alloy extrusions.

The joints between panels are aluminum alloy butt straps.


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RIBS & SPARS

Ribs:

There are 27 ribs, machined in aluminum alloy, installed in the outer




wing box of each outer wing. Each rib is continuous between the frontand rear spars. The junction between the center wing box and the outer

wing joint is at Rib 1. Rib 1 is the boundary of the lateral section of

the center wing box. Ribs 22 and 27 make the other lateral boundaries

of the fuel and vent tanks.

Spars:

The wing spars are machined in aluminum alloy. They give strength

to the wing box and they extend from Rib 1 to Rib 27.


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ACCESS HOLES/COVERS

There are twenty-one access covers installed in the bottom skin panels

of the outer wing box. All panels close the openings that give access




to the outer wing box.There are:

- seven non load-carrying access panels between Rib 1 and Rib 13,

clamped on the wing skin,

- fourteen load-carrying access panels between Rib 14 and Rib 27,

bolted through the skin panel.


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FIXED LEADING EDGE

GENERAL ARRANGEMENT

The fixed Leading Edge (LE) assembly is located forward of the front

spar of the wing-box.





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FIXED LEADING EDGE (continued)

STRUCTURE LAYOUT (1/2)

The fixed leading edge assembly is made of:

- the D-nose assembly, composed of aluminum alloy parts:




- the support ribs and riblets (riblets are installed between the wingbox front spar and the LE spar),

- the sub spar,

- the LE skin.

- three top surface access panels,

- bottom surface access panels, which are made of Carbon Fiber

Reinforced Plastic (CFRP) sandwich construction and are attached

with quick release fasteners.


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FIXED LEADING EDGE (continued)


Two pylon ribs are installed on each side of the engine pylon. These

ribs hold the pylon shroud panels.





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SLATS

GENERAL ARRANGEMENT

The wing leading edge is fitted of five slats, which make the movable

part of the wing leading edge.





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SLATS (continued)


Each slat has:

- a front spar or the stringers (girders),




- a rear spar,- a girder

- ribs,


- a trailing edge assembly.

Slat 1 is supported by 4 tracks, two of them being driven (track 2 and

3).

Slats 2 to 5 are supported by two tracks, both being driven.


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SLATS (continued)


When the slats are in retracted position, seals prevent airflow between

the slat and the wing.

Sl t 3 t 5 d i d th h t i f th bl d i t




Slats 3 to 5 are de-iced; the hot air comes from the bleed air systemand is supplied to these slats through a telescopic duct (not shown)

and piccolo tubes installed in the leading edges of the slats.


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FIXED TRAILING EDGE

STRUCTURE LAYOUT

The fixed trailing edge is located aft of the wing rear spar.

Its structure has:

an overwing panel and an under wing panel




- an overwing panel and an under wing panel,- a shroud box and a fixed shroud,

- a false rear spar,

- a main landing gear attachment,

- structures support for the trailing edge control surfaces,

- access panels.


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FIXED TRAILING EDGE (continued)

STRUCTURE LAYOUT (CONT'D)

This page deals with the fixed trailing edge inner structure.





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TRAILING EDGE DEVICES

GENERAL ARRANGEMENT

The trailing edge devices are:

- two flaps,

- one aileron,

fi il




- five spoilers.


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TRAILING EDGE DEVICES (continued)

FLAPS GENERAL ARRANGEMENT

Two flaps are installed on the TE of the outer wing. The inboard flap

is installed between Rib 1 and Rib 9 and the outboard flap is installed

between Ribs 9 and 20.

The flaps are connected to each other through an interconnection strut




The flaps are connected to each other through an interconnection strut.

In case of a drive station failure, this device carries the loads, which

result in such failure.


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INBOARD FLAP STRUCTURE - A318-A319-A320

The inboard flap is supported and driven by a fuselage track and

carriage at track 1 and a wing track carriage at track 2.

The inboard flap has:

- a leading edge with CFRP skin




- a leading edge with CFRP skin,

- a flap box with:

- skin panels and integrated stringers made of CFRP,

- track ribs and end ribs, made of aluminum alloy,

- other ribs made of aluminum alloy on the A318 and A319, and made

of CFRP or aluminum alloy on the A320,

- spars made of aluminum alloy on the A318 and A319, and made of

CFRP or aluminum alloy on the A320.

- a trailing edge made in an aluminum alloy sandwich construction.

A rubbing strip (not shown) made of stainless steel is bonded onto

the outer surface of the top skin.


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OUTBOARD FLAP STRUCTURE - A318-A319-A320

The outboard flap is supported and driven by two wing tracks and

carriages (tracks 3 and 4).

The outboard flap has:





a leading edge with CFRP skin,

- a flap box with:

- skin panels with integrated stringers and spars made of CFRP,

- track ribs and end ribs made of aluminum alloy,

- other ribs made of CFRP.

- a trailing edge of aluminum alloy sandwich construction.

A rubbing strip (not shown) made of stainless steel is bonded onto

the outer surface of the top skin.


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A321 FLAPS STRUCTURE

The A321 flaps are fowler flaps with a tab on the trailing edge.

The inboard flap has:

- a leading edge and a flap box made of aluminum alloy,

- a trailing edge made in an aluminum alloy sandwich construction.




g g y

The outboard flap has:


- a flap box with:

- skin panels and integrated stringers made of CFRP,

- spars made of CFRP,

- track end ribs made of aluminum alloy,

- other ribs made of CFRP.

The tab is made of honeycomb core with a skin made of aluminum

sheet metal.

The tab is operated by a linkage system.


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SPOILERS GENERAL ARRANGEMENT

There are five spoilers on the upper surface of the wing trailing edge.

Spoiler 1 is connected to the false rear spar, inboard of the kink

position.

Spoilers 2 thru 5 are connected to the middle and outer sections of




p

the rear spar, outboard of the kink position.

A rubbing strip is attached to the trailing edge of spoilers (1 & 2 only).

It prevents damage to spoilers when flaps are retracted.


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SPOILERS STRUCTURE LAYOUT

Spoilers are a wedge-shaped structure.

The top and bottom skins, the sides and the trailing edge profile of

the spoilers are made in CFRP sandwich construction.

The spoiler hinges fittings and the actuator attachment fittings are




made of aluminum alloy.


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TRAILING EDGE DEVICES - SPOILERS STRUCTURE LAYOUT


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AILERON STRUCTURE LAYOUT

The aileron is located outboard of the outer flap and is connected to

the wing box rear spar between Ribs 22 and 27.

It is manufactured using CFRP skin (bonded to a honeycomb core in

the center area), spar and ribs.




The aileron hinge fittings and the actuator attachment fittings are made

of aluminum alloy.


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TRAILING EDGE DEVICES - AILERON STRUCTURE LAYOUT


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STRUCTURE DAMAGE IDENTIFICATION D/O (3)

GENERAL

The types of damage on metallic and composite structures are described

in SRM 51-11-00 chapter dealing with damage classification.

The table provides, the term, the cause and the description for each type

of damage.

Damage results from many causes and can be generally categorized intofour main groups:




four main groups:

- mechanical action,

- chemical or electrochemical reaction,

- thermal action or cycling,

- inherent metallurgical characteristics.


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GENERAL


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TYPES OF DAMAGE ON STRUCTURE

SCRATCH

A scratch is a line of damage of any depth and length in the material,

which causes a cross sectional area change. A sharp object is usually

the cause.





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TYPES OF DAMAGE ON STRUCTURE - SCRATCH


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TYPES OF DAMAGE ON STRUCTURE (continued)

CORROSION

Corrosion is the destruction of metal by chemical or electrochemical

effect. Refer to SRM 51-22-00 for general information concerning

corrosion.

The different types of corrosion that can occur on the aircraft are:

pitting corrosion




- pitting corrosion,

- filiform corrosion,

- intergranular corrosion,

- galvanic corrosion.

- stress corrosion,

- biological corrosion,

- fretting corrosion,

- exfoliation corrosion.


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CORROSION (CONT'D)

This page deals with:

- pitting corrosion,

- filiform corrosion,

- intergranular corrosion,- galvanic corrosion




galvanic corrosion.


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CORROSION (CONT'D)

This page deals with:

- stress corrosion,

- biological corrosion,

- fretting corrosion,- exfoliation corrosion.




e o at o co os o .


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GOUGE

A gouge is a damaged area of any size, which results in a cross

sectional area change. It is usually caused by contact with a relatively

sharp object, which produces a continuous, sharp or smooth channel

like groove in the material.





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CRACK

A crack is a partial fracture or complete break in the material.





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DENT

A dent is a damaged area, which is pushed in, with respect to its usual

contour. There is no cross sectional area change in the material. The

edges of the damaged area are smooth.





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NICK

A nick is a small decrease of material due to, for example, a knock at

the edge of a member or a skin.





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DISTORTION

A distortion is any twisting, bending or permanent strain, which results

in misalignment or change of shape. It may be caused by an impact

from a foreign object, but is usually the result of a vibration or

movement of adjacent attached components. This group includesbending, buckling, deformation, imbalance, misalignment, pinching,

d t i ti




and twisting.


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ABRASION

An abrasion is a damaged area of any size, which causes change in a

cross sectional area because of scuffing, rubbing, scrapping or other

surface erosion. It is usually rough and irregular.





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DEBONDING

Debonding is the separation of material due to an adhesive failure.





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DELAMINATION

A delamination is when a separation of plies occurs in multi-laminate

material. The material being hit or when there is a resin failure for

any other reason can cause a delamination.





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FRETTING

A fretting is a surface damage at the interface between elements of

the joints resulting from very small angular or linear movements. The

result of fretting is usually the production of fine black powder

staining.





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CREASE

A crease is a damaged area, which is pushed in or folded back on

itself. The edges of the damaged area are sharp or well specified lines

or ridges.





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MARK

A mark is a damaged area of any size where a concentration of

scratches, nicks, chips, burrs or gouges etc. is shown. You must

consider the damage as an area and not as a serie of individual

scratches, gouges, etc.





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TYPES OF DAMAGE ON STRUCTURE - MARK


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WINDOW DAMAGE IDENTIFICATION D/O (3)

GENERAL

The windows include the cockpit windows (windshields, sliding windows,

and aft fixed windows), the cabin windows and the passenger/crew door

windows.

Information dealing with different types of damage on windows, is in

page block 6xx of the relevant AMM chapters:- AMM 56-11-11 for the windshields,

- AMM 56-12-11 for the sliding windows,

AMM 56 11 12 for the aft fixed windows




- AMM 56-11-12 for the aft fixed windows,

- AMM 56-21-13 for the cabin windows,

- AMM 56-31-00 for the passenger /crew door windows.


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GENERAL


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TYPES OF DAMAGE ON COCKPIT WINDOWS

RELEVANT ATA CHAPTERS

The types of damage on cockpit windows are mentioned in:

- AMM 56-11-11 page block 6xx for windshields,

- AMM 56-11-12 page block 6xx for aft fixed windows,

- AMM 56-12-11 page block 6xx for sliding windows.





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TYPES OF DAMAGE ON COCKPIT WINDOWS (continued)

DAMAGE IDENTIFICATION

The types of damage (illustrated) on cockpit windows are:

- Crack: line type defect through the depth of the ply,

- delamination: local separation of glass and interlayer,

- interlayer microflakes: due to moisture ingress in interlayer,- burning (on windshields only): discoloration of the slip pan due to

hot corner effect,

bubbles: appear between the inner face of the outer ply and the




- bubbles: appear between the inner face of the outer ply and the

interlayer,

- burn spot: due to degradation of the heating film element,

- discoloration: due to penetration of dust or sealant.

The types of damage not illustrated are:

- scratch: line type defect in the external surface of the window causing

a cross sectional change,

- chips: flakes of glass broken from the surface and the edges of the

window,

- transparency: halos on the surface of the window can make them

less transparent,

- rain repellent fluid residue on windshields only,

- damage on the soft liner on windshields (if supplied by SPS company

only).


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TYPES OF DAMAGE ON CABIN AND PASSENGER/CREW

DOOR WINDOWS

RELEVANT ATA CHAPTERS

The cabin and passenger/crew door windows have an inner and an

outer pane. The types of damage on cabin door windows are mentioned

in AMM 56-21-13 page block 6xx and on passenger/crew doorwindows in AMM 56-31-00 page block 6xx.





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TYPES OF DAMAGE ON CABIN AND PASSENGER/CREW DOOR WINDOWS - RELEVANT ATA CHAPTERS


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TYPES OF DAMAGE ON CABIN AND PASSENGER/CREW

DOOR WINDOWS (continued)


The types of damage are:

- crazing: small cracks that go in one or all directions,

- scratch: type line defect which causes a cross sectional area change,- crack: partial fracture or complete break of the window pane,

- orange peel effect: irregular cracks on or under the surface,

- chipping: flakes of stretched acrylic broken on the edge of the pane




- chipping: flakes of stretched acrylic broken on the edge of the pane,

- delamination: slate like separation of the material,

- pitting: impact by hard particles against the surface.


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SRM D/O (3)

GENERAL

The structure repair manual is a non-customized document.

It has been prepared in accordance with Air Transport Association of

America (ATA) specification 100.

The SRM includes descriptive information as well as specific instructions

and data to perform the assessment of structural damage and to perform

repairs. The manual content is approved by the French Airworthiness

Authority DGAC ("Direction Générale de l'Aviation Civile").

For most of the damage/defect discovered on the aircraft structure, the




SRM is the first document to be used to assess the damage, to identify

the affected structure and to determine the subsequent action or repair to

be performed.


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GENERAL


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MANUAL BREAKDOWN

The SRM is divided into seven main chapters (From ATA 51 to ATA

57) and the SRI (Structure Repair Inspection).

The manual also contains an introduction chapter (Chapter 00), and some

additional information pages (HIGHLIGHTS, RECORD OF

REVISIONS...) located just at the beginning of the manual.





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MANUAL BREAKDOWN


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FRONT PAGES

GENERAL

The front pages of the manual provides general information related

to the manual itself:

- revision transmittal sheet,

- highlights,- record of revisions approved,

- record of temporary revisions,

- list of effective temporary revisions.





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FRONT PAGES - GENERAL


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FRONT PAGES (continued)

HIGHLIGHTS

The highlights chapter deals with the identification, location of the

changes within the SRM from the previous revision. The type of

change (page revised, new, deleted) is also presented.





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FRONT PAGES - HIGHLIGHTS


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INTRODUCTION

GENERAL

The introduction chapter contains all necessary information and

explanations to enable a correct use of the manual.





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INTRODUCTION - GENERAL


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INTRODUCTION (continued)

AIRPLANE ALLOCATION LIST

It also contains the airplane allocation list, giving for each MSN

(Manufacturer Serial Number), the airplane type, the aircraft rank

within the customer version, the customer, the customer abbreviation

code and the registration number.





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INTRODUCTION - AIRPLANE ALLOCATION LIST


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WEIGHT VARIANT INFORMATION

This section of the introduction chapter enables the identification of

the aircraft weight variant according to the aircraft serie, engine types,

aircraft model and the modification associated to the weight variant

changes.





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INTRODUCTION - WEIGHT VARIANT INFORMATION


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ALPHANUMERICAL INDEX

This document provides a list of all PNs (Part numbers) referenced

within the SRM and gives the related ATA chapter, the figure number,

the configuration and the item number within the figure. This list is

updated at each SRM revision. It provides a quick access to the part,to identify within the manual using the part number as a key.





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INTRODUCTION - ALPHANUMERICAL INDEX


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SRI (STRUCTURAL REPAIR INSPECTIONS)

GENERAL

For permanent repairs with inspection program, inspections are quoted

along with the repair.

Due to the amount of inspections, these requirements have been

transferred in a separate appendix to the SRM:- for more clarity of the SRM,

- for better handling of the inspection requirements.

The chapter Structural Repair Instructions (SRI) gives all necessary

inspection instructions on structural repairs and allowable damage




inspection instructions on structural repairs and allowable damage

limits, and provides the airlines with information to integrate the

additional inspections to their own maintenance program.


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SRI (STRUCTURAL REPAIR INSPECTIONS) - GENERAL


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SRI (STRUCTURAL REPAIR INSPECTIONS) (continued)

GENERAL (CONT'D)

If an inspection is needed, an Inspection Instruction Reference (IIR)

will be specified within the related repair. The operator shall then

refer to the SRI chapter to find out all the inspection instructions using

the IIR as a key.





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SRI (STRUCTURAL REPAIR INSPECTIONS) - GENERAL (CONT'D)


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I.I.R. (INSPECTION INSTRUCTION REFERENCE)

The explanation of the IIR is given at the beginning of the SRI chapter.





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SRI (STRUCTURAL REPAIR INSPECTIONS) - I.I.R. (INSPECTION INSTRUCTION REFERENCE)


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SRI CONTENTS

At the beginning of the SRI chapter, a table of contents provides the

list of repairs concerned by special inspections requirements.





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SRI (STRUCTURAL REPAIR INSPECTIONS) - SRI CONTENTS


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SRI CONTENTS (CONT'D)

Within the introduction section of the SRI, the "Inspection Instruction

Schemes" table enables the operator to locate the inspection instruction

within the SRI (for inspection ref. 53-41-11-2-001-00, instructions

are to be found in paragraph 1.A.).





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SRI (STRUCTURAL REPAIR INSPECTIONS) - SRI CONTENTS (CONT'D)


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SRI CONTENTS (CONT'D)

Within the concerned paragraph, the inspection instruction are divided

into three main parts:

- the general information part, which reminds the inspection location,

- the inspection information table (A/C concerned, inspection

threshold, interval, required inspection methods),

- the inspection areas illustrations.





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SRI (STRUCTURAL REPAIR INSPECTIONS) - SRI CONTENTS (CONT'D)


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SRI CONTENTS (CONT'D)-INSPECTION AREAS

ILLUSTRATIONS (IF ANY - E.G. 53-00-11)

The areas to be inspected are defined on the illustration, with the

identification of the applicable inspection methods and applicable

NTM procedures.





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SRI (STRUCTURAL REPAIR INSPECTIONS) - SRI CONTENTS (CONT'D)-INSPECTION AREAS ILLUSTRATIONS (IF ANY - E.G. 53-00-11)


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MANUAL GENERAL USAGE PROCEDURE

When damage is discovered, the first step is to evaluate, classify and

accurately measure it. The SRM chapter 51-11-XX provides useful

information to perform this assessment in the best conditions (damage

definitions and classification, classification of structures, allowable

damage definitions and damage/defect reporting process).





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MANUAL GENERAL USAGE PROCEDURE


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MANUAL GENERAL USAGE PROCEDURE (continued)

CONT'D

The next step is the full identification of the affected area/structure.

This is achieved using the identification page block (pages 01-99) of

the related specific chapter/section (52-57). According to the original

structure data and the actual damage characteristics, it is then possible

to determine whether the damage is within the defined allowable limits

or not. This is done using the allowable damage page block (pages

101-199) of the related specific chapter/section. If the damage is within

the allowable limits, the subsequent actions are generally a slight

rework and a re-protection of the affected area using the standard




rework and a re protection of the affected area, using the standard

procedures of chapter 51.


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MANUAL GENERAL USAGE PROCEDURE - CONT'D


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MANUAL GENERAL USAGE PROCEDURE (continued)

CONT'D

If the damage is above the limits, you must check whether a repair is

available and/or applicable within the repair page block (pages

201-999).





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MANUAL GENERAL USAGE PROCEDURE - CONT'D


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NUMBERING SYSTEM AND PAGE BLOCK ALLOCATION

Each subject, within the SRM, is identified using a three-element

numbering system chapter/section and sub-section:

- the first element designates the chapter which is assigned by the ATA

spec. 100,

- the second element designates the section within the chapter. The first

digit is assigned by the ATA spec. 100. The second digit is assigned byAirbus S.A.S,

- the third element identifies the subsection (subject) within the section

and is assigned by Airbus S.A.S.

A standard page block allocation is used for all SRM chapters:




- pages 1 to 99 for structure identification,

- pages 101 to 199 for allowable damage,

- pages 201 to 999 for repairs.


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NUMBERING SYSTEM AND PAGE BLOCK ALLOCATION


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CHAPTER 51 (STRUCTURE GENERAL)

Information of a general nature or information applicable to more than

one chapter, is included in chapter 51.

NOTE: Note: the entry point within the SRM is always the specific

chapter 52 to 57, depending on the damage location.





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CHAPTER 51 (STRUCTURE GENERAL)


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CHAPTER 51 (STRUCTURE GENERAL) (continued)

CONTENTS

The table of contents is detailed in the following illustrations.





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CHAPTER 51 (STRUCTURE GENERAL) - CONTENTS


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CONTENTS (CONT'D)

Chapter 51 table of contents (2/3).





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CHAPTER 51 (STRUCTURE GENERAL) - CONTENTS (CONT'D)


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CONTENTS (CONT'D)

Chapter 51 table of contents (3/3).





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CHAPTER 51 (STRUCTURE GENERAL) - CONTENTS (CONT'D)


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CHAPTER 52 TO 57 CONTENTS

LAYOUT

Chapters 52 to 57 have the same layout, which conforms to the defined

page block allocation system (PB 1 to 99-identification, PB 101 to

199-allowable damage, PB 201 to 999-repairs). In addition, a table

of contents and a Service Bulletin (SB) list are provided at the

beginning of each chapter. Depending on the chapters, theModification/Service Bulletin list is to be found either at the chapter

level, or main section level.





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CHAPTER 52 TO 57 CONTENTS - LAYOUT


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CHAPTER 52 TO 57 CONTENTS (continued)

SERVICE BULLETIN LIST

Located at the beginning of each chapter, the SB list is a list of all

service bulletins referenced within the chapter.





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CHAPTER 52 TO 57 CONTENTS - SERVICE BULLETIN LIST


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SERVICE BULLETIN LIST (CONT'D)

The user can refer to this list to get more information regarding a

specific service bulletin in terms of:

- revision status of the SB,

- date of introduction within the SRM,

- description.





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CHAPTER 52 TO 57 CONTENTS - SERVICE BULLETIN LIST (CONT'D)


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MODIFICATION/SERVICE BULLETIN LIST

Located at:

- door level for chapter 52,

- fuselage section level for chapter 53,

- chapter level for chapter 54,

- main assembly level for chapter 55,- wing section level for chapter 57, this list is provided to enable the

user to determine the effectivity of the modification/SB.





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CHAPTER 52 TO 57 CONTENTS - MODIFICATION/SERVICE BULLETIN LIST


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MODIFICATION/SERVICE BULLETIN LIST (CONT'D)

This list provides, for a given modification number, its associated

suffix and the aircraft standard, and the effectivity expressed in MSN.





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CHAPTER 52 TO 57 CONTENTS - MODIFICATION/SERVICE BULLETIN LIST (CONT'D)


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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99)

EXAMPLE : COMPOSITE STRUCTURES

In the identification pages, the individual parts of the major

components are illustrated and listed in tabular form. Each

identification topic begins with an introduction page, which includes

a general information paragraph. The item number is the key between

the illustration and the identification table. For composite structures,the illustration provides identification of individual layers, orientation

and materials. A ply (layer) orientation reference is also given.





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - EXAMPLE : COMPOSITE STRUCTURES


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(continued)

EXAMPLE : METALLIC STRUCTURES

For metallic structure such as fuselage skin panels, the different

material thicknesses are provided using letter codes or shaded areas

as a key to the thickness tables. The associated identification table

provides the additional material, part number modification statusinformation.





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - EXAMPLE : METALLIC STRUCTURES


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(continued)

IDENTIFICATION TABLE DETAILED

The typical identification table associated with the illustration(s)

contains different columns:

- ITEM,

- NOMENCLATURE,- SPECIFICATION AND/OR SECTION CODE,

- THICKNESS IN MM (IN.) AND/OR PARTNUMBER,

- IC: INTERCHANGEABILITY,

- ACTION OR REPAIR,

- STATUS (MOD/PROP) SB/RC (MOD/PROP:




Modification/Proposal, RC: Recordable Concession),

In addition: the relevant assembly drawings are listed at the end of

the table (ASSY DRAWINGS:...........).


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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - IDENTIFICATION TABLE DETAILED


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(continued)

IDENTIFICATION TABLE DETAILED (ITEM COLUMN)

The item number is the link with the associated illustrations. This

column also indicates the different evolutions of the same item (with

a suffix letter: 1A, 1B,..) compared to the basic version (without

suffix). Each evolution is linked to a production modification givenin the column ''status (MOD/PROP)''.





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - IDENTIFICATION TABLE DETAILED (ITEM COLUMN)


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(continued)


(SPECIFICATION/SECTION CODE COLUMN)

This column provides information related to the type of material used

(material specification), and when possible/available, the appropriate

standards. To get more information regarding the materialspecifications the user can refer to the SRM chapter 51:

- 51-31-00 for metallic materials,

- 51-33-00 for non-metallic materials (composite materials).

When the appropriate standard is shown, the user can refer to the

Airbus S.A.S. Standard Manual(SM).





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - IDENTIFICATION TABLE DETAILED (SPECIFICATION/SECTION CODE COLUMN)


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(continued)

CHAPTER 51-31-00 EXAMPLE

This example shows the way to proceed to find out the material

information within the chapter 51-31-00 for metallic structure, starting

with the material specification. A first list provides the table to be

used (e.g. table 4 for 3.1364T42).





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - CHAPTER 51-31-00 EXAMPLE


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(continued)

CHAPTER 51-31-00 EXAMPLE (CONT'D)

The material table provides, for a given material specification, the

European substitute (if used by the other European manufacturers and

the United States substitutes).





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(continued)

STANDARD MANUAL EXTRACT

The standard manual provides the additional information linked to a

given standard number.





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - STANDARD MANUAL EXTRACT


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(continued)


(THICKNESS/PART NUMBER COLUMN)

The Part Number (PN) corresponding to the structural component is

shown in this column. Within the PN, the first nine characters give

the detailed drawing number of the component, which is an entry keyto the airbus drawing set. In general the "as drawn" parts are Left

Hand (LH) and are provided with an even part number (e.g.: ...202).

In the SRM, the identification table always states the LH part number

on the first line and the RH part number (when indicated) on the

second line.





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - IDENTIFICATION TABLE DETAILED (THICKNESS/PART NUMBER COLUMN)


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(continued)

IDENTIFICATION TABLE DETAILED (IC COLUMN)

The IC column provides the interchangeability status of the parts:

- 01: one way interchangeable (post-mod part to be used to replace

pre-mod part),

- 02: two ways interchangeable (post-mod or pre-mod parts can beused),

- 03: no interchangeability.





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - IDENTIFICATION TABLE DETAILED (IC COLUMN)


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(continued)

IDENTIFICATION TABLE DETAILED (ACTION OR

REPAIR COLUMN)

This column gives indication concerning the action or repair to be

performed. For existing general or specific repairs, the

chapter/section/sub-section is inserted. For a recommendation toreplace the part, the word "REPLACE" is inserted. Where left blank,

a case-by-case assessment has to be performed to determine the

relevant corrective action (part replacement, repair as per SRM or

specific repair as per Airbus SAS definition).





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - IDENTIFICATION TABLE DETAILED (ACTION OR REPAIR COLUMN)


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(continued)

IDENTIFICATION TABLE DETAILED (STATUS

MOD/PROP SB/RC COLUMN)

Linked with the Item column, the STATUS MOD/PRP, SB/RC column

gives the modification or service bulletin driving the different

evolutions of a structural component (listed in column Item). A prefixletter is used to identify the status BEFORE ("B" letter) or AFTER

("A" letter) SB or Modification. The suffix letter (A, B, C, D...)

indicated at the end of the MOD/PROP (and column "S" of the

MOD/SB list) shows the different effectivity within the same

MOD/PROP number.





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(continued)


MOD/PROP SB/RC COLUMN (CONT'D))

To find the relevant effectivity linked to a modification shown in the

STATUS column, the user must refer to the modification/service

bulletin list.

NOTE: the status before or after modification/SB and the relevant

modification solution (suffix letter) should not be forgotten.

Within the modification/service bulletin list, the effectivities

are expressed in MSN.

to find the relationship between the customer version number

( AFR 01 0016) d th MSN th f t th




(e.g.. AFR 01 0016) and the MSN, the user can refer to the

airplane allocation list of the introduction chapter of the

SRM.


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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - IDENTIFICATION TABLE DETAILED (STATUS MOD/PROP SB/RC COLUMN

(CONT'D))


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(continued)


MOD/PROP SB/RC COLUMN (CONT'D))

When the modification is linked to a SB, the SB number is also

mentioned, below the modification number. The SB list located at the

beginning of the chapter can be used to get more information.





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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99) - IDENTIFICATION TABLE DETAILED (STATUS MOD/PROP SB/RC COLUMN

(CONT'D))


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CONFIGURATIONS

GENERAL

When modification or service bulletin are extensive and the preceding

method of reflecting effectivity becomes cumbersome, thus distracting

from the continuity of subject matter, additional page blocks are

established. These additional page blocks are further identified by the

introduction of a configuration code (CONFIG-1, CONFIG-2), etc...)following the chapter/section/sub-section. Configuration codes are

always in ascending, sequential numerical order, i.e., CONFIG-1,

CONFIG-2, CONFIG-3, etc... The first example shown below

illustrates the change in the A320 vertical stabilizer spar box design

and manufacturing following modification 26500K4924H.





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CONFIGURATIONS - GENERAL


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CONFIGURATIONS (continued)

EXAMPLE

Here is an example of different configurations. The modification of

the design and production process of the carbon fiber vertical stabilizer

spar box leads to two configurations within the chapter 55. The

configuration number is indicated at the bottom of the related pages.

On this example, as highlighted, the configuration 1 is applicablebefore modification 26500K4924H. Configuration 2 is applicable

after modification 26500K4924H. It is of the operator duty to select

first in which configuration the concerned aircraft is, before going

further in the SRM investigation. This is done using the service

bulletin/modification list located at the beginning of the chapter or

section.





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CONFIGURATIONS - EXAMPLE


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ALLOWABLE DAMAGE PAGE BLOCK (PB 101)

GENERAL

The information to be found within allowable damage page block

enables the operator to define whether a damaged airplane may be

returned into service without repair. An allowable damage permitted

has no significant effect on the strength or fatigue life of the structure,

which must still be capable of fulfilling its function. Allowable damagemay require minimal rework such as cleanup or drilling of stop holes.

Basically the allowable page block contains different page types:

- general information pages,

- damage criteria tables,

- paragraph for each type of damage,

- damage measurement procedure,

- damage localization (zoning) figures,




g g g

- allowable damage diagram.


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ALLOWABLE DAMAGE PAGE BLOCK (PB 101) (continued)

ALLOWABLE DAMAGE INFORMATION

After a careful reading of the first information page(s), the first step

in the allowable damage determination, consists in the use of the

damage criteria table.





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DAMAGE CRITERIA TABLE

The damage criteria table enables the operator to determine the relevant

paragraph (e.g. 4A) that should be used, depending on the type of

damage (e.g.: allowable rework), and the affected structure (e.g. skin

plates).





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ALLOWABLE DAMAGE EFFECTIVITY

For some allowable damage (e.g. dent allowable damage), the

allowable damage effectivity per weight variant must be checked.

The table (e.g. table 102) at the beginning of the allowable damage

relevant paragraph shall be used. The allowable damage information

contained within the related paragraph is justified and applicable for

the listed weight variant only. If not, contact Airbus.





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WEIGHT VARIANT DETERMINATION

To define the "as delivered" weight variant of a concerned MSN, the

first step is to refer to the airplane allocation list of the introduction

chapter, which gives the corresponding type (e.g.: A320-232). Then,

refer to the weight variant identification list, also in the introduction

chapter.





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WEIGHT VARIANT DETERMINATION (CONT'D)

The weight variant identification list gives, for a given aircraft type,

the different possible "as delivered" weight variants. The next step

will be to determine which one of these possible weight variants is

applicable to the concerned MSN.





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To determine the applicable "as delivered" weight variant, the operator

must check which of the associated modification(s) is effective on the

concerned MSN. The Aircraft Inspection Report (AIR) can be used

for this purpose. Once found, the corresponding weight variant is

considered as the weight variant of concerned MSN at delivery.





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The "as delivered" weight variant may change after delivery, following

the embodiment of a service bulletin. It is the operator's responsibility

to check the embodiment of referenced SB, in order to determine the

relevant weight variant for the affected MSN. The information is given

into Table 2 "Service Bulletin/Weight Variant List".





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The table 2-"Service Bulletin/Weight Variant List" provides for the

related aircraft type (e.g.: A320-232) the different service bulletin

applicable. For the given MSN, the operator must check whether one

of the possible SB has been embodied or not. If yes, the weight variant

of the aircraft becomes the weight variant given by the table.





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For aircraft which have been subject to weight variant change through

a SB embodiment: the weight variant information to be used to identify

the effectivity of the given allowable damage (or repair information),

is the heaviest weight variant that the subject aircraft has been operated

with, since its delivery.





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ALLOWABLE DAMAGE DETERMINATION (CONT'D)

The allowable damage diagram to be used, generally depends on the

location of the damage on the concerned structure.

Illustrations are used to locate the damage and thus to define the

relevant diagram to be used (e.g. refer to diagram 102).





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ALLOWABLE DAMAGE PAGE BLOCK (PB 101) - ALLOWABLE DAMAGE DETERMINATION (CONT'D)


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ALLOWABLE DAMAGE DIAGRAM EXAMPLE

As a last step of the damage investigation, the use of the allowable

damage diagrams provides the necessary information concerning the

actions to be performed, depending on damage extent.





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(CONT'D)

Provided that no cracks have been detected, this first area of the

allowable damage rework diagram defines typical allowable damage

without any time limits. The surface protection needs to be restored.





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ALLOWABLE DAMAGE PAGE BLOCK (PB 101) - ALLOWABLE DAMAGE DIAGRAM EXAMPLE (CONT'D)


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(CONT'D)

This second area for reworks depth between 10 % and 25 % of the

nominal thickness, also defines allowable damage but with a time

limit.

In this example a repair has to be performed before 3000 flights at

the latest. The surface protection has to be restored.





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(CONT'D)

This area is defined for reworks depth between 25 % and 40 % of the

nominal thickness. It is stil l allowable damage, provided that no crack

is detected. But a repair has to be performed before 50 flights at the

latest.





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(CONT'D)

Stop drill and apply high-speed tape for one flight pressurized or

unpressurized only, or do a temporary repair as per 53-00-11 figure

210.

Temporary repair has to be replaced by a final repair within 2500

flights.





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(CONT'D)

For damage located in this area, the damage shall be stop drilled before

a ferry flight without cabin pressure. Install high-speed tape before

the ferry flight. A repair has to be performed.





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(CONT'D)

For damage located in this area, a repair has to be performed

immediately or a ferry flight may be allowed upon manufacturer's

authorization.





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REPAIRS PAGE BLOCK (PB 201)

GENERAL

The Repairs Page Block (PB 201), contains necessary information to

carry out permissible repairs.

Each of the repairs is described with illustrations and procedure

instructions, which includes repair applicability data and repair

material lists.





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REPAIRS PAGE BLOCK (PB 201) - GENERAL


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REPAIRS PAGE BLOCK (PB 201) (continued)

LIST OF AVAILABLE REPAIR SCHEMES

At the beginning of each repair page block, a list of available repair

schemes is provided for a quick assess to the repairs.

The repairs can be located in the general section of chapter 53 (e.g.:

53-00-11 for standard fuselage skin repairs), or directly covered within

the related section when the repair is specific (e.g. paragraph 5A, Fig.

201).





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REPAIRS PAGE BLOCK (PB 201) - LIST OF AVAILABLE REPAIR SCHEMES


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REPAIR INSTRUCTIONS

The relevant paragraph of the repair instructions contain information

related to:

- the applicability of the concerned repair,

- general reminders linked to repair principles and rules.

Note: an IIR appears within the repair instructions.





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REPAIRS PAGE BLOCK (PB 201) - REPAIR INSTRUCTIONS


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REPAIR INSTRUCTIONS (CONT'D)

All the materials (repair materials and consumable materials) are listed

within the concerned repair instructions. Consumable materials are

call-up using their Consumable Material List (CML) code (e.g.

09-013). For more information on these materials, the user can refer

to the SRM chapter 51-35-00 consumable materials and/or the CML

document.





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this SRM chapter(51-35-00) deals with consumable materials.





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Information concerning the consumable materials can also be found

in the CML.





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The repair instruction lists all the steps of the repair, with references

to the standard processes and practices covered within the chapter 51

when required.





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REPAIR ILLUSTRATIONS

Note: an IIR appears within the repairs illustrations.





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REPAIRS PAGE BLOCK (PB 201) - REPAIR ILLUSTRATIONS


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WINDOW DAMAGE ASSESSMENT D/O (3)

GENERAL

The windows are:

- the cockpit windows ( windshields, sliding windows, and aft fixed

windows),

- the cabin windows and,

- the passenger/crew door windows.

To find the corrective actions for each type of defect, refer to the page

block 6xx of the relevant AMM chapters:- AMM 56-11-11 for the windshields,

- AMM 56-12-11 for the sliding windows,

- AMM 56-11-12 for the aft fixed windows,

- AMM 56-21-13 for the cabin windows,

- AMM 56-31-00 for the passenger/crew door windows.





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GENERAL


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INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS

Refer to the relevant AMM chapter for each type of cockpit windows.





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(continued)

INVESTIGATION OF DAMAGE ON WINDSHIELDS

Refer to AMM 56-11-11 pages block 6xx to get the corrective actions

for the following types of defect on windshields:

- cracks,

- scratches,

- chips,- delaminating,

- discoloration,

- interlayer micro flakes,

- bubbles,

- burn spot,

- transparency,

- rain repellent fluid residue,

- damage on the soft liner (if supplied by SPS company only).





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INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS - INVESTIGATION OF DAMAGE ON WINDSHIELDS


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(continued)

INVESTIGATION OF DAMAGE ON SLIDING WINDOWS


for the following types of defect on sliding windows:

- cracks,

- scratches,

- chips,- delaminating,

- bubbles,

- discoloration or burning,


- transparency,

- crazing,

- burn spots.





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INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS - INVESTIGATION OF DAMAGE ON SLIDING WINDOWS


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(continued)

INVESTIGATION OF DAMAGE ON AFT FIXED

WINDOWS


for the following types of defect on aft fixed windows:

- cracks,

- scratches,- chips,

- delaminating,

- bubbles,

- discoloration or burning,

- transparency,


- burn spots.





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INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS - INVESTIGATION OF DAMAGE ON AFT FIXED WINDOWS


51 - STRUCTURE


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INVESTIGATION OF DAMAGE ON CABIN WINDOWS

AND PASSENGER/CREW DOOR WINDOWS

Refer to the relevant AMM for the cabin and passenger/crew door

windows.





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INVESTIGATION OF DAMAGE ON CABIN WINDOWS AND PASSENGER/CREW DOOR WINDOWS


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AND PASSENGER/CREW DOOR WINDOWS (continued)


Refer to AMM 56-21-13 page block 6xx to get the allowable damages

for the following types of defect on cabin windows:

- delaminating,

- scratches,

- pitting,- crazing,

- crazing with bulging,

- bulging,

- orange peel effect,

- chipping,

- cracks,

- vent hole damage.





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INVESTIGATION OF DAMAGE ON CABIN WINDOWS AND PASSENGER/CREW DOOR WINDOWS - INVESTIGATION OF DAMAGE

ON CABIN WINDOWS


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AND PASSENGER/CREW DOOR WINDOWS (continued)

INVESTIGATION OF DAMAGE ON PASSENGER/CREW

DOOR WINDOWS

Refer to AMM 56-31-00 page block 6xx to get the allowable damage

for the following types of defect on passenger/crew door windows:

- scratches,

- crazing,- bulging,

- crazing with bulging,

- delaminating.





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INVESTIGATION OF DAMAGE ON CABIN WINDOWS AND PASSENGER/CREW DOOR WINDOWS - INVESTIGATION OF DAMAGE

ON PASSENGER/CREW DOOR WINDOWS


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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

INTRODUCTION

The purpose of this example is to present, the complete procedure to be

followed when a damage is discovered, from the damage mapping draft

to the final structure damage assessment. This example was chosen as it

represents one of the more usual types of damage on an A/C and gives

an in depth investigation with all the different stages.





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INTRODUCTION


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IDENTIFICATION OF THE DAMAGE

The damage is located onto the fuselage skin, thus, all the information

regarding the identification of the part, allowable damage and repair, if

any, are to be found within the chapter 53 of the SRM. Information

concerning the damage classification and reporting are to be found within

the SRM chapter 51-11-00. The concerned damage is a dent with no

visible crack. At this stage, take visual reference to facilitate damage

location. Such as, a forward or aft passenger door, or a cargo door, above

or below cabin floor level at stringer (Stgr) 23, close to a longitudinal or

circumferential joint, etc...). If the dent is close to a rivet row, an internal

visual inspection is required to determine whether the internal structure

(frame, stringer, etc...) is also damaged or not.





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MAPPING

Using SRM 51-11-13 as a guide, the maximum information should be

taken from the aircraft before starting any assessment (measurement and

location of the maximum depth, distance of dent edges to nearest fastener

rows, existing closest skin joints or any other visible structure that will

help in the detailed location of the damage, etc...).





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MAPPING


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MAPPING (continued)

CONT'D

Using the data collected from the A/C, the mapping should be

completed by determining the exact location (in terms of frame

numbers and stringer numbers). For this purpose, refer to the beginning

of the chapter 53 (fuselage).





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MAPPING - CONT'D


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MAPPING (continued)

CONT'D

The illustration of chapter 53-00-00 enables the operator to determine

the circumferential joint related frame numbers.





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MAPPING - CONT'D


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MAPPING (continued)

CONT'D

Using the frame identification illustration of chapter 53-00-00, and

the data collected during the damage mapping, the frames surrounding

the damage can be determined. According to the mapping information,

the damage is located between the first and the second frame after the

circumferential joint located at Fr 24. Consequently, the damage is

located between Fr 25 and 26.





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MAPPING - CONT'D


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MAPPING (continued)

CONT'D

To complete the damage location, the stringers surrounding the damage

also need to be determined. For this purpose the "General panel

identification" illustrations, proposed within chapter 53-00-00 can be

used. According to the data collected on the A//C and the location of

the damage from the existing longitudinal skin joints, the affected

panel can be determined. For this example the damage is located on

panel 7 - lower side shell - located between Stgr 18LH & 32LH, and

Fr 24 & 35.





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MAPPING - CONT'D


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MAPPING (continued)

CONT'D

The information collected can be reported onto the damage mapping.

The damage is located between Fr 25 and 26. The stringer number

corresponds to the longitudinal skin joint from which the damage has

been located. Nevertheless, the exact stringer numbers surrounding

the damage need to be confirmed. For this purpose, the information

provided in the identification page block of the concerned panel has

to be used.





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MAPPING - CONT'D


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DETAILED IDENTIFICATION

The "fuselage section division" illustration of chapter 53-00-00 used

before enables the definition of the affected section: "Forward fuselage"

- section 13/14 - chapter 53-20-00. The general illustration of 53-20-00

identifies the main structural arrangement of the forward fuselage. Skin

plates are part of the "MAIN STRUCTURE" covered by section 53-21-00.





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DETAILED IDENTIFICATION (continued)

CONT'D

Following SRM 53-21-00 guidelines, the figure shows that the skin

panels (skin plates) are item number 1. The illustration associated

nomenclature informs us that the full identification of the skin panels

(skin plates) are covered by SRM 53-21-11.





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DETAILED IDENTIFICATION - CONT'D


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CONT'D

All the skin panels (plates) of the forward fuselage are listed within

the nomenclature located at the front page of SRM 53-21-11. Using

the information collected before (affected panel: lower side panel -

left, between Fr 24 & 35 and Stgr 18 & 32), the nomenclature provides

the figure number we have to refer to: "Skin plates - LWR parts LH

Fr 24 to Fr 35: REFER TO Figure 1".





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CONT'D

The figure 1 identifies two different panels (view A and C). The view

A concerns the skin panel located from Stgr 18LH to 32LH. The view

C concerns the skin panel located from Stgr 32LH to 41LH. According

to the damage mapping, the view A is concerned. The damage has

been located on panel 7 (between Stgr 18LH and 32LH).





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CONT'D

The view A, identifies all the different items which are part of the

panel (e.g. crack stoppers, doublers...), and a view indication identifies

the skin itself for more details (View D). The view D identifies the

different material thicknesses (letter codes), and all the stringer

locations.

There are two different panel configurations illustrated, showing the

basic version of the panel and an other possible version effective after

the embodiment of production modification(s). Modification numbers

are indicated at the bottom of the page. The next step of the

investigation is to define which of these panels is installed on the

concerned A/C.





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CONT'D

To identify the actual panel, the modification numbers indicated at

the bottom of the page have to be compared with the service

bulletin/modification list, located at the beginning of chapter 53-20-00.

The purpose is to check their effectivity in terms of Manufacturer

Serial Number (MSN).





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CONT'D

The view D of panel on figure 1, sheet 5 is valid after Modifications

(MODs) 27117P5234 or 2729P5353. Checking the Modification / SB

List at the beginning of chapter 53-20-00, MSN 2057 doesn't appear

in this list of MSN proposed for each of the modification. So the panel

installed on the A/C is a basic version, then, refer to view D figure 1,

sheet 3.





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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)DETAILED IDENTIFICATION (continued)

CONT'D

The damage is located between Fr 25 and 26, and is located between

the fourth and the fifth stringer from Stgr 18 LH (longitudinal skin

joint reference).

This information can be reported onto the illustration and gives:

- the material thickness of the area (code B, giving 1.4 mm (0.055

in)),

- the stringer location: the damage is located between Stgr 22LH and

23LH.





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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)DETAILED IDENTIFICATION (continued)

MAPPING (FINALIZATION)

The damage mapping can now be completed with the stringer numbers

and the nominal skin thickness in the dented area. The damage

assessment using the allowable damage page block is the next step.





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DETAILED IDENTIFICATION - MAPPING (FINALIZATION)


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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)DAMAGE ASSESSMENT

GENERAL

To start the damage assessment refer to the page block 101 of the

relevant chapter/section (53-21-11), and start to read carefully the

procedure. A special attention shall be paid to the notes and cautions.





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DAMAGE ASSESSMENT - GENERAL


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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)DAMAGE ASSESSMENT (continued)

WEIGHT VARIANT

A caution note indicates that the allowable damage effectivity per

A/C weight variant may have to be verified. The weight variant is a

criterion which is defined for each model of A/C and depending on

its Maximum Take Off Weight (MTWO), Maximum Landing Weight

(MLW), Maximum Zero Fuel Weight (MSFW). The allowable damage

limits are defined per weight variant and for a same model. The weight

variant can change, depending on the modification or Service Bulletin(SB) embodiment status. The actual weight variant of the affected

A/C has to be known before starting the assessment. Because of the

modifications, which could be embodied on the A/C, only the airline

engineering department shall give you this information. The actual

weight variant shall be compared with the data given in a table at the

beginning of allowable damage related paragraph.





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DAMAGE ASSESSMENT - WEIGHT VARIANT


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DAMAGE CRITERIA

A second caution note indicates that in some cases, an inspection may

be required to check for crack, even if the damage is determined as

being allowable.





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DAMAGE ASSESSMENT - DAMAGE CRITERIA


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DAMAGE CRITERIA (CONT'D)

Check the applicability of the allowable damage for dents.





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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D)


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To keep on with the damage assessment procedure, a note asks the

operator to refer to the damage criteria table 101.





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The allowable damage description/criteria table (101), shows two

types of dents:

- dent* referring to paragraph 4B,

- dent ** referring to paragraph 4C.

Note that an Inspection Instruction Reference (IIR) is indicated for

dents*. The first step is to define which paragraph is applicable to

reported dent. Dents are considered as fulfilling nearness/form criterionor out of nearness/form criterion, in accordance with their geometry

and their proximity to the nearest adjacent internal structure elements.

This must be determined according to the parameters defined in figure

104.





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1





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To define whether the dent fulfils or not the nearness/form criterion,

two criteria have to be checked:

- the first criterion consists in checking the smallest distance measured

from the dent edge to any fastener row (frame, stringer) distance B.

This distance should be minimum 15 mm (0.59 in),

- the second criterion consists in comparing the depth of the dent (D)

with the smallest distance measured from the deepest point of the dentto the closest adjacent structure (distance A).

The depth of the dent should be maximum 10% of the distance A. If

one of these criteria is not met, the dent **, and thus paragraph 4C

(dent not fulfilling criteria) should be taken to keep on with paragraph

4B.


1




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1





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The second criterion consists in comparing the depth of the dent (D)

with the smallest distance, measured from the deepest point of the

dent to the closest adjacent structure (distance A). If no access from

inside, the measurement is taken from outside, from the deepest point

of the dent to the closest fastener row (distance X). The distance A

will become distance X - 15 mm, which is the average considered

edge margin.


1




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01





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Check for the first criterion to be fulfilled. B distance : minimum 15

mm. The smallest distance measured between the edge of the dent

and the surrounding fastener rows is 29 mm, which is higher than 15

mm. The first criterion is met.


01




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01





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Check for the second criterion to be fulfilled: the depth of the dent

should be maximum 10% of the distance A. The distance measurement

has been done from outside: the smallest distance between the deepest

point of the dent and the surrounding fastener row is 66 mm. Since

measured from outside, distance A = 66 mm - 15 mm = 51 mm; 10%

of A = 5.1 mm. The second criterion is also fulfilled since the depth

of the dent (D = 4.5 mm) is smaller than 10% of A.


01




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01





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The dent fulfils "nearness/form criterion", then refer to paragraph 4.B.


001




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001





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As mentioned in a caution at the beginning of the allowable damage

pages, the allowable damage applicability have to be checked, using

the weight variant table (table 102) given at the beginning of the

paragraph. The information coming from the airline-engineering

department shows that the MSN 2057 is at weight variant 001.

Checking the table 102, weight variant 001 is included in, and thus

the following allowable damage information can be used.


001




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001





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Compare the dents in accordance with diagram 103.


001




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0001





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ALLOWABLE DENT DIAGRAM

The skin thickness in the dented area, and the depth of the dent, are

the keys to get into to diagram. You must refer to the data collected

before (damage mapping).


0001




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0001



DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM


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ALLOWABLE DENT DIAGRAM (CONT'D)

The skin thickness in the dented area is 1.4 mm (found in the

identification pages). The depth of the dent is 4.5mm (measured from

the A/C damage mapping). These two values are plotted onto the

diagram, which defines a point. The area where this point is located

defines the subsequent actions to be performed. For the concerned

dent, the actions to be performed are as follow: "check damage for

cracks by detailed visual examination. If clear, repair within 3000FC". Provided that no crack is detected by detailed visual inspection,

the dent is considered as an allowable damage with a time limit

(temporary allowable damage). The A/C can be released. But a repair

will have to be done before 3000 Flight Cycles (FC).


0001




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DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM (CONT'D)


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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)INTRODUCTION

The purpose of this example is to present you the complete procedure to

be followed when a damage is discovered, from the damage mapping

draft to the final structure damage assessment. This example was chosen

as it represents one of the more usual types of damage on the A/C and

enables to make an in depth investigation with all the different stages.


00001




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00001



INTRODUCTION


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any, are to be found within the chapter 53 of the SRM. All the information

regarding the damage classification or the rework, if any, are to be found

within SRM chapter 51. The applicable damage is a scratch with no

visible crack. At this stage: take visual reference to facilitate damage

location. Such as, a forward or an aft passenger door, or a cargo door,

above or below the cabin floor level at stringer (Stgr) 23, near a

longitudinal or circumferential skin joint, etc...

If the scratch is near a rivet row, an internal visual inspection is required

to determine whether the internal structure (frame, stringer, etc...) is also

damaged or not.


00001




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00001





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MAPPING

Using the SRM 51-11-13 as a guide, the maximum information should

be taken from the A/C before starting any assessment. (measurement and

location of the maximum depth, distance of rework edges to the nearest

fastener rows, existing closest skin joints or any other visible structure

that will help in the detailed location of the damage, etc...).


00001




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000001



MAPPING


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MAPPING (continued)

CONT'D

Using the data collected from the A/C, the mapping should be

completed by determining the exact location (in terms of frame

numbers and stringer numbers). For this purpose refer to the beginning

of chapter 53 - fuselage.


000001




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000001



MAPPING - CONT'D


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MAPPING (continued)

CONT'D

The illustration of chapter 53-00-00 enables the operator to determine

the circumferential skin joints related frame numbers.


000001




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000001



MAPPING - CONT'D


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MAPPING (continued)

CONT'D



the damage can be determined.

According to the mapping information, the damage is located between

the first and the second frame after the circumferential joint located

at frame (Fr) 64. Consequently the damage is located between Fr 62

and 63.


000001




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0000001



MAPPING - CONT'D


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MAPPING (continued)

CONT'D


also need to be determined. For this purpose, the "General Panel

Identification" illustrations, proposed within chapter 53-00-00 can be

used. According to the data collected on the A/C and the location of

the damage from the existing longitudinal skin joints, the affected

panel can be determined. For this example, the damage is located on

panel 8 - lower side shell - located between Stgr 18 & 32, and Fr 47& 64.


0000001




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0000001



MAPPING - CONT'D


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MAPPING (continued)

CONT'D

The collected information can be reported onto the damage mapping.

The damage is located between Fr 62 and 63. The stringer number

corresponds to the longitudinal skin joint from which the damage has

been located. Nevertheless, the exact stringer numbers surrounding

the damage need to be confirmed. For this purpose, the information

provided in the identification page block of the concerned panel has

to be used.


0000001




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0000001



MAPPING - CONT'D


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The "Fuselage Section Division" illustration of chapter 53-00-00 used

before, enables the definition of the affected section: rear fuselage -

section 17 - chapter 53-40-00". The general illustration of 53-40-00

identifies the main structural arrangement of the forward fuselage. The

skin plates are part of the main structure covered by the section 53-41-00.


00000001




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00000001





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CONT'D


panels (skin plates) are item number 5. The illustration associated

nomenclature informs us that the full identification of the skin panels

(skin plates) are covered by SRM 53-41-11.


00000001




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00000001





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CONT'D

All skin panels (plates) of the forward fuselage are listed within the

nomenclature located at the front page of SRM 53-41-11. Using the

information collected before (affected panel : lower side panel - left,

between Fr 47 & 64 and Stgr 18 & 32), the nomenclature provides

the figure number we have to refer to: "Skin plates - LWR parts LH

Fr 47 to Fr 64: Refer To Figure 6".


00000001




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000000001





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CONT'D

The figure 6, identifies two different panels configurations (view A

and B):

- the view A applies to the basic version of the skin panel,

- the view B applies to the evolution of the skin panel.

So, there are two different panel configurations illustrated, showing

the basic version of the panel and an other possible version effective

after the embodiment of production modification(s). The modification

numbers are indicated at the bottom of the page. The next step of the

investigation is to define which of these panels is installed on the

concerned A/C.


000000001




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000000001





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CONT'D

To identify the actual panel, the modification numbers indicated at

the bottom of the page have to be compared with the service

bulletin/modification list (located at the beginning of chapter

53-40-00). In this list, the effectivity in terms of Manufacturer Serial

Number (MSN) has to be checked.


000000001




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000000001





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CONT'D

The view B of the panel on figure 1, sheet 5 is valid after Modifications

(MODs) 21468K1489A, 22083K2232B, 24958K4082D or

31012K7082. Checking the Modification/Service Bulletin (SB) List

at the beginning of chapter 53-40-00, it appears that our MSN (2042)

is in this list of MSN proposed for the MOD number 31012K7082.

So, the panel installed on the A/C is a modified version, then refer to

view B figure 6, sheet 2.


C000000001




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C000000001





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CONT'D

The damage is located between Fr 62 and 63, and is located between

the sixth and the eighth stringer from Stgr 18 (longitudinal skin joint

reference).

This information can be reported onto the illustration and gives:

- the material thickness of the area (code C, giving 1.6 mm (0.063 in),

- the stringer location: the damage is located between Stgr 23LH and

25LH.


C000000001




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C000000001





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and the nominal skin thickness in the scratched area. The damage

assessment using the allowable damage page block is the next step.


C000000001




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C000000001





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DAMAGE ASSESSMENT

To start the damage assessments refer to the page block 101 of the relevant

chapter/section (53-41-11). And start to read carefully the procedure. A

special attention shall be paid to the notes and cautions.


C000000001




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DAMAGE ASSESSMENT


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DAMAGE ASSESSMENT (continued)

DAMAGE CRITERIA

Read carefully all the cautions, they could give you information on

the assessment. A second caution note indicates that in some cases,

an inspection may be required to check for crack, even if the damage

is determined as being allowable.


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In the allowable damage description/criteria table (101), the paragraph

4A has to be acknowledged for reworks.


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Check the applicability of the allowable damage for reworks.


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WEIGHT VARIANT

A caution note indicates that the allowable damage effectivity per

A/C weight variant may have to be verified. The weight variant is a

criterion, which is defined for each model of A/C and depending on

its Maximum Take Off Weight (MTOW), Maximum Landing Weight

(MLW), and Maximum Zero Fuel Weight (MZFW). The allowable

damage limits are defined per weight variant and for a same model.

The weight variant can change depending on modification or SB

embodiment status. The actual weight variant of the affected A/C has

to be known before starting the assessment. Because of the

modifications, which could be embodied on the A/C, only the

airline-engineering department shall give you this information. The

actual weight variant shall be compared with the data given in a table

in introduction pages.


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WEIGHT VARIANT (CONT'D)


pages, the allowable damage applicability has to be checked, using

the weight variant exclusion table (table 4) given at the introduction

of the SRM. The information coming from the airline-engineering

department shows that the MSN 2042 is at weight variant 010.

Checking the table 4, the weight variant 010 is included in and thus,

the following allowable damage information can be used.


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DAMAGE ASSESSMENT - WEIGHT VARIANT (CONT'D)


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Two diagrams are given, one for riveted areas and one for unriveted

areas.


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The parameters of the rework have to be determined to be sure that

we are dealing with the correct diagram.


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To determine which diagram to use, we have to check if the damage

is located in a riveted area. A riveted area extends from less than 15

mm (0.590 in) all around a rivet. In the applicable damage, the riveted

area and the unriveted area have to be acknowledged. The maximum

depths of the scratch in riveted and unriveted areas have to be

acknowledged too.


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To be allowable, the rework width has to be equal or longer than forty

times the depth.


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To complete the diagram, the maximum depth of the rework has to

be expressed as a percentage of the damaged skin thickness.

Those two values (found before) are plotted onto the diagram, which

defines a point. The area where this point is located defines the

subsequent actions to be performed.


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Check for the first criterion to be fulfilled: the width of the damage

must be at least 40 x T. The depth of the depression in the riveted area

is 0.2 mm; 40*0.2 = 8 mm. The width of the depression is 18.5 mm,

which is higher than 8 mm. So, the first criterion is met.


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To complete the diagram, the maximum depth of the rework in the

riveted area has to be expressed as a percentage of the damaged skin

thickness.


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ALLOWABLE REWORK DIAGRAM

The depth of the rework as percentage of the skin thickness in

unriveted area, and the length of the rework, are the key to get into to

diagram. You must refer to the data collected before. These two values

are plotted onto the diagram, which defines a point. The area where

this point is located defines the subsequent actions to be performed.

For the concerned rework, read the note: "Check damage for cracks.

Remove damage up to depression depth "T" (section view). Renew

surface protection and repair after 50 flights at the latest". Providedthat no crack is detected by detailed visual inspection, the rework is

considered as an allowable damage with a time limit (temporary

allowable damage). The A/C can be released. But a repair will have

to be done before 50 flights.


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DAMAGE ASSESSMENT - ALLOWABLE REWORK DIAGRAM


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ALLOWABLE REWORK DIAGRAM (CONT'D)

This diagram enables to determine if the damage is allowable, and

the condition of allowability.


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DAMAGE ASSESSMENT - ALLOWABLE REWORK DIAGRAM (CONT'D)


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DAMAGE ASSESSMENT - CONCLUSION

In the allowable example, two assessments have been done, the more

restrictive one has to be acknowledged. So, the damage has to be checked

for cracks, damage up to depression depth has to be removed, the surface

has to be renewed and the A/C has to be repaired after 50 flights at the

latest.


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DAMAGE ASSESSMENT - CONCLUSION


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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

INTRODUCTION

The purpose of this example is to present you, the complete procedure

to be followed when a damage is discovered, from the damage mapping

draft to the final structure damage assessment. This example was chosen

as it represents one of the more usual types of damage on the A/C and

enables to make an in depth investigation with all the different stages.


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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) May 10, 2006

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INTRODUCTION


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any, are to be found within chapter 53 of the SRM. Information

concerning the damage classification and reporting are to be found within

SRM chapter 51-11-00. The applicable damage is a scratch with no visible

crack.

At this stage: take visual reference to facilitate damage location.

Such as, forward or an aft passengers door, or a cargo door, above or

below cabin floor level at stringer (Stgr) 23, near a longitudinal orcircumferential joint, etc...).

If the scratch is near a rivet row, an internal visual inspection is required

to determine whether the internal structure (frame, stringer, etc...) is also

damaged or not.


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MAPPING

DRAFT

Using the SRM 51-11-13 as a guide, the maximum information should

be taken from the A/C before starting any assessment (measurement

and location of the maximum depth, distance from dent edges to the

nearest fastener rows, existing closest skin joints or any other visible

structure that will help in the detailed location of the damage, etc...).


regarding the identification of the part, allowable damage and repair,

if any, are to be found within the chapter 53 of the SRM.

NOTE: on the affected panel, there are no stringer rivet rows, thus,

stringers, if any, should be welded onto the skin: it is not

possible to identify the stringer references.

As a consequence, it is necessary to measure the distance from a

longitudinal skin joint to the dent maximum depth, in order to get a

reference for the location of the dent. This reference will be compared

with the welded stringer references coming from the SRM, page blocks

101 and 201 (see next pages).


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MAPPING - DRAFT


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MAPPING (continued)

CIRCUMFERENTIAL SKIN JOINT IDENTIFICATION

The illustration of the chapter 53-00-00 enables the operator to

determine the circumferential joints related frame numbers. In the

given example, the damage is located in the aft center fuselage section,

with the circumferential skin joints at Fr 47/54 and Fr 64.


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MAPPING - CIRCUMFERENTIAL SKIN JOINT IDENTIFICATION


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MAPPING (continued)

CIRCUMFERENTIAL SKIN JOINT IDENTIFICATION

(CONT'D)



the damage can be determined. According to the mapping information,

the damage is located between the fourth and the fifth frame before

the circumferential skin joint located at frame (Fr) 64. Consequently

the damage is located between Fr 59 and 60.


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MAPPING - CIRCUMFERENTIAL SKIN JOINT IDENTIFICATION (CONT'D)


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MAPPING (continued)

LONGITUDINAL SKIN JOINT AND PANEL

IDENTIFICATION (CONT'D)


also need to be determined. For this purpose, the "General Panel

Identification" illustrations, proposed within chapter 53-00-00, can

be used. According to the data collected on the A/C and the location

of the damage from the existing longitudinal skin joints, the affected

panel can be determined. For this example, the damage is located on

panel 5 - lower side shell - located between Stgr 32 LH and 41 RH,and Fr 47/54 and 64. As seen before, this panel is a welded

skin/stringer panel, thus, refer to page block 101 to determine the

stringers position.


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MAPPING - LONGITUDINAL SKIN JOINT AND PANEL IDENTIFICATION (CONT'D)


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MAPPING (continued)

WELDED STRINGERS POSITION

First, refer to the page block 101 of the relevant chapter/section

(53-41-11) and start to read carefully the procedure. Refer to the

allowable damage description/criteria table to find the concerned

paragraph (4F): "Fuselage Skin Plates Fr 47 / 54 Thru Fr 64 Between

Stgr 32 LH and Stgr 41 RH (Welded Panel)".


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MAPPING - WELDED STRINGERS POSITION


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MAPPING (continued)

WELDED STRINGERS POSITION (CONT'D)

Read the notes within the relevant paragraph to find information about

the definition and determination of undisturbed skin (unwelded and

unriveted).


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MAPPING - WELDED STRINGERS POSITION (CONT'D)


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MAPPING (continued)


This figure shows how unwelded and unriveted areas, welded areas,

riveted areas and coupling areas are defined. Two methods of

measurement are given, we look at measurement from outside, a flag

refers to SRM chapter 53-41-11 page block 201 to get stringer

positions on a welded panel.


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MAPPING (continued)


This diagram provides the distances from the lap joint 41RH to all

welded stringers, at each frame location. Therefore a new mapping is

required. The distance from the lap joint (Stgr 41RH) to the dent

maximum depth (at Fr 60) becomes 825 mm.


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MAPPING (continued)


This distance (825 mm) must be compared with the distances from

the lap joint 41RH to the welded stringers, to locate the dent.


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MAPPING (continued)


We conclude that the dent is located between Stgr 40 and 41LH and

it is now possible to finalize the draft (see next page).


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The "fuselage section division" illustration of chapter 53-00-00 used

before enables the definition of the affected section: aft center fuselage,

part of rear fuselage - section 16/17 - chapter 53-40-00. The general

illustration of the chapter 53-40-00 identifies the main structural

arrangement of the rear fuselage. The skin plates are part of the main

structure covered by the section 53-41-00.


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DETAILED IDENTIFICATION (CONT'D)


panels (skin plates) are item number 5. The identification table informs

us that the full identification of the skin panels (skin plates) are covered

by SRM 53-41-11.


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DETAILED IDENTIFICATION - DETAILED IDENTIFICATION (CONT'D)


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The identification table of SRM 53-41-11 refers to the figure 3 sheet

2 for the skin panel located between Fr 58A & Fr 64, and Stgr 41RH

& Stgr 32LH.


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The damaged panel is illustrated onto two sheets. According to the

figure 3 sheet 1, the damaged panel is item number 1.


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There is no modification associated to item 1 thus: it is the basic panel.

No other item/Part Number (P/N) with associated Modification

(MOD)/Service Bulletins (SB) status is available in the nomenclature

table so, the skin panel of MSN 2218 is the Part Number (P/N)

D53479410202.


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DETERMINATION OF SKIN THICKNESS IN DENTED

AREA

The damage is located between Fr 59 & 60, and Stgr 40LH & 41LH.

This information can be reported onto the illustration and gives the

skin thickness in the dented area (code B, giving 1.6 mm (0.063 in)).


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DETAILED IDENTIFICATION - DETERMINATION OF SKIN THICKNESS IN DENTED AREA


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and the nominal skin thickness in the dented area. As the mapping is

at scale 1:1, we can measure the distance between Stgr 40LH and the

deepest point of the dent (59 mm), and the distance between Stgr 40

and the edge of the dent (49 mm). The damage assessment using the

allowable damage page block 101 is the next step.


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DAMAGE ASSESSMENT

GENERAL

To start the damage assessments refer to the page block 101 of the

relevant chapter/section (53-41-11), and start to read carefully the

procedure. A special attention shall be paid to the notes and cautions.


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DAMAGE ASSESSMENT - GENERAL


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WEIGHT VARIANT

In the relevant paragraph 4F (fuselage skin plates Fr 47/54 thru Fr 64

between Stgr 32 LH and Stgr 41 RH (welded panel)), a caution note

indicates that the allowable damage effectivity per A/C weight variant

may have to be verified.


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WEIGHT VARIANT (CONT'D)

The weight variant is a criterion which is defined for each model of

A/C and depending on its Maximum Take Off Weight (MTOW),

Maximum Landing Weight (MLW), and Maximum Zero Fuel Weight

(MZFW). The allowable damage limits are defined per weight variant

and, for a same model, the weight variant can change, depending on

modification or SB embodiment status. The current weight variant of

the affected A/C has to be known before starting the assessment. If

the A/C weight variant is not within the table, a damage report has to

be sent to Airbus.

Depending on SB/Mod since A/C delivery, only the airline-engineering

department is able to give you the current A/C weight variant. The

current weight variant shall be compared with the data given in a table

at the beginning of allowable damage related paragraph.


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DAMAGE ASSESSMENT - WEIGHT VARIANT (CONT'D)


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DAMAGE CRITERIA

A second caution note indicates that in some cases, an inspection

program has to be followed.


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Check the applicability of the allowable damage for dents.


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Refer to paragraph 4F for dents. Dents are considered as fulfilling

nearness/form criterion or out of nearness/form criterion in accordance

with their geometry and their proximity to the nearest adjacent internal

structure elements. This must be determined according to the

parameters defined in figure 114 and diagram 102.


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AREA OF RIVETED SUBSTRUCTURE

In areas of riveted substructure, to define whether the dent fulfils

nearness/form criterion, two criteria have to be checked. Refer to

figure 114 sheet 1. B must be at least 15 mm, where B is the smallest

distance measured from the dent edge to any fastener row or any

cutout. D must be maximum 10 % of A, where D is the maximum

depth of the dent and A is the smallest distance measured from D

point to the closest adjacent structure.


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DAMAGE ASSESSMENT - AREA OF RIVETED SUBSTRUCTURE


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AREA OF RIVETED SUBSTRUCTURE (CONT'D)

Check whether the first criterion is fulfilled: B distance is minimum

15 mm.

The smallest distance measured between the edge of the dent and the

surrounding fastener rows is 95 mm, which is higher than 15 mm.

The first criterion is met.


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DAMAGE ASSESSMENT - AREA OF RIVETED SUBSTRUCTURE (CONT'D)


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The second criterion consists in comparing the maximum depth of

the dent (D) with the smallest distance measured from the deepest

point of the dent to the closest adjacent structure (distance A). If no

access from inside, the measurement is taken from outside. In this

case, A is X-15 mm, where X is the distance between the deepest

point and the closest fastener row.


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Check whether the second criterion is fulfilled: D less or equal to 10

% of A. The depth of the dent should be maximum 10% of the distance

A. The smallest distance between the deepest point of the dent and

the surrounding fastener rows is 221 mm. Since measured from

outside, distance A = 221 mm - 15 mm = 206 mm. The second

criterion is met: D = 3 mm is smaller than 10 % of A.


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AREA OF UNRIVETED SUBSTRUCTURE

In areas of unriveted substructure, to define whether the dent fulfils

the nearness/form criterion, two criteria have to be checked. Refer to

figure 114 sheet 2. Dent should be out of the welded area. D is

maximum 10 % of A, where D is the maximum depth of the dent and

A is the distance measured from D point to the boundary of the welded

area. Figure 114 sheet 2 informs us to refer to figure 115 for welded

areas.


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DEFINITION OF AREAS (CONT'D)

This figure defines that a welded area is delimited by 25 mm up and

down the stringer.


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DAMAGE ASSESSMENT - DEFINITION OF AREAS (CONT'D)


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AREA OF UNRIVETED SUBSTRUCTURE (CONT'D)

The welded area of 50 mm width has been reported on the mapping;

the first criterion is fulfilled as the dent is out of the welded area.

Check whether the second criteria is fulfilled: D 10 % A. The dent

fulfils both nearness form/criterion for unriveted area and riveted area,

thus continue the damage assessment.


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pages, the allowable damage applicability have to be checked, using

the weight variant table (table 107) given at the beginning of the

paragraph. The information coming from the airline-engineering

department shows that the MSN 2218 is at weight variant 004. Since

the weight variant 004 is within table 107, continue the damage

assessment.


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Before starting comparing the dents in accordance with diagram 102

as mentioned in paragraph 2, read the paragraph 3.

We have checked that the dent is out of riveted areas and welded areas,

if we look at figure 115 we see that the dent is also out of coupling

areas.


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The key to the allowable damage diagram is the skin thickness in

dented area and the dent depth. You must refer to the data collected

before (damage mapping). The diagram is associated to requirements

already checked at an early stage.


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ALLOWABLE DENT DIAGRAM (CONT'D)

The skin thickness in the dented area is 1.6 mm (found in the

identification pages). The depth of the dent is 4 mm (measured from

the A/C-damage mapping). These two values are plotted onto the

diagram, which defines a point. The area where this point is located

defines the subsequent actions to be performed. For the concerned

dent read the note: "Check Damage For Cracks By Detailed Visual

Examination. If Clear, repair Within 3000 FC". Provided that no crack

is detected by detailed visual inspection, the dent is considered as an

allowable damage with a time limit (temporary allowable damage).The A/C can be released. But a repair will have to be done before

3000 Flight Cycles (FC). If cracked, contact Airbus or repair before

next flight.


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STRUCTURE PROTECTIONS & AWARENESS D/O (3)

SOURCES OF DAMAGE

Throughout its operational life, the aircraft structure is subjected to

different types of damage:

- fatigue damage (cracking),

- accidental damage (e.g. bird impact, ground handling,...),

- deteriorations due to environmental and operating conditions (lightning

strike, corrosion, ...).


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SOURCES OF DAMAGE


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SOURCES OF DAMAGE (continued)

DAMAGE DETECTION/PREVENTION

Concerning fatigue damage, the aircraft is designed and justified, to

be free of significant fatigue cracking during its Design Service Goal

(DSG). The scheduled structure inspection programs are prepared to

detect any fatigue cracking before it reaches a critical length.

Inspections for corrosion are also part of the scheduled maintenance

program. Nevertheless, the maximum protection schemes and attention

is paid to protect the aircraft structure against known environmental

aggressions. In addition, the basic protections should be kept in good

conditions and some basic precautions should also be considered.


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SURFACE PROTECTIONS

- the material,

- the function,

- the location.


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SURFACE PROTECTIONS


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SURFACE PROTECTIONS (continued)

PROTECTIVE TREATMENT AREAS - FUSELAGE

- difficult access, and/or high risk of accidental damage.


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SURFACE PROTECTIONS (continued)

TYPE OF PROTECTIVE TREATMENTS

- Type 2 - heavy-duty corrosion preventive compound: grease-like

coatings containing corrosion inhibitors which protect against corrosive

agents.


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3U06


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SEALANTS

SEALING IN TYPICAL FUSELAGE AREAS

In some specified areas of the aircraft, for example the lower shell, a

protective layer is put on the sealant. This layer makes sure that other

materials (for example, fuel, hydraulic oil, engine oil and waste fluids

from the toilets and galleys) do not cause a deterioration of the sealant.


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SEALANTS SEALING IN TYPICAL FUSELAGE AREAS


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SEALANTS (continued)

SEALING IN TYPICAL FUEL TANK AREAS

In the fuel tanks, the sealant is used to prevent fuel leaks and corrosion

of the fuel tank.


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3U0

SEALANTS SEALING IN TYPICAL FUEL TANK AREAS


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DRAINAGE

- special drain valves installed in those parts of the fuselage and which

are pressurized in flight.

The drain holes and drain valves are usually at the lowest part of the

fuselage. It is important that any unwanted liquids get to the drain holes

or valves. The structure of the lower fuselage is constructed so that a path

is given for these liquids. When you do a repair, make sure that you keep

this path free of unwanted materials.


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DRAINAGE


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DRAINAGE


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COMPOSITE PARTS PROTECTION

COMPOSITE DAMAGES

Composite structures can be damaged by lightning strikes or handling

operations. The environmental conditions may be the source of damage

like rain, dust. The structure can also be affected by impact of foreign

objects or birds for example. At the design stage, the structure has the

maximum protection against these different sources of damage.


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COMPOSITE PARTS PROTECTION - COMPOSITE DAMAGES


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COMPOSITE PARTS PROTECTION (continued)

LIGHTNING STRIKE PROTECTION

- Zone 3: this zone includes all of the aircraft surfaces that are not in

Zone 1 and 2. In Zone 3, there is a low probability of attachment of

a lightning strike. However, high lightning currents can go through

Zone 3 by direct conduction between two attachment points. Zone 3

currents will also go into Zones 1 and 2.


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COMPOSITE PARTS PROTECTION - LIGHTNING STRIKE PROTECTION


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RADOME

This AMM extract deals with an example of lightning strike protection

in Zone 1, the radome. The radome is a sandwich structure with quartz

fiber skins; it is protected using copper straps on the external surface,

and bonding braids connecting the aluminum alloy frame to the

fuselage structure.


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COMPOSITE PARTS PROTECTION - RADOME


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ELEVATORS AND RUDDER

This second example shows the lightning strike protection of the

elevators and rudder trailing edges and tip, which are also located in

Zone 1. The elevators and the rudder are basically carbon fiber

structures. Their trailing edges are made of an aluminum alloy profile

and their tips are also made of aluminum alloy.


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COMPOSITE PARTS PROTECTION - ELEVATORS AND RUDDER


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ELECTRICAL CONTINUITY

The Nose Landing Gear doors are located in Zone 2. Their protection

and the electrical continuity is achieved using a metallic grid installed

at the manufacturing stage on the top of the composite layers. Note

that in most cases, this grid should be restored when damaged, as per

the Structural Repair Manual (SRM) procedures.


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COMPOSITE PARTS PROTECTION - ELECTRICAL CONTINUITY


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HANDLING OF COMPOSITE STRUCTURES

To keep composite structures in good and serviceable conditions, the

operator should avoid any damage during handling and/or maintenance

operations (such as chopped tools, take care of no step areas, ...).

Chemical strippers are not authorized on composite structures (the

resin system may be deteriorated). The protection like paint schemes

and special layers (e.g. tedlar layers on inside surfaces) should be kept

in good condition. The drying of composites is also essential before

hot bonding repair operations.


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COMPOSITE PARTS PROTECTION - HANDLING OF COMPOSITE STRUCTURES


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ENVIRONMENTAL & IMPACT PROTECTION

The impact protection of the Trimmable Horizontal Stabilizer (THS)

leading edge is achieved by a metallic cover plate.


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COMPOSITE PARTS PROTECTION - ENVIRONMENTAL & IMPACT PROTECTION


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ENVIRONMENTAL & IMPACT PROTECTION (CONT'D)

- titanium fasteners.


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COMPOSITE PARTS PROTECTION - ENVIRONMENTAL & IMPACT PROTECTION (CONT'D)


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DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

SESSION OBJECTIVES

SESSION SET-UP

DAMAGE ASSESSMENT PROCEDURE


DETAILED IDENTIFICATION OF THE DAMAGED

PART

ALLOWABLE DAMAGE-GENERAL

DAMAGE CRITERIA

ALLOWABLE DENT DIAGRAM USAGE/FINAL

DECISION

CONCLUSION

DAMAGE LOCATION


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SESSION OBJECTIVES ... DAMAGE LOCATION


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MAPPING

DRAFT


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MAPPING - DRAFT


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MAPPING (continued)

FINALIZATION


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MAPPING - FINALIZATION


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SESSION OBJECTIVES

SESSION SET-UP


DAMAGE IDENTIFICATION/LOCATION


PART

ALLOWABLE DAMAGE - GENERAL

APPLICABLE ALLOWABLE DAMAGE DIAGRAM

ALLOWABLE SCRATCH DIAGRAM USAGE/FINAL

DECISION

CONCLUSION

DAMAGE LOCATION


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MAPPING

DRAFT


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MAPPING - DRAFT


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MAPPING (continued)

FINALIZATION


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MAPPING - FINALIZATION


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A318 DAMAGE ASSESSMENT EX. 3 OPERAT. SCENARIO(3)

SESSION OBJECTIVES

SESSION SET-UP



STRINGERS LOCATION


PART

ALLOWABLE DAMAGE - GENERAL

DAMAGE CRITERIA

CONCLUSION

DAMAGE LOCATION

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A318 DAMAGE ASSESSMENT EX. 3 OPERAT. SCENARIO(3) May 10, 2006

Page 702


MAINTENANCE COURSE T1 (V2500 A5/ME) A318 DAMAGE ASSESSMENT EX 3 OPERAT SCENARIO(3) M 10 2006


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MAPPING

DRAFT

MAINTENANCE COURSE T1 (V2500 A5/ME) A318 DAMAGE ASSESSMENT EX 3 OPERAT SCENARIO(3) M 10 2006


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MAPPING - DRAFT

MAINTENANCE COURSE T1 (V2500 A5/ME) A318 DAMAGE ASSESSMENT EX 3 OPERAT SCENARIO(3) May 10 2006


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MAPPING (continued)

RESULT



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MAPPING - RESULT



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51 Structure

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