51 Structure.pdf

 Single Aisle   TECHNICAL TRAINING MANUAL   T1+T2 (CFM 56) (Lvl 2&3)   STRUCTURE 

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Single Aisle TECHNICAL TRAINING MANUAL

STRUCTURE

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Structure General (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Doors D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18 Fuselage D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34 A318 Fuselage D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72 Pylons/Nacelles D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 108 Stabilizers D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134 A318 Stabilizers D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 168 Windows D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Wings D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216 Structure Damage Identification D/O (3) . . . . . . . . . . . . . . . . . . . . . 264 Window Damage Identification D/O (3) . . . . . . . . . . . . . . . . . . . . . 296 SRM D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 306 Window Damage Assessment D/O (3) . . . . . . . . . . . . . . . . . . . . . . . 340 Damage Assessment Example 1 D/O (3) . . . . . . . . . . . . . . . . . . . . . 356 Damage Assessment Example 2 D/O (3) . . . . . . . . . . . . . . . . . . . . . 420 A318 Damage Assessment Example 3 D/O (3) . . . . . . . . . . . . . . . . 484 Structure Protections & Awareness D/O (3) . . . . . . . . . . . . . . . . . . . 556 Damage Assessment Ex. 1 Operational Scenario (3) . . . . . . . . . . . . 588 Damage Assessment Ex. 2 Operational Scenario (3) . . . . . . . . . . . . 596 A318 Damage Assessment EX. 3 OPER Scenario (3) . . . . . . . . . . . 604

T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

TABLE OF CONTENTS

Oct 25, 2008 Page 1

Single Aisle TECHNICAL TRAINING MANUAL

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

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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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STRUCTURE GENERAL (3)

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AIRCRAFT MATERIALS - METALLIC MATERIALS & COMPOSITE MATERIALS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE GENERAL (3)

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STRUCTURE GENERAL (3) 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

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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 GENERAL (3)

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Single Aisle TECHNICAL TRAINING MANUAL

STRUCTURE PROTECTION - AIRCRAFT DRAINAGE & SURFACE PRETREATMENT T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE GENERAL (3)

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STRUCTURE GENERAL (3) STRUCTURE PROTECTION (continued) PAINT SYSTEM

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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 GENERAL (3)

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STRUCTURE PROTECTION - PAINT SYSTEM T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE GENERAL (3)

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STRUCTURE GENERAL (3) STRUCTURE PROTECTION (continued) NO STEP AREAS Protective mats are required on the horizontal stabilizer as it is a carbon fiber structure.

JACKING POINTS

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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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Single Aisle TECHNICAL TRAINING MANUAL

STRUCTURE PROTECTION & JACKING POINTS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE GENERAL (3)

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STRUCTURE GENERAL (3) 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

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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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STRUCTURE GENERAL (3)

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Single Aisle TECHNICAL TRAINING MANUAL

A318 STRUCTURE DIFFERENCES - GENERAL STRUCTURE CHANGES T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE GENERAL (3)

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STRUCTURE GENERAL (3) A318 STRUCTURE DIFFERENCES (continued) VERTICAL STABILIZER

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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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STRUCTURE GENERAL (3)

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A318 STRUCTURE DIFFERENCES - VERTICAL STABILIZER T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE GENERAL (3)

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STRUCTURE GENERAL (3) A318 STRUCTURE DIFFERENCES (continued) LASER BEAM WELDING

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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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STRUCTURE GENERAL (3)

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A318 STRUCTURE DIFFERENCES - LASER BEAM WELDING T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE GENERAL (3)

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STRUCTURE GENERAL (3) A318 STRUCTURE DIFFERENCES (continued) CARGO DOORS

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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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STRUCTURE GENERAL (3)

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A318 STRUCTURE DIFFERENCES - CARGO DOORS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE GENERAL (3)

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DOORS D/O (3) GENERAL

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The fuselage has: - 4 passenger/crew doors, - 2 or 4, emergency exits depending on the A/C type or option, - 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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DOORS D/O (3)

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GENERAL T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DOORS D/O (3)

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DOORS D/O (3) PASSENGER COMPARTMENT DOORS PASSENGER/CREW DOOR

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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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DOORS D/O (3)

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PASSENGER COMPARTMENT DOORS - PASSENGER/CREW DOOR T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DOORS D/O (3)

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DOORS D/O (3) PASSENGER COMPARTMENT DOORS (continued) EMERGENCY EXIT DOORS

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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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DOORS D/O (3)

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PASSENGER COMPARTMENT DOORS - EMERGENCY EXIT DOORS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DOORS D/O (3)

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DOORS D/O (3) CARGO COMPARTMENT DOORS FWD & AFT CARGO DOORS

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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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DOORS D/O (3)

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CARGO COMPARTMENT DOORS - FWD & AFT CARGO DOORS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DOORS D/O (3)

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DOORS D/O (3) CARGO COMPARTMENT DOORS (continued) BULK CARGO DOOR (A320 & A321 ONLY)

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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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DOORS D/O (3)

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CARGO COMPARTMENT DOORS - BULK CARGO DOOR (A320 & A321 ONLY) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DOORS D/O (3)

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DOORS D/O (3) ACCESS & SERVICE DOORS

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The access doors are installed in the aircraft for inspection of the structure and to give access to maintenance. Service doors are installed in the fuselage 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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DOORS D/O (3)

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DOORS D/O (3) LANDING GEAR DOORS NOSE LANDING GEAR (NLG) DOORS

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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 to allow 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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DOORS D/O (3)

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LANDING GEAR DOORS - NOSE LANDING GEAR (NLG) DOORS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DOORS D/O (3)

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DOORS D/O (3) LANDING GEAR DOORS (continued) MAIN LANDING GEAR (MLG) DOORS

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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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DOORS D/O (3)

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LANDING GEAR DOORS - MAIN LANDING GEAR (MLG) DOORS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DOORS D/O (3)

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FUSELAGE D/O (3) GENERAL FUSELAGE LAYOUT

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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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FUSELAGE D/O (3)

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GENERAL - FUSELAGE LAYOUT T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

FUSELAGE D/O (3)

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FUSELAGE D/O (3) GENERAL (continued) FUSELAGE BREAKDOWN

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Compared with the A320, the A321 forward fuselage is eight frame bays 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. To fulfil an optional request of cabin capacity of at least 160 passengers on A319, a second emergency exit is installed, both side, in section 15.

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FUSELAGE D/O (3)

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GENERAL - FUSELAGE BREAKDOWN T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

FUSELAGE D/O (3)

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FUSELAGE D/O (3) NOSE FORWARD FUSELAGE GENERAL ARRANGEMENT

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The nose forward fuselage includes section 11, between Fr 1 and Fr 12, 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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FUSELAGE D/O (3)

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NOSE FORWARD FUSELAGE - GENERAL ARRANGEMENT T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

FUSELAGE D/O (3)

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FUSELAGE D/O (3) NOSE FORWARD FUSELAGE (continued) FORWARD & AFT UPPER STRUCTURES

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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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FUSELAGE D/O (3)

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NOSE FORWARD FUSELAGE - FORWARD & AFT UPPER STRUCTURES T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

FUSELAGE D/O (3)

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FUSELAGE D/O (3) NOSE FORWARD FUSELAGE (continued) LOWER STRUCTURE

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This part of section 11/12 contains the nose landing gear bay, access and 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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FUSELAGE D/O (3)

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NOSE FORWARD FUSELAGE - LOWER STRUCTURE T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

FUSELAGE D/O (3)

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FUSELAGE D/O (3) FORWARD FUSELAGE GENERAL ARRANGEMENT

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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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FUSELAGE D/O (3) FORWARD FUSELAGE (continued) TYPICAL STRUCTURE

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This section is of conventional construction mainly composed of chemically milled skin panels, frames made from sheet metal and extruded stringer profiles. 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 means of sheet metal cleats. The structure of the cabin floor has: - cross beams, - seat tracks, - floor support struts, - floor panels.

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FUSELAGE D/O (3) FORWARD FUSELAGE (continued) LONGITUDINAL SKIN JOINTS

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The longitudinal joints are generally longitudinal lap joints with three rivet - row joints. As the skin is, in most areas, 1.6 thick, it is reinforced by bonded doubler straps of at least 0.6 mm to allow countersunk riveting. For chemical milled skins the maximum thickness is at least 2.2 mm in the joint areas. In addition, at each intersection of frames and lap joints, a 1 mm titanium alloy strap has been added to provide good damage tolerance capabilities.

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FUSELAGE D/O (3) FORWARD FUSELAGE (continued) CIRCUMFERENTIAL SKIN JOINTS

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At the typical joints the skin panels are connected by a circumferential strap and three rivet rows. The stringers are coupled by joint pieces. In the region of the circumferential joints all stringers are riveted to the skin. The panels of the lower skin joints of section 15 and 16 are milled such that they will overlap each other after being riveted. The joint is supported by a frame, which has a T-shaped cross section. The stringers are connected with a milled stringer joint fitting.

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FUSELAGE D/O (3) CENTER FUSELAGE GENERAL ARRANGEMENT

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The fuselage center section (section 15) extends from Fr 35 to Fr 47 for 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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FUSELAGE D/O (3) CENTER FUSELAGE (continued) KEEL BEAM

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The longitudinal structural continuity of the lower fuselage in this area 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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FUSELAGE D/O (3) CENTER FUSELAGE (continued) BELLY FAIRING

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The belly fairing includes a substructure made of aluminum alloy frames 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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FUSELAGE D/O (3) REAR FUSELAGE - A319 & A320 GENERAL ARRANGEMENT

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The rear fuselage assembly is a pressurized area, which extends from Fr 47 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 (RH side). 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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FUSELAGE D/O (3) REAR FUSELAGE - A321 GENERAL ARRANGEMENT

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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 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 is installed in a separate compartment below each door.

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FUSELAGE D/O (3) CONE/REAR FUSELAGE GENERAL ARRANGEMENT

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This section comprises the un-pressurized part of the rear fuselage extending from Fr 70 to Fr 87. It includes: - the mounting structures for the vertical and horizontal stabilizers, - the rear pressure bulkhead, - a support pad used during jacking operations, - 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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FUSELAGE D/O (3) CONE/REAR FUSELAGE (continued) REAR PRESSURE BULKHEAD

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The rear pressure bulkhead installed at Fr 70, divides the pressurized rear 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) CONE/REAR FUSELAGE (continued) VERTICAL STABILIZER ATTACHMENT FITTINGS

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The vertical stabilizer spar box attachment fittings are located at Fr 70, Fr 72 and Fr 74. They have six fail safe yokes, which transmit the vertical and longitudinal stabilizer loads into the fuselage frames via shear bolts. Transversal loads are transmitted via rods between stabilizer and attachment fittings. The upper segments of frames 70, 72 and 74 are machined from plates while the lower segments are made from sheet metal.

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FUSELAGE D/O (3) CONE/REAR FUSELAGE (continued) THS ATTACHMENT FITTINGS

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The fuselage area between Fr 73 and Fr 77 houses the horizontal stabilizer. 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 and vertical 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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FUSELAGE D/O (3) CONE/REAR FUSELAGE (continued) TAIL CONE

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The tail cone unit is located aft of Fr 77 and houses the APU. This section is connected to section 19 by means of four lugs and one spigot. The inner skins and forward wall of the APU compartment are made from titanium alloy to create a fire containment compartment.

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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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A318 FUSELAGE D/O (3) GENERAL (continued) FRAME/SKIN/STRINGER ASSEMBLY

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Standard frames have a common z-shape section made from formed sheet. 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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The nose forward fuselage has section 11, from Fr 1 to Fr 12 and section 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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A318 FUSELAGE D/O (3) NOSE FORWARD FUSELAGE (continued) UPPER STRUCTURE

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The upper structure between Fr 1 and Fr 12 has closed frames and opened 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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A318 FUSELAGE D/O (3) NOSE FORWARD FUSELAGE (continued) LOWER STRUCTURE

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This part of section 11/12 contains the nose landing gear bay, access and 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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A318 FUSELAGE D/O (3) FORWARD FUSELAGE GENERAL ARRANGEMENT

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This region of the fuselage lies between Fr 24 and 35. It contains the front 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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A318 FUSELAGE D/O (3) FORWARD FUSELAGE (continued) TYPICAL STRUCTURE

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This section is of conventional construction, having chemically milled skin 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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A318 FUSELAGE D/O (3) FORWARD FUSELAGE (continued) LONGITUDINAL SKIN JOINTS

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The longitudinal joints are generally longitudinal lap joints with three rivet - row joints. As the skin is, in most areas, 1.6 thick, it is reinforced by bonded doubler straps of at least 0.6 mm to allow countersunk riveting. For chemical milled skins the maximum thickness is at least 2.2 mm in the joint areas. In addition, at each intersection of frames and lap joints, a 1 mm titanium alloy strap has been added to provide good damage tolerance capabilities.

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A318 FUSELAGE D/O (3) FORWARD FUSELAGE (continued) CIRCUMFERENTIAL SKIN JOINTS

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At the typical joints the skin panels are connected by a circumferential strap and three rivet rows. The stringers are coupled by joint pieces. In the region of the circumferential joints all stringers are riveted to the skin. The panels of the lower skin joints of section 15 and 16 are milled such that they will overlap each other after being riveted. The joint is supported by a frame, which has a T-shaped cross section. The stringers are connected with a milled stringer joint fitting.

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A318 FUSELAGE D/O (3) CENTER FUSELAGE GENERAL ARRANGEMENT

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The fuselage center section extends from Fr 35 to Fr 47/54, and integrates 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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A318 FUSELAGE D/O (3) CENTER FUSELAGE (continued) KEEL BEAM

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In this area, the longitudinal structural continuity of the lower fuselage is 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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A318 FUSELAGE D/O (3) CENTER FUSELAGE (continued) BELLY FAIRING

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The belly fairing has a substructure made of aluminum alloy frames and 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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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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A318 FUSELAGE D/O (3) CONE/REAR FUSELAGE GENERAL ARRANGEMENT

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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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A318 FUSELAGE D/O (3) CONE/REAR FUSELAGE (continued) REAR PRESSURE BULKHEAD

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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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A318 FUSELAGE D/O (3) CONE/REAR FUSELAGE (continued) VERTICAL STABILIZER ATTACHMENT FITTINGS

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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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A318 FUSELAGE D/O (3) CONE/REAR FUSELAGE (continued) THS ATTACHMENT FITTINGS

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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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A318 FUSELAGE D/O (3) CONE/REAR FUSELAGE (continued) TAIL CONE

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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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CONE/REAR FUSELAGE - TAIL CONE T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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PYLONS/NACELLES D/O (3) GENERAL

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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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PYLONS/NACELLES D/O (3) PYLONS - GENERAL ARRANGEMENT

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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/NACELLES D/O (3) PYLONS PRIMARY STRUCTURE - PYLON BOX GENERAL ARRANGEMENT

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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/NACELLES D/O (3) PYLONS PRIMARY STRUCTURE - PYLON BOX (continued) MAIN ASSEMBLY

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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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PYLONS/NACELLES D/O (3) PYLONS PRIMARY STRUCTURE - PYLON BOX (continued) PYLON TO WING ATTACHMENT

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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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PYLONS/NACELLES D/O (3) PYLONS PRIMARY STRUCTURE - PYLON BOX (continued) PYLON TO ENGINE ATTACHMENT

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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/NACELLES D/O (3) PYLONS SECONDARY STRUCTURE GENERAL ARRANGEMENT

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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/NACELLES D/O (3) PYLONS SECONDARY STRUCTURE (continued) FORWARD FAIRING

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The forward fairing can be divided into two sections; the cantilever structure between Rib 01 and Rib 05, and the structure between Rib 05 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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PYLONS/NACELLES D/O (3) PYLONS SECONDARY STRUCTURE (continued) PYLON TO WING CENTER FILLETS

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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 Carbon Fiber Reinforced Plastic (CFRP) and Aramid Fiber Reinforced Plastic (AFRP) sandwich construction.

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PYLONS SECONDARY STRUCTURE - PYLON TO WING CENTER FILLETS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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PYLONS/NACELLES D/O (3) PYLONS SECONDARY STRUCTURE (continued) AFT FAIRING

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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 inconel. 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 from AFRP sandwich construction and the tail cone is made from AFRP or CFRP monolithic construction.

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PYLONS/NACELLES D/O (3) PYLONS SECONDARY STRUCTURE (continued) LOWER FAIRING

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A fairing located under the pylon box (lower fairing) makes sure there is a continuity of the aerodynamic profile between the pylon box and the 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 structure is mainly made of aluminum alloy, stainless steel and inconel alloys.

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PYLONS/NACELLES D/O (3) PYLON TO NACELLE JUNCTION

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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 and Rib 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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PYLONS/NACELLES D/O (3) NACELLES - GENERAL

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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 D/O (3) STABILIZERS - GENERAL ARRANGEMENT

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Stabilizers are composed of: Trimmable Horizontal Stabilizer (THS), elevators, the vertical stabilizer and rudder.

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STABILIZERS D/O (3) TRIMMABLE HORIZONTAL STABILIZER (THS) GENERAL ARRANGEMENT

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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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STABILIZERS D/O (3) TRIMMABLE HORIZONTAL STABILIZER (THS) (continued) SPAR BOXES

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The complete spar box assembly has the LH and RH boxes and the center 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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STABILIZERS D/O (3) TRIMMABLE HORIZONTAL STABILIZER (THS) (continued) MAIN SUPPORT FITTINGS

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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 ball nut and a jack crew. 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 at Rib 3. All fittings are made of CFRP.

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STABILIZERS D/O (3) TRIMMABLE HORIZONTAL STABILIZER (THS) (continued) ELEVATOR ATTACHMENT FITTINGS

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

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STABILIZERS D/O (3) TRIMMABLE HORIZONTAL STABILIZER (THS) (continued) LEADING EDGE

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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 edges stretch from Rib 1 to Rib 25 of the leading edge structure. Each section has a stainless steel protection, bonded to the leading edge. The leading edge intersection is attached to Rib 1 of the leading edge substructure and to the spar box. A rubber strip is attached to the intersection. It seals the gap between the fuselage skin and the leading edge intersection.

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STABILIZERS D/O (3) TRIMMABLE HORIZONTAL STABILIZER (THS) (continued) TIP

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The tips of the THS are the LH and RH outer fairings. The tips are made 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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STABILIZERS D/O (3) TRIMMABLE HORIZONTAL STABILIZER (THS) (continued) TRAILING EDGE

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The trailing edge shapes 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 nine intermediate ribs, and by 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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STABILIZERS D/O (3) ELEVATORS - STRUCTURE LAYOUT

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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. The outboard closing rib and the tip are made from aluminum alloy, like the actuator and hinge attachment fittings. Rivets attach an aluminum profile to the trailing edge to act as lightning strike protection. Six hinge fittings attach each elevator to the spar box of the THS. Two fittings attach the servo control units. You can remove the tips and the inboard end caps. Each elevator has three hoisting points and four static dischargers.

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STABILIZERS D/O (3) VERTICAL STABILIZER GENERAL ARRANGEMENT

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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 trailing edge, - the tip, - the attach fittings.

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STABILIZERS D/O (3) VERTICAL STABILIZER (continued) SPAR BOX - STRUCTURE LAYOUT

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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 co-bonded stringers, all laminated in CFRP. Three pairs of main attach fittings made of CFRP attach the spar box to the fuselage.

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STABILIZERS D/O (3) VERTICAL STABILIZER (continued) SPAR BOX - STRUCTURE LAYOUT (CONT'D)

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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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STABILIZERS D/O (3) VERTICAL STABILIZER (continued) LEADING EDGE

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The vertical stabilizer leading edge has three 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 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 can be bonded to the most forward part of the leading edge to prevent erosion damage.

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STABILIZERS D/O (3) VERTICAL STABILIZER (continued) TIP - STRUCTURE LAYOUT

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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 lightning strike protection strap is bonded along the top of the tip.

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STABILIZERS D/O (3) VERTICAL STABILIZER (continued) TRAILING EDGE

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The trailing edge components are attached to the rear spar of the vertical stabilizer. The supporting structure is made from aluminium alloy fittings, hinge arms 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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STABILIZERS D/O (3) RUDDER GENERAL ARRANGEMENT

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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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STABILIZERS D/O (3) RUDDER (continued) STRUCTURE LAYOUT

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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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Stabilizers are composed of the Trimmable Horizontal Stabilizer (THS), the elevators, the vertical stabilizer and the rudder.

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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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The complete spar box assembly has the LH and RH spar boxes joined together 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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The front spar joint at Rib 1 made of CFRP supports the trim actuator hinge 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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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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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. All components are laminated in CFRP. 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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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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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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A318 STABILIZERS D/O (3) ELEVATORS STRUCTURE LAYOUT

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The structure of each elevator has a front spar, a top and a bottom skin panel and four ribs. All components are laminated in CFRP. 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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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 trailing edge, - the tip, - the attach fittings.

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A318 STABILIZERS D/O (3) VERTICAL STABILIZER (continued) STRUCTURE LAYOUT

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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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A318 STABILIZERS D/O (3) VERTICAL STABILIZER (continued) SPAR BOX - STRUCTURE LAYOUT (CONT'D)

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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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A318 STABILIZERS D/O (3) VERTICAL STABILIZER (continued) LEADING EDGE

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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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A318 STABILIZERS D/O (3) VERTICAL STABILIZER (continued) TIP - STRUCTURE LAYOUT

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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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A318 STABILIZERS D/O (3) VERTICAL STABILIZER (continued) TRAILING EDGE

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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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A318 STABILIZERS D/O (3) RUDDER GENERAL ARRANGEMENT

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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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A318 STABILIZERS D/O (3) RUDDER (continued) STRUCTURE LAYOUT

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The rudder main structure is the primary structural component of the rudder. It has an assembly of two side skin panels, a front spar, a bottom closing 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

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The windows are installed in: - the cockpit, - the cabin, - the doors.

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WINDOWS D/O (3) COCKPIT WINDOWS GENERAL ARRANGEMENT

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There are two types of windows: - 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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WINDOWS D/O (3) COCKPIT WINDOWS (continued) WINSHIELDS STRUCTURE

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The windshield panels are made of several layers of different materials depending on the windows supplier (LUCAS-ACT, PPG, SPS), and are interchangeable. They are held by three retainers bolted onto the outer surface of the frame. They are installed with an anti-icing and defogging system.

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WINDOWS D/O (3) COCKPIT WINDOWS (continued) FIXED SIDE WINDOWS

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The fixed side windows have two layers of different materials depending on the windows supplier (LUCAS-ACT, PPG, SPS), and are interchangeable between the different suppliers. They are held by retainers bolted onto the inner surface of the frame. They are installed with an integral anti - icing and defogging system.

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WINDOWS D/O (3) COCKPIT WINDOWS (continued) SLIDING WINDOWS

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The sliding windows have several layers of different materials depending on the windows supplier (LUCAS-ACT, PPG, SPS), and are interchangeable between the different suppliers. Each window has an anti-icing and defogging system. The sliding windows are installed on a mobile frame, which is controlled from inside the cockpit, and the crew can use them as emergency exits.

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WINDOWS D/O (3) CABIN WINDOWS GENERAL ARRANGEMENT

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The windows are installed in frames and create a smooth surface with the fuselage skin. The cabin windows are installed and removed from the inside of the aircraft. The window sets consists of a seal, inner and outer panes made of stretched acrylic resin held together by the seal. A retainer ring and eye bolts keeps the window set in position. A vent hole in the inner pane is used to only pressurize the outer pane in normal operation. The inner pane will maintain the cabin pressure in case of outer pain failure.

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WINDOWS D/O (3) DOOR WINDOWS STRUCTURE LAYOUT

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The passenger/crew doors and emergency exit doors have a circular window. They are used for inspection and observation in order to check from outside if the cabin is pressurized or to verify if the outside is clear for door opening/slide deployment. The windows have a seal, inner and outer panes made of stretched acrylic, held in position by a retainer ring. A vent hole in the inner pane allows to pressurize only the outer pane in normal operation. The inner pain is able to maintain the cabin pressure in case of outer pane failure.

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WINGS D/O (3) GENERAL

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The aircraft wing is in the continuity of the structure going through the fuselage which is divided into three parts: - the center wing box, - the left outer wing and, - the right outer wing.

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WINGS D/O (3) CENTER WING BOX GENERAL ARRANGEMENT

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The center wing is installed in the center fuselage between the main 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, - top and bottom skin panels, - 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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WINGS D/O (3) CENTER WING BOX (continued) WING ROOT JOINT

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An upper cruciform fitting and a lower triform fitting ensure the junction between the center wing box and the outer wing box. The upper cruciform fitting makes the junction between the center wing 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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WINGS D/O (3) OUTER WING GENERAL ARRANGEMENT

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Each outer wing has: - a main structure (outer wing box), - 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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WINGS D/O (3) OUTER WING BOX GENERAL ARRANGEMENT

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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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WINGS D/O (3) OUTER WING BOX (continued) SKIN PANELS

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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 stiffened by stringers machined in aluminum alloy extrusions. The joints between panels are aluminum alloy butt straps.

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WINGS D/O (3) OUTER WING BOX (continued) RIBS & SPARS

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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 front and 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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WINGS D/O (3) OUTER WING BOX (continued) ACCESS HOLES/COVERS

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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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WINGS D/O (3) FIXED LEADING EDGE GENERAL ARRANGEMENT

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The fixed Leading Edge (LE) assembly is located forward of the front spar of the wing-box.

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WINGS D/O (3) FIXED LEADING EDGE (continued) STRUCTURE LAYOUT (1/2)

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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 wing box 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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WINGS D/O (3) FIXED LEADING EDGE (continued) STRUCTURE LAYOUT (2/2)

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

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WINGS D/O (3) SLATS GENERAL ARRANGEMENT

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

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WINGS D/O (3) SLATS (continued) STRUCTURE LAYOUT (1/2)

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Each slat has: - a front spar or the stringers (girders), - a rear spar, - a girder - ribs, - top and bottom skin panels, - 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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WINGS D/O (3) SLATS (continued) STRUCTURE LAYOUT (2/2)

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When the slats are in retracted position, seals prevent airflow between the slat and the wing. Slats 3 to 5 are de-iced; the hot air comes from the bleed air system and 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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WINGS D/O (3) FIXED TRAILING EDGE STRUCTURE LAYOUT

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The fixed trailing edge is located aft of the wing rear spar. Its structure has: - 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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WINGS D/O (3) FIXED TRAILING EDGE (continued) STRUCTURE LAYOUT (CONT'D)

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This page deals with the fixed trailing edge inner structure.

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WINGS D/O (3) TRAILING EDGE DEVICES GENERAL ARRANGEMENT

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The trailing edge devices are: - two flaps, - one aileron, - five spoilers.

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WINGS D/O (3) TRAILING EDGE DEVICES (continued) FLAPS GENERAL ARRANGEMENT

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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. In case of a drive station failure, this device carries the loads, which result in such failure.

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WINGS D/O (3) TRAILING EDGE DEVICES (continued) INBOARD FLAP STRUCTURE - A318-A319-A320

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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 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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WINGS D/O (3) TRAILING EDGE DEVICES (continued) OUTBOARD FLAP STRUCTURE - A318-A319-A320

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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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WINGS D/O (3) TRAILING EDGE DEVICES (continued) A321 FLAPS STRUCTURE

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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. The outboard flap has: - a leading edge with CFRP skin, - 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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WINGS D/O (3) TRAILING EDGE DEVICES (continued) SPOILERS GENERAL ARRANGEMENT

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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 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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WINGS D/O (3) TRAILING EDGE DEVICES (continued) SPOILERS STRUCTURE LAYOUT

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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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WINGS D/O (3) TRAILING EDGE DEVICES (continued) AILERON STRUCTURE LAYOUT

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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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STRUCTURE DAMAGE IDENTIFICATION D/O (3) GENERAL

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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 into four main groups: - mechanical action, - chemical or electrochemical reaction, - thermal action or cycling, - inherent metallurgical characteristics.

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE SCRATCH

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) CORROSION

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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, - filiform corrosion, - intergranular corrosion, - galvanic corrosion. - stress corrosion, - biological corrosion, - fretting corrosion, - exfoliation corrosion.

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TYPES OF DAMAGE ON STRUCTURE - CORROSION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) CORROSION (CONT'D)

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This page deals with: - pitting corrosion, - filiform corrosion, - intergranular corrosion, - galvanic corrosion.

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TYPES OF DAMAGE ON STRUCTURE - CORROSION (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) CORROSION (CONT'D)

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This page deals with: - stress corrosion, - biological corrosion, - fretting corrosion, - exfoliation corrosion.

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TYPES OF DAMAGE ON STRUCTURE - CORROSION (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) GOUGE

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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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TYPES OF DAMAGE ON STRUCTURE - GOUGE T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) CRACK

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

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TYPES OF DAMAGE ON STRUCTURE - CRACK T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) DENT

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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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TYPES OF DAMAGE ON STRUCTURE - DENT T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) NICK

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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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TYPES OF DAMAGE ON STRUCTURE - NICK T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) DISTORTION

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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 includes bending, buckling, deformation, imbalance, misalignment, pinching, and twisting.

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TYPES OF DAMAGE ON STRUCTURE - DISTORTION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) ABRASION

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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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TYPES OF DAMAGE ON STRUCTURE - ABRASION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) DEBONDING

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

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TYPES OF DAMAGE ON STRUCTURE - DEBONDING T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) DELAMINATION

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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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TYPES OF DAMAGE ON STRUCTURE - DELAMINATION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) FRETTING

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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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TYPES OF DAMAGE ON STRUCTURE - FRETTING T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) CREASE

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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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TYPES OF DAMAGE ON STRUCTURE - CREASE T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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STRUCTURE DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON STRUCTURE (continued) MARK

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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WINDOW DAMAGE IDENTIFICATION D/O (3) GENERAL

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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 56-21-13 for the cabin windows, - AMM 56-31-00 for the passenger /crew door windows.

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WINDOW DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON COCKPIT WINDOWS RELEVANT ATA CHAPTERS

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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 - RELEVANT ATA CHAPTERS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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WINDOW DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON COCKPIT WINDOWS (continued)

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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 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 COCKPIT WINDOWS - DAMAGE IDENTIFICATION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

WINDOW DAMAGE IDENTIFICATION D/O (3)

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WINDOW DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON CABIN AND PASSENGER/CREW DOOR WINDOWS RELEVANT ATA CHAPTERS

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The cabin and passenger/crew door windows have an inner and an outer pane. The types of damage on cabin windows are mentioned in AMM 56-21-13 page block 6xx and on passenger/crew door windows 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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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WINDOW DAMAGE IDENTIFICATION D/O (3) TYPES OF DAMAGE ON CABIN AND PASSENGER/CREW DOOR WINDOWS (continued) DAMAGE IDENTIFICATION

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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, - delamination: slate like separation of the material, - pitting: impact by hard particles against the surface.

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TYPES OF DAMAGE ON CABIN AND PASSENGER/CREW DOOR WINDOWS - DAMAGE IDENTIFICATION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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SRM D/O (3) GENERAL

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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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SRM D/O (3) MANUAL BREAKDOWN

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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SRM D/O (3) FRONT PAGES

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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SRM D/O (3) INTRODUCTION CHAPTER

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The introduction chapter contains all necessary information and explanations to enable a correct use of the manual.

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INTRODUCTION CHAPTER T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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SRM D/O (3) STRUCTURAL REPAIR INSPECTIONS (SRI) CHAPTER

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For permanent repairs with inspection program, inspections are quoted along with the repair. Due to the amount of common inspection methods, 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 Inspections (SRI) gives all necessary inspection instructions on structural damage, threshold and intervals.

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STRUCTURAL REPAIR INSPECTIONS (SRI) CHAPTER T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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SRM D/O (3) NUMBERING SYSTEM AND PAGE BLOCK ALLOCATION

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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 third element identifies the sub-section (subject) within the section and is assigned by Airbus S.A.S. The 3 first digits are assigned by the ATA spec. 100. The following 3 digits are assigned by Airbus, 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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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SRM D/O (3) CHAPTER 51 (STANDARD PRACTICES AND STRUCTURES) Information about general purposes or applicable to more than one chapter, is included in chapter 51.

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NOTE: Note: the entry in the SRM is always the specific chapter 52 to 57, depending on the damage location.

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CHAPTER 51 (STANDARD PRACTICES AND STRUCTURES) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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SRM D/O (3) CHAPTER 52 TO 57 CONTENTS LAYOUT

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Chapters 52 to 57 are specifically attributed as per ATA 2200 specification (ex ATA100). Chapters 52 to 57 all 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, the Modification/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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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SRM D/O (3) CHAPTER 52 TO 57 CONTENTS (continued) modification:service bulletin list

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Located at: - section level for chapter 52, - section level for chapter 53, - chapter level for chapter 54, - section level for chapter 55, - section level for chapter 57, this list is provided to enable the user to determine the effectivity of the modification/SB. 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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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SRM D/O (3) CHAPTER 52 TO 57 CONTENTS (continued) identification pages

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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. 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 status information.

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SRM D/O (3) CHAPTER 52 TO 57 CONTENTS (continued) IDENTIFICATION TABLE DETAILED To find the relevant effectivity linked to a modification shown in the STATUS column, the user must refer to the modification/service bulletin list.

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NOTE: 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.Note: to find the relationship between the customer version number (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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SRM D/O (3) CHAPTER 52 TO 57 CONTENTS (continued) ALLOWABLE DAMAGE PAGE BLOCK

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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 damage may require minimal rework such as cleanup or drilling of stop holes. - general information pages, - damage criteria tables, - paragraph for each type of damage, - damage measurement procedure, - damage localization (zoning) figures, - allowable damage diagram.

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SRM D/O (3) CHAPTER 52 TO 57 CONTENTS (continued) REPAIRS PAGE BLOCK

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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). 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 list.

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SRM D/O (3) SRM GENERAL USAGE PROCEDURE

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SRM general usage procedure.

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WINDOW DAMAGE ASSESSMENT D/O (3) GENERAL

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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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WINDOW DAMAGE ASSESSMENT D/O (3) INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS

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

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WINDOW DAMAGE ASSESSMENT D/O (3) INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS (continued) INVESTIGATION OF DAMAGE ON WINDSHIELDS

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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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WINDOW DAMAGE ASSESSMENT D/O (3) INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS (continued) INVESTIGATION OF DAMAGE ON SLIDING WINDOWS

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Refer to AMM 56-12-11 pages block 6xx to get the corrective actions for the following types of defect on sliding windows: - cracks, - scratches, - chips, - delaminating, - bubbles, - discoloration or burning, - interlayer micro flakes, - transparency, - crazing, - burn spots.

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WINDOW DAMAGE ASSESSMENT D/O (3) INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS (continued) INVESTIGATION OF DAMAGE ON AFT FIXED WINDOWS

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Refer to AMM 56-11-12 pages block 6xx to get the corrective actions for the following types of defect on aft fixed windows: - cracks, - scratches, - chips, - delaminating, - bubbles, - discoloration or burning, - transparency, - interlayer micro flakes, - burn spots.

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WINDOW DAMAGE ASSESSMENT D/O (3) INVESTIGATION OF DAMAGE ON CABIN WINDOWS AND PASSENGER/CREW DOOR WINDOWS

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

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WINDOW DAMAGE ASSESSMENT D/O (3) INVESTIGATION OF DAMAGE ON CABIN WINDOWS AND PASSENGER/CREW DOOR WINDOWS (continued) INVESTIGATION OF DAMAGE ON CABIN WINDOWS

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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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WINDOW DAMAGE ASSESSMENT D/O (3) INVESTIGATION OF DAMAGE ON CABIN WINDOWS AND PASSENGER/CREW DOOR WINDOWS (continued) INVESTIGATION OF DAMAGE ON PASSENGER/CREW DOOR WINDOWS

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) INTRODUCTION

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) IDENTIFICATION OF THE DAMAGE

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) MAPPING

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) MAPPING (continued) CONT'D

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) MAPPING (continued) CONT'D

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The illustration of chapter 53-00-00 enables the operator to determine the circumferential joint related frame numbers.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) MAPPING (continued) CONT'D

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) MAPPING (continued) CONT'D

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) MAPPING (continued) CONT'D

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DETAILED IDENTIFICATION

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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

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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)

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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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT GENERAL

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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 EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) WEIGHT VARIANT

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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 EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA

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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 EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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Check the applicability of the allowable damage for dents.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) 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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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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 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 104.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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 dent to 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.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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The dent fulfils "nearness/form criterion", then refer to paragraph 4.B.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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Compare the dents in accordance with diagram 103.

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) ALLOWABLE DENT DIAGRAM

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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).

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DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3) DAMAGE ASSESSMENT (continued) ALLOWABLE DENT DIAGRAM (CONT'D)

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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 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).

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DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) INTRODUCTION

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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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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) IDENTIFICATION OF THE DAMAGE

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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. 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.

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IDENTIFICATION OF THE DAMAGE T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) MAPPING

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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...).

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) MAPPING (continued) CONT'D

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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.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) MAPPING (continued) CONT'D

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The illustration of chapter 53-00-00 enables the operator to determine the circumferential skin joints related frame numbers.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) MAPPING (continued) CONT'D

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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 frame (Fr) 64. Consequently the damage is located between Fr 62 and 63.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) MAPPING (continued) CONT'D

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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 8 - lower side shell - located between Stgr 18 & 32, and Fr 47 & 64.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) MAPPING (continued) CONT'D

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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.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DETAILED IDENTIFICATION

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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.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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Following SRM 53-41-00 guidelines, the figure shows that the skin 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.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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".

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DETAILED IDENTIFICATION (continued) CONT'D

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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 skin 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.

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DETAILED IDENTIFICATION (continued) MAPPING (FINALIZATION)

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The damage mapping can now be completed with the stringer numbers and the nominal skin thickness in the scratched area. The damage assessment using the allowable damage page block is the next step.

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DETAILED IDENTIFICATION - MAPPING (FINALIZATION) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT

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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 EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA

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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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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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In the allowable damage description/criteria table (101), the paragraph 4A has to be acknowledged for reworks. Repair categories are applicable depending on the location of the damage and the weight variant of the aircraft.

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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) WEIGHT VARIANT

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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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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) WEIGHT VARIANT (CONT'D)

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The information coming from the airline engineering department shows that the weight variant of MSN 2042 to be used for allowable damage reading is still the weight variant at aircraft delivery which is given in the airplane allocation list available in the introduction chapter of the SRM.

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DAMAGE ASSESSMENT - WEIGHT VARIANT (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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Check the applicability of the allowable damage for reworks.

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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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Two diagrams are given, one for riveted areas and one for unriveted areas.

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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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Both diagrams refer to Figure 101 sheet 1 that provides guidelines for rework in riveted and unriveted areas.

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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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 x 0.2 = 8 mm. The width of the depression is 18 mm, which is higher than 8 mm. So, the first criterion is met.

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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) ALLOWABLE REWORK DIAGRAM

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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". Provided that no crack is detected by an NDT inspection as per NTM 51-10-08 Pb 601, 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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT (continued) ALLOWABLE REWORK DIAGRAM (CONT'D)

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This diagram enables to determine if the damage in riveted area is allowable, and the condition of allowability.

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DAMAGE ASSESSMENT - ALLOWABLE REWORK DIAGRAM (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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DAMAGE ASSESSMENT EXAMPLE 2 D/O (3) DAMAGE ASSESSMENT - CONCLUSION

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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 as per NTM 51-10-08 Pb 601, 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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EXAMPLE 2 D/O (3)

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) INTRODUCTION

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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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INTRODUCTION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) IDENTIFICATION OF THE DAMAGE

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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 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 or circumferential 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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IDENTIFICATION OF THE DAMAGE T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) 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...). 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.

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) MAPPING (continued) CIRCUMFERENTIAL SKIN JOINT IDENTIFICATION

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) MAPPING (continued) CIRCUMFERENTIAL SKIN JOINT IDENTIFICATION (CONT'D)

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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 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) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) MAPPING (continued) LONGITUDINAL SKIN JOINT AND PANEL IDENTIFICATION (CONT'D)

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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 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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) MAPPING (continued) WELDED STRINGERS POSITION

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) MAPPING (continued) WELDED STRINGERS POSITION (CONT'D)

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) MAPPING (continued) WELDED STRINGERS POSITION (CONT'D)

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) MAPPING (continued) WELDED STRINGERS POSITION (CONT'D)

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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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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) MAPPING (continued) WELDED STRINGERS POSITION (CONT'D)

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DETAILED IDENTIFICATION (continued) DETAILED IDENTIFICATION (CONT'D)

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Following SRM 53-41-00 guidelines, the figure shows that the skin 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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DETAILED IDENTIFICATION (continued) 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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DETAILED IDENTIFICATION (continued) DETAILED IDENTIFICATION (CONT'D)

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DETAILED IDENTIFICATION (continued) DETAILED IDENTIFICATION (CONT'D)

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DETAILED IDENTIFICATION (continued) DETERMINATION OF SKIN THICKNESS IN DENTED AREA

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DETAILED IDENTIFICATION (continued) MAPPING (FINALIZATION)

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The damage mapping can now be completed with the stringer numbers 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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT GENERAL

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) WEIGHT VARIANT

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) WEIGHT VARIANT (CONT'D)

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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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A second caution note indicates that in some cases, an inspection program has to be followed.

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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Check the applicability of the allowable damage for dents.

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) AREA OF RIVETED SUBSTRUCTURE

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) AREA OF RIVETED SUBSTRUCTURE (CONT'D)

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) 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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) AREA OF RIVETED SUBSTRUCTURE (CONT'D)

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) AREA OF UNRIVETED SUBSTRUCTURE

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) DEFINITION OF AREAS (CONT'D)

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This figure defines that a welded area is delimited by 25 mm up and down the stringer.

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) AREA OF UNRIVETED SUBSTRUCTURE (CONT'D)

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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 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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) DAMAGE CRITERIA (CONT'D)

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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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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) ALLOWABLE DENT DIAGRAM

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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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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

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DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) DAMAGE ASSESSMENT (continued) ALLOWABLE DENT DIAGRAM (CONT'D)

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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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DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) SOURCES OF DAMAGE

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) SOURCES OF DAMAGE (continued) DAMAGE DETECTION/PREVENTION

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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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SOURCES OF DAMAGE - DAMAGE DETECTION/PREVENTION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) SURFACE PROTECTIONS

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- the material, - the function, - the location.

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SURFACE PROTECTIONS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) SURFACE PROTECTIONS (continued) PROTECTIVE TREATMENT AREAS - FUSELAGE

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- difficult access, and/or high risk of accidental damage.

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SURFACE PROTECTIONS - PROTECTIVE TREATMENT AREAS - FUSELAGE T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) SURFACE PROTECTIONS (continued) TYPE OF PROTECTIVE TREATMENTS

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- Type 2 - heavy-duty corrosion preventive compound: grease-like coatings containing corrosion inhibitors which protect against corrosive agents.

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SURFACE PROTECTIONS - TYPE OF PROTECTIVE TREATMENTS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) SEALANTS SEALING IN TYPICAL FUSELAGE AREAS

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) SEALANTS (continued) SEALING IN TYPICAL FUEL TANK AREAS

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In the fuel tanks, the sealant is used to prevent fuel leaks and corrosion of the fuel tank.

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SEALANTS - SEALING IN TYPICAL FUEL TANK AREAS T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) DRAINAGE

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- 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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) COMPOSITE PARTS PROTECTION COMPOSITE DAMAGES

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) COMPOSITE PARTS PROTECTION (continued) LIGHTNING STRIKE PROTECTION

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- 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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) COMPOSITE PARTS PROTECTION (continued) RADOME

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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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STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) COMPOSITE PARTS PROTECTION (continued) ELEVATORS AND RUDDER

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) COMPOSITE PARTS PROTECTION (continued) ELECTRICAL CONTINUITY

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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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STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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COMPOSITE PARTS PROTECTION - ELECTRICAL CONTINUITY T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) COMPOSITE PARTS PROTECTION (continued) HANDLING OF COMPOSITE STRUCTURES

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) COMPOSITE PARTS PROTECTION (continued) ENVIRONMENTAL & IMPACT PROTECTION

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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 T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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STRUCTURE PROTECTIONS & AWARENESS D/O (3) COMPOSITE PARTS PROTECTION (continued) ENVIRONMENTAL & IMPACT PROTECTION (CONT'D)

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- titanium fasteners.

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STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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COMPOSITE PARTS PROTECTION - ENVIRONMENTAL & IMPACT PROTECTION (CONT'D) T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

STRUCTURE PROTECTIONS & AWARENESS D/O (3)

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DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3) SESSION OBJECTIVES SESSION SET-UP DAMAGE ASSESSMENT PROCEDURE IDENTIFICATION OF THE DAMAGE DETAILED IDENTIFICATION OF THE DAMAGED PART ALLOWABLE DAMAGE-GENERAL DAMAGE CRITERIA ALLOWABLE DENT DIAGRAM USAGE/FINAL DECISION CONCLUSION

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DAMAGE LOCATION

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SESSION OBJECTIVES ... DAMAGE LOCATION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

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DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3) MAPPING

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DRAFT

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DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

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Single Aisle TECHNICAL TRAINING MANUAL

MAPPING - DRAFT T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

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DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3) MAPPING (continued)

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FINALIZATION

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DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

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Single Aisle TECHNICAL TRAINING MANUAL

MAPPING - FINALIZATION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

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DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

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ALLOWABLE DENT DIAGRAM

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DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

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Single Aisle TECHNICAL TRAINING MANUAL

ALLOWABLE DENT DIAGRAM T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

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DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3) SESSION OBJECTIVES SESSION SET-UP DAMAGE ASSESSMENT PROCEDURE DAMAGE IDENTIFICATION/LOCATION DETAILED IDENTIFICATION OF THE DAMAGED PART ALLOWABLE DAMAGE - GENERAL APPLICABLE ALLOWABLE DAMAGE DIAGRAM ALLOWABLE SCRATCH DIAGRAM USAGE/FINAL DECISION CONCLUSION

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DAMAGE LOCATION

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DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3)

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Single Aisle TECHNICAL TRAINING MANUAL

SESSION OBJECTIVES ... DAMAGE LOCATION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3)

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DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3) MAPPING

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DRAFT

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DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3)

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Single Aisle TECHNICAL TRAINING MANUAL

MAPPING - DRAFT T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3)

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DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3) MAPPING (continued)

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FINALIZATION

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DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3)

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Single Aisle TECHNICAL TRAINING MANUAL

MAPPING - FINALIZATION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3)

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DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3)

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ALLOWABLE DENT DIAGRAM

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DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3)

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Single Aisle TECHNICAL TRAINING MANUAL

ALLOWABLE DENT DIAGRAM T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

DAMAGE ASSESSMENT EX. 2 OPERATIONAL SCENARIO (3)

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A318 DAMAGE ASSESSMENT EX. 3 OPER SCENARIO (3) SESSION OBJECTIVES SESSION SET-UP DAMAGE ASSESSMENT PROCEDURE DAMAGE IDENTIFICATION STRINGERS LOCATION DETAILED IDENTIFICATION OF THE DAMAGED PART ALLOWABLE DAMAGE - GENERAL DAMAGE CRITERIA CONCLUSION

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DAMAGE LOCATION

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A318 DAMAGE ASSESSMENT EX. 3 OPER SCENARIO (3)

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SESSION OBJECTIVES ... DAMAGE LOCATION T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EX. 3 OPER SCENARIO (3)

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A318 DAMAGE ASSESSMENT EX. 3 OPER SCENARIO (3) MAPPING

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DRAFT

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A318 DAMAGE ASSESSMENT EX. 3 OPER SCENARIO (3)

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Single Aisle TECHNICAL TRAINING MANUAL

MAPPING - DRAFT T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EX. 3 OPER SCENARIO (3)

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A318 DAMAGE ASSESSMENT EX. 3 OPER SCENARIO (3) MAPPING (continued)

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RESULT

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MAPPING - RESULT T1+T2 (CFM 56) (Lvl 2&3)  51 - STRUCTURE

A318 DAMAGE ASSESSMENT EX. 3 OPER SCENARIO (3)

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AIRBUS S.A.S. 31707 BLAGNAC cedex, FRANCE STM REFERENCE U7Z08431 OCTOBER 2008 PRINTED IN FRANCE AIRBUS S.A.S. 2008 ALL RIGHTS RESERVED AN EADS COMPANY