ata-51 airbus

B r i t i s h A i r w a y s E n g i n e e r i n g T r a i n i n g

Your Course NotesThese notes have been prepared by BritishAirways Engineering Training to provide asource of reference during your period oftraining.

The information presented is as correct aspossible at the time of printing and is notsubject to amendment action.

They will be useful to you during yourtraining, but I must emphasise that theappropriate Approved Technical Publicationsmust always be used when you are actuallyworking on the aircraft.

I trust your stay with us will be informativeand enjoyable.

JOHN QUINLISKTraining and Quality Delivery Manager

STRUCTUREStructure General (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2Doors D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18Fuselage D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34A318 Fuselage D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68Pylons/Nacelles D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 100Stabilizers D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 126A318 Stabilizers D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160Windows D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 194Wings D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208Structure Damage Identification D/O (3) . . . . . . . . . . . . . . . . . . . . . 256Window Damage Identification D/O (3) . . . . . . . . . . . . . . . . . . . . . 288SRM D/O (3) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 298Window Damage Assessment D/O (3) . . . . . . . . . . . . . . . . . . . . . . . 438Damage Assessment Example 1 D/O (3) . . . . . . . . . . . . . . . . . . . . . 454Damage Assessment Example 2 D/O (3) . . . . . . . . . . . . . . . . . . . . . 518A318 Damage Assessment Example 3 D/O (3) . . . . . . . . . . . . . . . . 582Structure Protections & Awareness D/O (3) . . . . . . . . . . . . . . . . . . . 654Damage Assessment Ex. 1 Operational Scenario (3) . . . . . . . . . . . . 686Damage Assessment Ex. 2 Operational Scenario (3) . . . . . . . . . . . . 694A318 Damage Assessment Ex. 3 Operat. Scenario(3) . . . . . . . . . . . 702

MAINTENANCE COURSE - T1 (CFM56-5B/ME) 51 - STRUCTURE

TABLE OF CONTENTS May 11, 2006Page 1

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STRUCTURE GENERAL (3)

AIRCRAFT MATERIALS

METALLIC MATERIALSThe basic A/C structure is made of aluminum alloys with stainlesssteel and titanium alloys in specific areas.

COMPOSITE MATERIALSComposite 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 primarystructures, whilst Aramid Fiber Reinforced Plastic (AFRP) and GlassFiber Reinforced Plastic (GFRP) are only used for secondarystructures.


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AIRCRAFT MATERIALS - METALLIC MATERIALS & COMPOSITE MATERIALS




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STRUCTURE PROTECTION

AIRCRAFT DRAINAGEWings and fuselage have different types of drains. Holes and gaps aremeant to be used for a natural drainage of the fluid collection points.Drain holes are drilled before application of pretreatments. Remotedrains are used when natural drainage is not possible.

SURFACE PRETREATMENTThe protection of the structure against corrosion is achieved by meansof appropriate surface pretreatment of the metallic parts.Aluminum alloys: the primary protection is generally a pure aluminumcladding. The main pretreatment used is the unsealed chromic acidanodizing.Titanium alloys: surface interfaying with aluminum alloy parts arezinc sprayed. The other titanium alloy surfaces are left bare. Titaniumfasteners are either sulphuric acid anodized or aluminum coated.Composite materials are left bare.




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STRUCTURE PROTECTION (continued)

PAINT SYSTEMBefore the final paint system, all aluminum parts are primed. Thepaint system used includes polyurethane primers and paint on theexternal surfaces, and epoxy primers and polyurethane paint on theinternal surfaces. Anti-slip paint is the overwing escape zones.




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STRUCTURE PROTECTION - PAINT SYSTEM




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STRUCTURE PROTECTION (continued)

NO STEP AREASProtective mats are required on the horizontal stabilizer as it is a carbonfiber structure.

JACKING POINTS

Three jacking points are provided, one below each wing outboard of thepylon and one in front of the NLG bay.




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A318 STRUCTURE DIFFERENCES

The Main structure differences between the A319/A320/A321 and theA318 are due to the reduced length of the fuselage. There are severalgeneral structure changes, laser beam welded structures and the verticalstabilizer fin tip extension.

GENERAL STRUCTURE CHANGESThe main general structure changes are:- on section 17, due to reduced length of the fuselage, the longitudinalbeams, the seat rails and the Z-profiles are replaced by new ones. Thecrossbeams at FRame 52, FR53 and FR54 are removed. Newcrossbeams are installed between FR55 to FR64,- due to its location in the non-cylindrical part of the fuselage, a newcargo 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 theA319/A320/A321 one's,- the aft part of the belly fairing is modified due to an overlap withnon-cylindrical part of the fuselage. To avoid interference with cargocompartment door, the A318 belly fairing is two panels shorter thanthe A319/A320/A321 one's.




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A318 STRUCTURE DIFFERENCES - GENERAL STRUCTURE CHANGES




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A318 STRUCTURE DIFFERENCES (continued)

VERTICAL STABILIZERCompared with A319/A320/A321 A/Cs, the A318 vertical stabilizerfin tip is 750 mm (29,5 in.) longer.The new developed tip is completely made of GFRP. There is anadditional fin leading edge panel. There is a new spar and a new CFRPadaptor box, between the fin base and the fin tip.The metallic rudder tip is longer by 100 mm in vertical direction. Therudder trailing edge is increased in width by 50 mm.




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A318 STRUCTURE DIFFERENCES - VERTICAL STABILIZER




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LASER BEAM WELDINGThe technology used for the A319/A320/A321 A/Cs is rivetedskin/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 ontothem.




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CARGO DOORSThe A318 forward and aft cargo doors are smaller. The new cargodoor 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 nocontainerized cargo system option.




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DOORS D/O (3)

GENERAL

The fuselage has:- 4 passenger/crew doors,- 2 or 4, emergency exits depending on the A/C type,- 2 cargo compartment doors,- 1 bulk cargo compartment door (A320 & A321 only),- landing gear bay doors and access doors for servicing and maintenance.


DOORS D/O (3) May 10, 2006Page 18


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GENERAL




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DOORS D/O (3)

PASSENGER COMPARTMENT DOORS

PASSENGER/CREW DOORThe aircraft has four type C passenger doors, located on each side ofthe fuselage at frame (Fr) 16/20 and 66/68.Normal operation of the door is possible from the inside and theoutside of the aircraft. Arming of the emergency operation is onlypossible from inside.The doors are of fail-safe, plug-type construction. The structure is ofconventional design, composed of an outer skin, frame segments andbeams. Edgemembers built a surrouding frame on which hinge fittingsand locking mechanisms are installed. The loads resulting from cabinpressure are transferred by stop fittings located on each side of thedoor 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)

PASSENGER COMPARTMENT DOORS (continued)

EMERGENCY EXIT DOORSOn A318 and A319 aircraft are two Type III overwing emergencyexits 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 andcontain a standard size passenger cabin window.The A321 aircraft has four Type "C" emergency exits, one on eachside of the fuselage sections 14A and 16A, between Fr 35.1 and 35.3Aand between Fr 47.2A and 47.4. The structural design and operationof these plug-type exits is similar to the passenger doors.In an emergency, these exits can be opened manually; they are operatedlike 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)

CARGO COMPARTMENT DOORS

FWD & AFT CARGO DOORSTwo doors in the lower RH side of the fuselage provide access to themain cargo compartments.These doors are designed to carry loads from differential pressure andcircumferential loads of the frames from the fuselage. With thisconsideration, 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 thedoor, and the lower ends are attachment for the locking hooks. TheA318 cargo doors cutout is reduced by 534 mm (one frame pitch).




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DOORS D/O (3)

CARGO COMPARTMENT DOORS (continued)

BULK CARGO DOOR (A320 & A321 ONLY)The bulk cargo compartment, at the rear, has a conventional plug-typedoor, located between Fr 60 and 62.The door is operated, locked and unlocked manually and can be openedfrom the outside.It is opened by pushing inward and upward and is locked in the openposition onto the ceiling of the compartment. (In this compartment,nets are provided to maintain the clearance for the door opening). Theweight of the door is compensated by a torsion bar. The door isconnected to the door locking warning system.




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




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DOORS D/O (3)

ACCESS & SERVICE DOORS

The access doors are installed in the aircraft for inspection of the structureand to give access to maintenance. Service doors are installed in thefuselage to give access to the servicing of systems.All access and service doors are opened and closed manually.Access and service doors are illustrated as follows:- Avionics compartment door: there are four avionic compartment doorslike the one illustrated. This avionics compartment access door is installedin the lower shell of the fuselage between Fr 3 and Fr 5 in a pressurizedarea. The door can be opened from the inside or the outside.- APU doors: The APU access doors are installed in the fuselage tail conein Zone 310. These doors are located in the lower part of the fuselagebetween Fr 80A and Fr 84A. The doors give you access to the APU formaintenance.There are also access and service doors - not-illustrated: These doors arelocated in the fuselage and belly fairing for water, waste, external powerand maintenance.




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DOORS D/O (3)

LANDING GEAR DOORS

NOSE LANDING GEAR (NLG) DOORSLanding gear doors give protection to the landing gear when theaircraft is in flight.The nose and auxiliary landing gear doors have five parts:- two forward doors, hydraulically actuated, which can be closed withthe gear in the extended or retracted position. These doors are madefrom CFRP (Carbon Fiber Reinforced Plastic) sandwich materialswith a honeycomb core. They are hinged to the landing gear baylongitudinal edges.- two aft doors, linked to the gear by a rotating rod, which are madefrom CFRP sandwich materials with an honeycomb core. The purposeof these doors hinged to the landing gear bay rear lateral edge, is toallow the forward doors to be retracted when the gear is extended.- one small door (fixed door) attached to the landing gear leg is madefrom aluminum alloy.




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LANDING GEAR DOORS - NOSE LANDING GEAR (NLG) DOORS




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DOORS D/O (3)

LANDING GEAR DOORS (continued)

MAIN LANDING GEAR (MLG) DOORSThe main landing gear doors are made from CFRP sandwich materialswith a honeycomb core, and have three parts:- a main door, hydraulically actuated, which is hinged to the fuselagekeel beam parallel to the aircraft center line and can be closed withthe 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 theupper end of the main leg (hinged fairing door).All doors are part of the fuselage belly fairing and wing lower surfacein closed position.




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FUSELAGE D/O (3)

GENERAL

FUSELAGE LAYOUTThe 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).


FUSELAGE D/O (3) May 10, 2006Page 34


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FUSELAGE D/O (3)

GENERAL (continued)

FUSELAGE BREAKDOWNCompared with the A320, the A321 forward fuselage is eight framebays longer (additional section 14A, extending between frames (Fr)35 and 35.8).The A321 rear fuselage is five frame bays longer (additional section16A, 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 baysand four frame bays shorter.




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GENERAL - FUSELAGE BREAKDOWN




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FUSELAGE D/O (3)

NOSE FORWARD FUSELAGE

GENERAL ARRANGEMENTThe nose forward fuselage includes section 11, between Fr 1 and Fr12, and section 12, from Fr 12 to Fr 24.The pressurized zone extends from Fr 1 to Fr 24.The unpressurized zones are the radome, forward of Fr 1 and the noselanding gear bay.The structure of the nose forward fuselage has three parts:- the forward upper structure, between Fr 1 and 11, which makes theflight deck,- the aft upper structure, between Fr 12 and 24, which makes theforward part of the passenger cabin,- the lower structure between Fr 1 and 24.




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FUSELAGE D/O (3)

NOSE FORWARD FUSELAGE (continued)

FORWARD & AFT UPPER STRUCTURESThe 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 oftitanium alloy for bird impact requirements.The aft upper structure, between Fr 12 and Fr 24, is the forwardpassenger 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)


LOWER STRUCTUREThis part of section 11/12 contains the nose landing gear bay, accessand service door cutouts.The nose landing gear bay is shaped by three machined panelsreinforced by horizontal and vertical extruded sections attached to thecorresponding frames. The lower parts of Fr 9 and Fr 20 are theforward and rear limits of the gear bay.The lower fuselage comprises three skin panels. The central panel hasan opening for access between Fr 3 and 5 and the opening for the noselanding gear bay between Fr 9 and 20.The right hand side panel has two openings for access, between Fr 12and 14 and Fr 21 and 23.




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FORWARD FUSELAGE

GENERAL ARRANGEMENTThis area of the fuselage lies between Fr 24 and Fr 35.It contains the front part of the passenger cabin and beneath the cabinfloor and the forward cargo compartment. The forward cargo door ison 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 includesthe emergency exit cut-outs (one on each side of the fuselage) betweenFr 35.1 and Fr 35.2A.




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FUSELAGE D/O (3)

FORWARD FUSELAGE (continued)

TYPICAL STRUCTUREThis section is of conventional construction composed primarily ofchemically milled skin panels, frames and stringers made in sheetmetal.The standard frames have a common Z-shaped section made fromformed sheet, which provides a continuous structural member attachedto the skin and stringers by sheet metal cleats.The structure of the cabin floor has:- cross beams,- seat tracks,- floor support struts,- floor panels.




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CENTER FUSELAGE

GENERAL ARRANGEMENTThe fuselage center section (section 15) extends from Fr 35 to Fr 47for A320, from Fr 35.8 to Fr 47 for A321 and from Fr 35 to Fr 47/51for A319.The upper section includes part of the passenger compartment.The passenger floor structure is made of longitudinal beams, seat andsupport 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 ofthe lower fuselage,- the belly fairing supporting structure, panels and doors.




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CENTER FUSELAGE (continued)

KEEL BEAMThe longitudinal structural continuity of the lower fuselage in thisarea is maintained by the keel beam.This beam is an aluminum alloy box structure, including skins,stringers and ribs, and provides attachments for the main landing geardoors and door actuators.In its center area, the keel beam side walls are connected to thewing-box aft lower panel.




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BELLY FAIRINGThe belly fairing includes a substructure made of aluminum alloyframes and webs which are attached to the fuselage via fittings androds.This substructure supports the panels made of composite materials.The belly fairing also includes the landing gear doors, external accesspanels and access doors for maintenance.




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REAR FUSELAGE - A319 & A320 GENERALARRANGEMENT

The rear fuselage assembly is a pressurized area, which extends from Fr47 to Fr 70.The A319 and A320 rear fuselage is divided into two sections (the A321has 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 passengercabin and the aft passenger/crew doors located between Fr 66 and Fr 68.The lower part contains the aft cargo compartment. The aft cargocompartment door is installed between Fr 52A and Fr 56 (RH side); thebulk cargo compartment door is installed between Fr 60 and Fr 62 (RHside).The design of section 16/17 is similar to that of forward fuselage sections(typical skin, stringer and frame arrangement).Skin panels of the lower area have support attachment structures for thebelly fairing rear part.




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REAR FUSELAGE - A321 GENERAL ARRANGEMENT

The A321 rear fuselage assembly is a pressurized area, which extendsfrom Fr 47 to Fr 70.The A321 rear fuselage is divided into three sections:- section 16/17 and 18 which are similar to the A320,- section 16A,The section 16A of the A321 fuselage extends from Fr 47 to Fr 47.5.The section 16A includes the passenger cabin part in the upper section,and beneath the cabin floor, the forward part of the rear cargocompartment.The section 16A is of similar construction to section 16/17 but includesthe emergency exit cut-outs (one on each side of the fuselage) betweenFr 47.2A and Fr 47.4.The slide/slide-raft is installed in a separatecompartment below each door.




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FUSELAGE D/O (3)

CONE/REAR FUSELAGE

GENERAL ARRANGEMENTThis section comprises the un-pressurized part of the rear fuselageextending from Fr 70 to Fr 87.It includes:- the mounting structures for the vertical and horizontal stabilizers,- the rear pressure bulkhead,- a jacking point,- attachment structure for the tail cone, which houses the AuxiliaryPower 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 stringersand frames.The side skin panels include the horizontal stabilizer cut-out. Thelower panel has an access door for this section where a maintenancefloor is installed.




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CONE/REAR FUSELAGE - GENERAL ARRANGEMENT




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FUSELAGE D/O (3)

CONE/REAR FUSELAGE (continued)

REAR PRESSURE BULKHEADThe rear pressure bulkhead installed at Fr 70, divides the pressurizedrear fuselage from the cone/rear fuselage, which is not pressurized.It is made of a spherical membrane, and four aluminum alloy sheetsegments joined together on the inner surface by means of four "I"profile sections. Four additional "I" profile radial stiffeners are alsoinstalled.




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FUSELAGE D/O (3)


VERTICAL STABILIZER ATTACHMENT FITTINGSThe vertical stabilizer spar box attachment fittings are located at Fr70, Fr 72 and Fr 74.They have six fail safe yokes, which transmit the vertical stabilizerloads into the fuselage frames via shear bolts.The upper segments frames 70, 72 and 74 are machined from plateswhile the lower segments are made from sheet metal.




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FUSELAGE D/O (3)


THS ATTACHMENT FITTINGSThe fuselage area between Fr 73 and Fr 77 houses the horizontalstabilizer.There is a large cut-out between Fr 73 and Fr 77, which is surroundedby machined beams. A system of diagonal struts is installed on thehorizontal plane in the upper and lower areas of the cutout to increasethe rigidity of this open section.The machined frame 77 supports the tailplane hinge bearings and thelateral load fittings. They introduce horizontal stabilizer loads intothe fuselage structure, via the central bracing structure and the upperand lower bracing structures.Frame 77 also includes four lugs for the attachment of the tail coneunit.




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FUSELAGE D/O (3)


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




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A318 FUSELAGE D/O (3)

GENERAL

FUSELAGE LAYOUTThe fuselage is divided into five main parts: the nose forward fuselage(section 11/12), the forward fuselage (section 13/14), the centerfuselage (section 15), the rear fuselage (sections 16/17 and 18) andthe cone/rear fuselage (section 19/19.1).


A318 FUSELAGE D/O (3) May 10, 2006Page 68


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GENERAL (continued)

FRAME/SKIN/STRINGER ASSEMBLYStandard frames have a common z-shape section made from formedsheet. These frames are continuous structural members attached tothe 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 Stgr41RH.




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NOSE FORWARD FUSELAGE

GENERAL ARRANGEMENTThe nose forward fuselage has section 11, from Fr 1 to Fr 12 andsection 12, from Fr 12 to Fr 24. The pressurized area extends from Fr1 to Fr 24. The unpressurized areas are the radome, forward of Fr 1,and the nose landing gear bay.




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UPPER STRUCTUREThe upper structure between Fr 1 and Fr 12 has closed frames andopened frames at level of openings for:- the windshield and side windows,- the forward pressure bulkhead,- the flight deck floor support structure,- skin panels and windshield frames.The upper structure between Fr 12 and Fr 24 makes the forwardpassenger compartment and contains the two forward passenger/crewdoors.




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NOSE FORWARD FUSELAGE - UPPER STRUCTURE




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LOWER STRUCTUREThis part of section 11/12 contains the nose landing gear bay, accessand service door cutouts. The nose landing gear bay is made ofmachined flat panels stabilized laterally and longitudinally by struts.The struts are attached respectively to frames and flight deckcrossbeams.




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FORWARD FUSELAGE

GENERAL ARRANGEMENTThis region of the fuselage lies between Fr 24 and 35. It contains thefront part of the passenger cabin and, beneath the cabin floor, theforward cargo compartment. The forward cargo door is locatedbetween Fr 24A and 28 on the RH side of the fuselage




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FORWARD FUSELAGE - GENERAL ARRANGEMENT




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FORWARD FUSELAGE (continued)

TYPICAL STRUCTUREThis section is of conventional construction, having chemically milledskin panels, frames and stringers made from sheet metal. The standardframes have a common Z-shaped section made from formed sheet.They are continuous structural members attached to the skin andstringers by sheet metal cleats. A skin panel with laser beam weldedstringer is installed between Fr 24A and 35, and between Stgr 18LHand 32LH.




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FORWARD FUSELAGE - TYPICAL STRUCTURE




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CENTER FUSELAGE

GENERAL ARRANGEMENTThe fuselage center section extends from Fr 35 to Fr 47/54, andintegrates the center wing box. The upper section contains a part ofthe passenger compartment, with two overwing emergency exit doorcutouts. The pressure boundary is delimited by the forward bulkheadat Fr 35, the upper skin panel of the center wing box prolonged by apressure diaphragm up to frame 46 and ending by an inclined pressurebulkhead. Beneath the cabin floor are the air conditioning, hydraulicand main landing gears, in conjunction with a belly fairing.




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KEEL BEAMIn this area, the longitudinal structural continuity of the lower fuselageis maintained by a keel beam located between Fr 35.8 and 46. Thekeel beam transmits the overall fuselage vertical bending loads. Thisbeam is a box structure having attachments for the main landing geardoors and door actuators. In its center region, the keel beam side wallsare connected to the bottom skin panels of the center wing box.




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CENTER FUSELAGE - KEEL BEAM




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BELLY FAIRINGThe belly fairing has a substructure made of aluminum alloy framesand webs, attached to the fuselage via fittings and rods. Thissubstructure supports the panels, made of sandwich construction. Thebelly fairing also incorporates the landing gear doors, external accesspanels and access doors for maintenance.




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REAR FUSELAGE - GENERAL ARRANGEMENT

The rear fuselage assembly is a pressurized area, which extends from Fr47/54 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 forthe belly fairing rear part. The aft cargo door cutout is located betweenFr 57A and 60 on the RH side of the fuselage. Aft passenger door cutoutsare located between Fr 66 and 68.




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CONE/REAR FUSELAGE

GENERAL ARRANGEMENTThis section is the unpressurized part of the rear fuselage, aft of Fr70. It has the mounting structure for vertical and horizontal stabilizersand houses the Auxiliary Power Unit (APU). It is divided into twomain 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 panelhas a door, which gives access to this section where a maintenancefloor is installed.




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REAR PRESSURE BULKHEADThe Fr 70 supports the rear pressure bulkhead, designed as a pressurediaphragm. It is made of aluminum alloy. The bulkhead is attachedto the inside of the fuselage with a connecting strap, made of aluminumalloy.




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CONE/REAR FUSELAGE - REAR PRESSURE BULKHEAD




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VERTICAL STABILIZER ATTACHMENT FITTINGSThe vertical stabilizer spar box attachment fittings are three pairs offail safe yokes, made from forging aluminum alloy. They transmit thefin loads into the fuselage and are located at Fr 70, 72 and 74. At thoselocations, the upper frame segments are made of integrally machinedplates.




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THS ATTACHMENT FITTINGSTo house the Trimmable Horizontal Stabilizer (THS), there is a largecutout in the fuselage between Fr 74 and 77. Frame 77 is made ofintegrally machined plates and carries the THS bearing loads with thevertical link fittings. The side loads are carried through an eye bolt,linked to:- the side load fitting on the rear spar of the THS,- and oblique struts attached to the lower and upper areas of Fr 77.




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TAIL CONEThe tail cone unit is located aft of Fr 77 and houses the AuxiliaryPower Unit (APU). This section is connected to section 19 by meansof four lugs and one spigot.




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PYLONS/NACELLES D/O (3)

GENERAL

The function of the engine pylons installed under each wing is:- to support the engine,- to transmit the engine thrust to the aircraft,- to enable the routing and attachment of all the systems connected withthe engine (electrical wiring, hydraulic, bleed air and fuel lines).The nacelle gives the engine an aerodynamic shape and supports thethrust reverser system.Information concerning structure of the nacelle can be found within thenacelle manufacturer documentation.


PYLONS/NACELLES D/O (3) May 10, 2006Page 100


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GENERAL




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PYLONS - GENERAL ARRANGEMENT

The pylon has:- a primary structure attached to the wing and supporting the engine,- a secondary structure, essentially fairings, housing most of the systems.




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PYLONS - GENERAL ARRANGEMENT




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PYLONS PRIMARY STRUCTURE - PYLON BOX

GENERAL ARRANGEMENTThe pylon box is the primary structure. It supports the engine by twopoints and is attached to the wing at three points.




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PYLONS PRIMARY STRUCTURE - PYLON BOX (continued)

MAIN ASSEMBLYThe pylon box is composed of ribs, two upper spars and one lowerspar, and panels mainly made from steel and titanium alloys.




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PYLON TO WING ATTACHMENTThe forward pylon to wing attach fitting has a double lugged forkattachments connected to the wing fitting by means of four shackles.This fitting located at Rib 4 is made of titanium alloy and carriesvertical loads.The aft pylon to wing attach fitting has a single fail safe lug connectedto the wing fitting by means of two shackles. This fitting located atRib 10 is made of titanium alloy and carries vertical and side loads.Immediately behind the forward attach fitting a spherical bearingtransmits the thrust to a spigot bolted to the bottom wing skin panel.




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PYLON TO ENGINE ATTACHMENTAt The forward engine to pylon attach fitting there is a pyramidattached 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 mountlocated at Rib 3 for CFM 56-5 engine configuration or at Rib 4 forIAE V2500 engine configuration. This fitting reacts to vertical loads,side loads and roll movement.




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PYLONS SECONDARY STRUCTURE

GENERAL ARRANGEMENTThe secondary structure is composed of:- the forward fairing,- the pylon to wing center fillets,- the aft fairing,- the lower fairing.




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PYLONS SECONDARY STRUCTURE - GENERAL ARRANGEMENT




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PYLONS SECONDARY STRUCTURE (continued)

FORWARD FAIRINGThe forward fairing can be divided into two sections; the cantileverstructure between Rib 01 and Rib 05, and the structure between Rib05 and Rib 9.The cantilever structure gives an aerodynamic contour between theengine nose cowl and the pylon box structure. It routes all systemsand the bleed air from the engine to the fuselage.The structure between Rib 05 and Rib 9 gives an aerodynamic contourbetween the cantilever structure and the wing leading edge, and enablesthe 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 SECONDARY STRUCTURE - FORWARD FAIRING




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PYLON TO WING CENTER FILLETSThe pylon to wing center fillets give an aerodynamic contour betweenthe pylon box and the wing bottom skin panel.The pylon-to-wing center fillets are made of aluminum alloy ribs.These ribs support the panels made of hybrid Carbon Fiber ReinforcedPlastic (CFRP)/Aramid Fiber Reinforced Plastic (AFRP) sandwichconstruction.




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AFT FAIRINGThe aft fairing is a removable secondary structure composed of twoparts:- 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 10and by one point to the wing box at the false rear spar.The fixed fairing is assembled of ribs and skin panels made ofaluminum alloy, and includes a lower aft fairing made in AFRPsandwich construction.The movable fairing is hinged at Rib 14 and linked to the flap by arod attached to the fairing by a serrated plate system.The internal structure of the movable fairing is mainly made ofaluminum alloy. The side panels are made in CFRP or AFRP sandwichconstruction.




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LOWER FAIRINGA fairing located under the pylon box (lower fairing) makes sure thereis a continuity of the aerodynamic profile between the pylon box andthe engine nozzle.Its function is:- to supply thermal protection to the pylon from the engine exhaustgases,- to smooth out protrusions with minimal aerodynamic drag changes.The lower fairing is made of stainless steel alloy sheet except for thebottom removable sole which is made of inconel 625 alloy.




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PYLON TO NACELLE JUNCTION

The pylon to nacelle junction has:- Fan cowl door attachments.The hinge fittings of the fan cowl doors are located at Rib 01, Rib 03 andRib 05. They are made of titanium and installed on the forward secondarystructure.- Thrust reverser doors attachmentsThe hinge fittings of the thrust reverser doors are located at Rib 1 andRib 2. They are made of titanium and installed on the primary structure(pylon box). An other hinge (tie-bar) goes through the secondary structure.




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NACELLES - GENERAL

The nacelle cowling includes the inlet cowl, the fan cowl, the thrustreverser 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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NACELLES - GENERAL




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STABILIZERS D/O (3)

STABILIZERS - GENERAL ARRANGEMENT

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


STABILIZERS D/O (3) May 10, 2006Page 126


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STABILIZERS D/O (3)

TRIMMABLE HORIZONTAL STABILIZER (THS)

GENERAL ARRANGEMENTThe 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 stabilizerand support all the other components.




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STABILIZERS D/O (3)

TRIMMABLE HORIZONTAL STABILIZER (THS)(continued)

SPAR BOXESThe complete spar box assembly has the LH and RH boxes and thecenter joint.Each spar box includes top and bottom skin panels, a front spar, a rearspar and thirteen ribs (from Rib 2 thru Rib 14).The LH and RH spar boxes are laminated in Carbon Fiber ReinforcedPlastic (CFRP).The center joint is made from titanium and connects the LH and RHspar boxes to make one single unit.




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STABILIZERS D/O (3)


MAIN SUPPORT FITTINGSA hydromechanical actuator enables the adjustment of the angle ofincidence of the THS. The actuator is connected to a dual fitting (frontspar fitting) at the forward end of Rib 1, by means of hinge arms.The THS is attached to the cone rear fuselage structure at two pivotpoints (rear support fittings). They are installed on each side of theTHS centerline at Rib 3. All fittings are made of CFRP.




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STABILIZERS D/O (3)


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




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TRIMMABLE HORIZONTAL STABILIZER (THS) - ELEVATOR ATTACHMENT FITTINGS




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STABILIZERS D/O (3)


LEADING EDGEThe leading edge has an aerodynamic shape at the front of the THS.On each side of the THS centerline, the THS leading edge includes:- three leading edge primary ribs,- one inboard leading edge section,- one outboard leading edge section and,- one leading edge intersection.All components are laminated in CFRP.The front part of the inboard and outboard leading edge sections hasa stainless steel protection, bonded to the leading edge.The leading edge intersection is fitted to Rib 1 and to the spar box. Arubber strip is fitted to the intersection. It seals the gap between thefuselage skin and the leading edge intersection.




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STABILIZERS D/O (3)


TIPThe tips of the THS are the LH and RH outer fairings. The tips aremade of aluminum alloy and include rib and skin panels. The tips areattached to the leading edge rib 25 and to the upper and lower shellsof the spar box.




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STABILIZERS D/O (3)


TRAILING EDGEThe trailing edge shapes an aerodynamic surface between the THSspar box and the elevator.On each side of the THS centerline, the trailing edge panels aresupported by six intermediate ribs, and by the six hinge elevator armsupports.The access panels are laminated in CFRP bonded to a honeycombcore.On each side there are four panel assemblies on the top surface andfour access panels on the bottom surface. A rubber seal is installedbetween the panel assemblies and the access panels along the trailingedges to prevent dirtiness.




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STABILIZERS D/O (3)

ELEVATORS - STRUCTURE LAYOUT

The structure of each elevator includes:- a front spar,- top and bottom skin panels,- four ribs.All components are laminated in CFRP; the top and bottom panels aremade in sandwich construction.Rivets attach an aluminum profile to the trailing edge to make the trailingedge stronger.Six hinge fittings attach each elevator to the spar box of the THS. Twofittings attach the servo control units. You can remove the leading edgeaccess panels, 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 ARRANGEMENTThe vertical stabilizer is attached to the top of the rear fuselage. Itsupports the rudder, which is operated by three servo control units.The High Frequency (HF) antenna and the Very high frequencyOmnibearing 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 LAYOUTThe spar box is the primary structural component of the verticalstabilizer.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 sidepanels with stiffeners, all laminated in CFRP.Three pairs of main attach fittings made of CFRP attach the spar boxto the fuselage.




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VERTICAL STABILIZER - SPAR BOX - STRUCTURE LAYOUT




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STABILIZERS D/O (3)


SPAR BOX - STRUCTURE LAYOUT (CONT'D)The seven rudder hinge arms and the three actuator hinge fittings aremade from aluminum alloy and are attached to the spar box rear spar.




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




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STABILIZERS D/O (3)


LEADING EDGEThe vertical stabilizer leading edge has three removable sections.They are attached to the forward edge of the spar box side panels andto the leading edge ribs. The lower section gives access to the HFantenna (see ATA 53 fuselage description for the lower section).The three sections give an aerodynamic shape to the front of thevertical stabilizer.The three sections are laminated in Glass Fiber Reinforced Plastic(GFRP) bonded to a honeycomb core. A stainless steel cover is bondedto the inner surfaces of the sections and protects them against hail andbird impact damage.




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VERTICAL STABILIZER - LEADING EDGE




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STABILIZERS D/O (3)


TIP - STRUCTURE LAYOUTThe vertical stabilizer tip is laminated in GFRP bonded to ahoneycomb core. It is attached to the leading edge end rib and thestabilizer spar box. An aluminum lightning strike protection strap isbonded along the top of the tip.




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VERTICAL STABILIZER - TIP - STRUCTURE LAYOUT




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STABILIZERS D/O (3)


TRAILING EDGEThe trailing edge is attached to the rear of the vertical stabilizer.It has a basic aluminum alloy supporting structure made of sparsections and profiles, and four access panels on each side. The panelsgive access to the rudder servo control actuators and the hinge arms.The panels are laminated in CFRP and GFRP bonded to a honeycombcore.




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STABILIZERS D/O (3)

RUDDER

GENERAL ARRANGEMENTThe rudder is one of the primary flight controls of the aircraft.The components of the rudder are:- the main structure,- the leading edge,- the tip,- the hinge and actuator fittings.




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RUDDER - GENERAL ARRANGEMENT




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STABILIZERS D/O (3)

RUDDER (continued)

STRUCTURE LAYOUTThe rudder main structure is the primary structural component of therudder.It is an assembly of two CFRP sandwich panels, CFRP laminates frontspar, top and bottom closing ribs.An access panel, installed on the left hand side shell, gives access tothe No. 7 rudder hinge fittings. At the other locations, cutouts in theside shells give access to the adjacent hinge fittings.Three actuators and seven rudder hinge fittings are attached to theforward face of the rudder main structure, and rivets attach them tothe spar and to the skin panels.




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RUDDER - STRUCTURE LAYOUT




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A318 STABILIZERS D/O (3)

STABILIZERS

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


A318 STABILIZERS D/O (3) May 10, 2006Page 160


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TRIMMABLE HORIZONTAL STABILIZER (THS)

GENERAL ARRANGEMENTThe THS main structure has the LH and RH side spar boxes, theleading edge, the trailing edge, the THS tip and the attachment fittings.The spar boxes are the primary structure of the horizontal stabilizerand support all the other components.




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SPAR BOXESThe complete spar box assembly has the LH and RH spar boxes joinedtogether with a center joint to make one single unit. Each spar boxincludes top and bottom skin panels, a front spar, a rear spar andthirteen ribs (from Rib 2 to Rib 14), all parts being laminated in CarbonFiber Reinforced Plastic (CFRP).The center joint includes a web (Rib 1) made of CFRP and upper andlower fittings made of titanium.




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MAIN SUPPORT FITTINGSThe front spar joint at Rib 1 made of CFRP supports the trim actuatorhinge arms.The THS is attached to the cone/rear fuselage at Rib 3.On each spar box side, the attachment fittings include a THS rearsupport fitting of fail safe design with a lower and upper supportfittings, and a side load fitting. All fittings are made of CFRP exceptthe side load fitting made of aluminum alloy.




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TRIMMABLE HORIZONTAL STABILIZER (THS) - MAIN SUPPORT FITTINGS




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ELEVATOR ATTACHMENT FITTINGSOn 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 servocontrol actuators.




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LEADING EDGEAt 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 hasa 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 betweenthe fuselage skin and the intersection.




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TIPThe tips of the THS are the LH and RH outer fairings. The tips areattached to the leading edge Rib 25 and to the top and bottom skinpanels of the spar box. The tips are made from aluminum alloy.




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TRIMMABLE HORIZONTAL STABILIZER (THS) - TIP




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TRAILING EDGEThe trailing edge has an aerodynamic surface between the THS sparbox and the elevator.On each side of the THS centerline, the trailing edge panels aresupported 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 andfour access panels on the bottom surface. A rubber seal is installedbetween the panel assemblies and the access panels along the trailingedges to prevent dirtiness.




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ELEVATORS

STRUCTURE LAYOUTThe structure of each elevator has a front spar, a top and a bottomskin panel and four ribs.All components are laminated in CFRP.Six hinge fittings attach each elevator to the spar box of the THS andtwo fittings attach the servo control units. You can remove the leadingedge access panels, the tips and the inboard end caps. Each elevatorhas three hoisting points and four static dischargers.




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VERTICAL STABILIZER

GENERAL ARRANGEMENTThe vertical stabilizer is attached to the top of the rear fuselage. Itsupports the rudder, which is operated by three servo control units.The High Frequency (HF) antenna and the Very high frequencyOmnibearing 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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VERTICAL STABILIZER - GENERAL ARRANGEMENT




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STRUCTURE LAYOUTThe spar box is the primary structural component of the verticalstabilizer.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 panelswith integrated stiffeners, all laminated in CFRP.Three pairs of primary attach fittings made of CFRP attach the sparbox to the fuselage.




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VERTICAL STABILIZER - STRUCTURE LAYOUT




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SPAR BOX - STRUCTURE LAYOUT (CONT'D)The seven rudder hinge arms and the three actuator hinge fittings aremade from aluminum alloy and are attached to the spar box rear spar.




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




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LEADING EDGEThe vertical stabilizer leading edge has four removable sections. Theyare attached to the forward edge of the spar box side panels and to theleading edge ribs. The lower section gives access to the HighFrequency (HF) antenna.The four sections give an aerodynamic shape to the front of the verticalstabilizer.The four sections are laminated in Glass Fiber Reinforced Plastic(GFRP) bonded to a honeycomb core. A protective foil is bonded tothe inner surfaces of the sections and protects them against hail andbird impact damage.




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VERTICAL STABILIZER - LEADING EDGE




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TIP - STRUCTURE LAYOUTThe vertical stabilizer tip is laminated in GFRP bonded to ahoneycomb core. It is attached to the leading edge end rib and thestabilizer spar box. An aluminum strap is bonded along the top of thetip for lightning strike protection.




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VERTICAL STABILIZER - TIP - STRUCTURE LAYOUT




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TRAILING EDGEThe trailing edge is attached to the rear of the vertical stabilizer.It has a basic aluminum alloy supporting structure made of sparsections and profiles and four access panels 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 bondedto a honeycomb core.




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VERTICAL STABILIZER - TRAILING EDGE




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RUDDER

GENERAL ARRANGEMENTThe rudder is one of the primary flight controls of the aircraft.The main components of the rudder are:- the main structure,- the leading edge,- the tip,- the hinge fittings and the actuator fittings.




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RUDDER - GENERAL ARRANGEMENT




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RUDDER (continued)

STRUCTURE LAYOUTThe rudder main structure is the primary structural component of therudder.It has an assembly of two side skin panels, a front spar, a bottomclosing rib and a top closing rib. All components of the rudder mainstructure are laminated in CFRP and are attached to the rudder mainstructure.Seven rudder hinge fittings and three actuator fittings are installed onthe front spar of the rudder (all fittings are made from aluminum alloy).An aluminum profile is installed on the trailing edge of the rudder forlightning strike protection.




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RUDDER - STRUCTURE LAYOUT




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WINDOWS D/O (3)

GENERAL

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


WINDOWS D/O (3) May 10, 2006Page 194


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GENERAL




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COCKPIT WINDOWS

GENERAL ARRANGEMENTThere 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 andinstalled from the exterior.The sliding windows are installed as follows:- on a mobile frame with a mechanism controlled from the cockpit.




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COCKPIT WINDOWS - GENERAL ARRANGEMENT




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WINDOWS D/O (3)

COCKPIT WINDOWS (continued)

WINSHIELDS STRUCTUREThe windshield panels are made of several layers of different materialsdepending on the windows supplier (LUCAS-ACT, PPG, SPS), andare interchangeable. They are held by three retainers bolted onto theouter surface of the frame. They are installed with an anti-icing anddefogging system.




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COCKPIT WINDOWS - WINSHIELDS STRUCTURE




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WINDOWS D/O (3)


FIXED SIDE WINDOWSThe fixed side windows have of two layers of different materialsdepending on the windows supplier (LUCAS-ACT, PPG, SPS), andare interchangeable. They are held by retainers bolted onto the innersurface of the frame. They are installed with an integral anti - icingand defogging system.




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COCKPIT WINDOWS - FIXED SIDE WINDOWS




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WINDOWS D/O (3)


SLIDING WINDOWSThe sliding windows have several layers of different materialsdepending on the windows supplier (LUCAS-ACT, PPG, SPS), andare interchangeable. Each panel has an anti-icing and defoggingsystem. The sliding windows are installed on a mobile frame, whichis controlled from the cockpit, and the crew can use them as emergencyexits.




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COCKPIT WINDOWS - SLIDING WINDOWS




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WINDOWS D/O (3)

CABIN WINDOWS

GENERAL ARRANGEMENTThe windows are installed in frames and make a smooth surface withthe fuselage skin.The cabin windows are installed and removed from the inside of theaircraft.The windows have a circular seal, inner and outer panes made ofstretched acrylic resin held together by a retainer ring, and eye bolts.A vent hole in the inner pane lets the cabin pressure maintained in thewindow.




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CABIN WINDOWS - GENERAL ARRANGEMENT




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WINDOWS D/O (3)

DOOR WINDOWS

STRUCTURE LAYOUTThe passenger/crew doors and emergency exit doors have a circularwindow. They are used for inspection and observation in order tocheck if the cabin is pressurized and the escape slide is armed.The windows have a circular seal, inner and outer panes made ofstretched acrylic, held together by a retainer ring. A vent hole in theinner pane lets the cabin pressure maintained in the window.




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DOOR WINDOWS - STRUCTURE LAYOUT




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WINGS D/O (3)

GENERAL

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


WINGS D/O (3) May 10, 2006Page 208


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GENERAL




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WINGS D/O (3)

CENTER WING BOX

GENERAL ARRANGEMENTThe center wing is installed in the center fuselage between the mainframes (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 doneat the left hand and right hand sides rib 1.The access for maintenance to the center wing box is done throughtwo triangular openings in the rear spar.




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CENTER WING BOX - GENERAL ARRANGEMENT




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WINGS D/O (3)

CENTER WING BOX (continued)

WING ROOT JOINTAn upper cruciform fitting and a lower triform fitting ensure thejunction between the center wing box and the outer wing box.The upper cruciform fitting makes the junction between the centerwing box top skin panels, the outer wing box top skin panels, fuselageand Rib 1.The lower triform fitting and a safety butt-strap fitting make thejunction between the center wing box bottom skin panels, the outerwing box bottom skin panels and Rib 1.




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CENTER WING BOX - WING ROOT JOINT




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WINGS D/O (3)

OUTER WING

GENERAL ARRANGEMENTEach 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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OUTER WING - GENERAL ARRANGEMENT




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WINGS D/O (3)

OUTER WING BOX

GENERAL ARRANGEMENTThe outer wing box tapers from Rib 1 (the wing root) to Rib 27 hold:- the wing spars (front and rear),- the ribs,- the top and bottom skin panels,- the top and bottom stringers,- the wing-root joint.




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OUTER WING BOX - GENERAL ARRANGEMENT




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WINGS D/O (3)

OUTER WING BOX (continued)

SKIN PANELSThe top and the bottom surfaces of the outer wing box are made ofskin panels machined from aluminum alloy.There are three panels on each surface. The skin panels are stiffenedby stringers machined in aluminum alloy extrusions.The joints between panels are aluminum alloy butt straps.




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OUTER WING BOX - SKIN PANELS




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WINGS D/O (3)


RIBS & SPARSRibs:There are 27 ribs, machined in aluminum alloy, installed in the outerwing box of each outer wing. Each rib is continuous between the frontand rear spars. The junction between the center wing box and the outerwing joint is at Rib 1. Rib 1 is the boundary of the lateral section ofthe center wing box. Ribs 22 and 27 make the other lateral boundariesof the fuel and vent tanks.Spars:The wing spars are machined in aluminum alloy. They give strengthto the wing box and they extend from Rib 1 to Rib 27.




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OUTER WING BOX - RIBS & SPARS




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WINGS D/O (3)


ACCESS HOLES/COVERSThere are twenty-one access covers installed in the bottom skin panelsof the outer wing box. All panels close the openings that give accessto 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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OUTER WING BOX - ACCESS HOLES/COVERS




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FIXED LEADING EDGE

GENERAL ARRANGEMENTThe fixed Leading Edge (LE) assembly is located forward of the frontspar of the wing-box.




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FIXED LEADING EDGE - GENERAL ARRANGEMENT




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FIXED LEADING EDGE (continued)

STRUCTURE LAYOUT (1/2)The fixed leading edge assembly is made of:- the D-nose assembly, composed of aluminum alloy parts:- the support ribs and riblets (riblets are installed between the wingbox front spar and the LE spar),- the sub spar,- the LE skin.- three top surface access panels,- bottom surface access panels, which are made of Carbon FiberReinforced Plastic (CFRP) sandwich construction and are attachedwith quick release fasteners.




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WINGS D/O (3)

FIXED LEADING EDGE (continued)

STRUCTURE LAYOUT (2/2)Two pylon ribs are installed on each side of the engine pylon. Theseribs hold the pylon shroud panels.




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SLATS

GENERAL ARRANGEMENTThe wing leading edge is fitted of five slats, which make the movablepart of the wing leading edge.




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SLATS - GENERAL ARRANGEMENT




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WINGS D/O (3)

SLATS (continued)

STRUCTURE LAYOUT (1/2)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 and3).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)When the slats are in retracted position, seals prevent airflow betweenthe slat and the wing.Slats 3 to 5 are de-iced; the hot air comes from the bleed air systemand is supplied to these slats through a telescopic duct (not shown)and piccolo tubes installed in the leading edges of the slats.




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SLATS - STRUCTURE LAYOUT (2/2)




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FIXED TRAILING EDGE

STRUCTURE LAYOUTThe 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)This page deals with the fixed trailing edge inner structure.




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FIXED TRAILING EDGE - STRUCTURE LAYOUT (CONT'D)




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WINGS D/O (3)

TRAILING EDGE DEVICES

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




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TRAILING EDGE DEVICES (continued)

FLAPS GENERAL ARRANGEMENTTwo flaps are installed on the TE of the outer wing. The inboard flapis installed between Rib 1 and Rib 9 and the outboard flap is installedbetween 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, whichresult in such failure.




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INBOARD FLAP STRUCTURE - A318-A319-A320The inboard flap is supported and driven by a fuselage track andcarriage 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 madeof CFRP or aluminum alloy on the A320,- spars made of aluminum alloy on the A318 and A319, and made ofCFRP 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 ontothe outer surface of the top skin.




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OUTBOARD FLAP STRUCTURE - A318-A319-A320The outboard flap is supported and driven by two wing tracks andcarriages (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 ontothe outer surface of the top skin.




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A321 FLAPS STRUCTUREThe 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 aluminumsheet metal.The tab is operated by a linkage system.




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TRAILING EDGE DEVICES - A321 FLAPS STRUCTURE




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SPOILERS GENERAL ARRANGEMENTThere are five spoilers on the upper surface of the wing trailing edge.Spoiler 1 is connected to the false rear spar, inboard of the kinkposition.Spoilers 2 thru 5 are connected to the middle and outer sections ofthe 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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TRAILING EDGE DEVICES - SPOILERS GENERAL ARRANGEMENT




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SPOILERS STRUCTURE LAYOUTSpoilers are a wedge-shaped structure.The top and bottom skins, the sides and the trailing edge profile ofthe spoilers are made in CFRP sandwich construction.The spoiler hinges fittings and the actuator attachment fittings aremade of aluminum alloy.




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TRAILING EDGE DEVICES - SPOILERS STRUCTURE LAYOUT




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AILERON STRUCTURE LAYOUTThe aileron is located outboard of the outer flap and is connected tothe wing box rear spar between Ribs 22 and 27.It is manufactured using CFRP skin (bonded to a honeycomb core inthe center area), spar and ribs.The aileron hinge fittings and the actuator attachment fittings are madeof aluminum alloy.




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TRAILING EDGE DEVICES - AILERON STRUCTURE LAYOUT




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STRUCTURE DAMAGE IDENTIFICATION D/O (3)

GENERAL

The types of damage on metallic and composite structures are describedin SRM 51-11-00 chapter dealing with damage classification.The table provides, the term, the cause and the description for each typeof damage.Damage results from many causes and can be generally categorized intofour main groups:- mechanical action,- chemical or electrochemical reaction,- thermal action or cycling,- inherent metallurgical characteristics.


STRUCTURE DAMAGE IDENTIFICATION D/O (3) May 10, 2006Page 256


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GENERAL




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TYPES OF DAMAGE ON STRUCTURE

SCRATCHA 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 usuallythe cause.




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TYPES OF DAMAGE ON STRUCTURE - SCRATCH




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TYPES OF DAMAGE ON STRUCTURE (continued)

CORROSIONCorrosion is the destruction of metal by chemical or electrochemicaleffect. Refer to SRM 51-22-00 for general information concerningcorrosion.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




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CORROSION (CONT'D)This page deals with:- pitting corrosion,- filiform corrosion,- intergranular corrosion,- galvanic corrosion.




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TYPES OF DAMAGE ON STRUCTURE - CORROSION (CONT'D)




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CORROSION (CONT'D)This page deals with:- stress corrosion,- biological corrosion,- fretting corrosion,- exfoliation corrosion.




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TYPES OF DAMAGE ON STRUCTURE - CORROSION (CONT'D)




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GOUGEA gouge is a damaged area of any size, which results in a crosssectional area change. It is usually caused by contact with a relativelysharp object, which produces a continuous, sharp or smooth channellike groove in the material.




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TYPES OF DAMAGE ON STRUCTURE - GOUGE




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




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TYPES OF DAMAGE ON STRUCTURE - CRACK




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DENTA dent is a damaged area, which is pushed in, with respect to its usualcontour. There is no cross sectional area change in the material. Theedges of the damaged area are smooth.




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TYPES OF DAMAGE ON STRUCTURE - DENT




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NICKA nick is a small decrease of material due to, for example, a knock atthe edge of a member or a skin.




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TYPES OF DAMAGE ON STRUCTURE - NICK




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DISTORTIONA distortion is any twisting, bending or permanent strain, which resultsin misalignment or change of shape. It may be caused by an impactfrom a foreign object, but is usually the result of a vibration ormovement of adjacent attached components. This group includesbending, buckling, deformation, imbalance, misalignment, pinching,and twisting.




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ABRASIONAn abrasion is a damaged area of any size, which causes change in across sectional area because of scuffing, rubbing, scrapping or othersurface erosion. It is usually rough and irregular.




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




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DELAMINATIONA delamination is when a separation of plies occurs in multi-laminatematerial. The material being hit or when there is a resin failure forany other reason can cause a delamination.




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FRETTINGA fretting is a surface damage at the interface between elements ofthe joints resulting from very small angular or linear movements. Theresult of fretting is usually the production of fine black powderstaining.




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CREASEA crease is a damaged area, which is pushed in or folded back onitself. The edges of the damaged area are sharp or well specified linesor ridges.




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MARKA mark is a damaged area of any size where a concentration ofscratches, nicks, chips, burrs or gouges etc. is shown. You mustconsider the damage as an area and not as a serie of individualscratches, gouges, etc.




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WINDOW DAMAGE IDENTIFICATION D/O (3)

GENERAL

The windows include the cockpit windows (windshields, sliding windows,and aft fixed windows), the cabin windows and the passenger/crew doorwindows.Information dealing with different types of damage on windows, is inpage block 6xx of the relevant AMM chapters:- AMM 56-11-11 for the windshields,- AMM 56-12-11 for the sliding windows,- AMM 56-11-12 for the aft fixed windows,- AMM 56-21-13 for the cabin windows,- AMM 56-31-00 for the passenger /crew door windows.


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GENERAL




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TYPES OF DAMAGE ON COCKPIT WINDOWS

RELEVANT ATA CHAPTERSThe 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




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TYPES OF DAMAGE ON COCKPIT WINDOWS (continued)

DAMAGE IDENTIFICATIONThe 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 tohot corner effect,- bubbles: appear between the inner face of the outer ply and theinterlayer,- 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 causinga cross sectional change,- chips: flakes of glass broken from the surface and the edges of thewindow,- transparency: halos on the surface of the window can make themless transparent,- rain repellent fluid residue on windshields only,- damage on the soft liner on windshields (if supplied by SPS companyonly).




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TYPES OF DAMAGE ON CABIN AND PASSENGER/CREWDOOR WINDOWS

RELEVANT ATA CHAPTERSThe cabin and passenger/crew door windows have an inner and anouter pane. The types of damage on cabin door windows are mentionedin AMM 56-21-13 page block 6xx and on passenger/crew doorwindows in AMM 56-31-00 page block 6xx.




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TYPES OF DAMAGE ON CABIN AND PASSENGER/CREWDOOR WINDOWS (continued)

DAMAGE IDENTIFICATIONThe 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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SRM D/O (3)

GENERAL

The structure repair manual is a non-customized document.It has been prepared in accordance with Air Transport Association ofAmerica (ATA) specification 100.The SRM includes descriptive information as well as specific instructionsand data to perform the assessment of structural damage and to performrepairs. The manual content is approved by the French AirworthinessAuthority DGAC ("Direction Générale de l'Aviation Civile").For most of the damage/defect discovered on the aircraft structure, theSRM is the first document to be used to assess the damage, to identifythe affected structure and to determine the subsequent action or repair tobe performed.


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GENERAL


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SRM D/O (3)

MANUAL BREAKDOWN

The SRM is divided into seven main chapters (From ATA 51 to ATA57) and the SRI (Structure Repair Inspection).The manual also contains an introduction chapter (Chapter 00), and someadditional information pages (HIGHLIGHTS, RECORD OFREVISIONS...) located just at the beginning of the manual.


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MANUAL BREAKDOWN


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SRM D/O (3)

FRONT PAGES

GENERALThe front pages of the manual provides general information relatedto the manual itself:- revision transmittal sheet,- highlights,- record of revisions approved,- record of temporary revisions,- list of effective temporary revisions.


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SRM D/O (3)

FRONT PAGES (continued)

HIGHLIGHTSThe highlights chapter deals with the identification, location of thechanges within the SRM from the previous revision. The type ofchange (page revised, new, deleted) is also presented.


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SRM D/O (3)

INTRODUCTION

GENERALThe introduction chapter contains all necessary information andexplanations to enable a correct use of the manual.


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INTRODUCTION - GENERAL


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SRM D/O (3)

INTRODUCTION (continued)

AIRPLANE ALLOCATION LISTIt also contains the airplane allocation list, giving for each MSN(Manufacturer Serial Number), the airplane type, the aircraft rankwithin the customer version, the customer, the customer abbreviationcode and the registration number.


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SRM D/O (3)


WEIGHT VARIANT INFORMATIONThis section of the introduction chapter enables the identification ofthe aircraft weight variant according to the aircraft serie, engine types,aircraft model and the modification associated to the weight variantchanges.


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SRM D/O (3)


ALPHANUMERICAL INDEXThis document provides a list of all PNs (Part numbers) referencedwithin the SRM and gives the related ATA chapter, the figure number,the configuration and the item number within the figure. This list isupdated at each SRM revision. It provides a quick access to the part,to identify within the manual using the part number as a key.


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SRI (STRUCTURAL REPAIR INSPECTIONS)

GENERALFor permanent repairs with inspection program, inspections are quotedalong with the repair.Due to the amount of inspections, these requirements have beentransferred in a separate appendix to the SRM:- for more clarity of the SRM,- for better handling of the inspection requirements.The chapter Structural Repair Instructions (SRI) gives all necessaryinspection instructions on structural repairs and allowable damagelimits, and provides the airlines with information to integrate theadditional inspections to their own maintenance program.


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SRI (STRUCTURAL REPAIR INSPECTIONS) (continued)

GENERAL (CONT'D)If an inspection is needed, an Inspection Instruction Reference (IIR)will be specified within the related repair. The operator shall thenrefer to the SRI chapter to find out all the inspection instructions usingthe IIR as a key.


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I.I.R. (INSPECTION INSTRUCTION REFERENCE)The explanation of the IIR is given at the beginning of the SRI chapter.


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SRI CONTENTSAt the beginning of the SRI chapter, a table of contents provides thelist of repairs concerned by special inspections requirements.


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SRI CONTENTS (CONT'D)Within the introduction section of the SRI, the "Inspection InstructionSchemes" table enables the operator to locate the inspection instructionwithin the SRI (for inspection ref. 53-41-11-2-001-00, instructionsare to be found in paragraph 1.A.).


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SRI CONTENTS (CONT'D)Within the concerned paragraph, the inspection instruction are dividedinto three main parts:- the general information part, which reminds the inspection location,- the inspection information table (A/C concerned, inspectionthreshold, interval, required inspection methods),- the inspection areas illustrations.


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SRI CONTENTS (CONT'D)-INSPECTION AREASILLUSTRATIONS (IF ANY - E.G. 53-00-11)The areas to be inspected are defined on the illustration, with theidentification of the applicable inspection methods and applicableNTM procedures.


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MANUAL GENERAL USAGE PROCEDURE

When damage is discovered, the first step is to evaluate, classify andaccurately measure it. The SRM chapter 51-11-XX provides usefulinformation to perform this assessment in the best conditions (damagedefinitions and classification, classification of structures, allowabledamage definitions and damage/defect reporting process).


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MANUAL GENERAL USAGE PROCEDURE (continued)

CONT'DThe next step is the full identification of the affected area/structure.This is achieved using the identification page block (pages 01-99) ofthe related specific chapter/section (52-57). According to the originalstructure data and the actual damage characteristics, it is then possibleto determine whether the damage is within the defined allowable limitsor not. This is done using the allowable damage page block (pages101-199) of the related specific chapter/section. If the damage is withinthe allowable limits, the subsequent actions are generally a slightrework and a re-protection of the affected area, using the standardprocedures of chapter 51.


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MANUAL GENERAL USAGE PROCEDURE (continued)

CONT'DIf the damage is above the limits, you must check whether a repair isavailable and/or applicable within the repair page block (pages201-999).


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NUMBERING SYSTEM AND PAGE BLOCK ALLOCATION

Each subject, within the SRM, is identified using a three-elementnumbering system chapter/section and sub-section:- the first element designates the chapter which is assigned by the ATAspec. 100,- the second element designates the section within the chapter. The firstdigit is assigned by the ATA spec. 100. The second digit is assigned byAirbus S.A.S,- the third element identifies the subsection (subject) within the sectionand is assigned by Airbus S.A.S.A standard page block allocation is used for all SRM chapters:- pages 1 to 99 for structure identification,- pages 101 to 199 for allowable damage,- pages 201 to 999 for repairs.


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CHAPTER 51 (STRUCTURE GENERAL)

Information of a general nature or information applicable to more thanone chapter, is included in chapter 51.

NOTE: Note: the entry point within the SRM is always the specificchapter 52 to 57, depending on the damage location.


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CHAPTER 51 (STRUCTURE GENERAL) (continued)

CONTENTSThe table of contents is detailed in the following illustrations.


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CONTENTS (CONT'D)Chapter 51 table of contents (2/3).


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CONTENTS (CONT'D)Chapter 51 table of contents (3/3).


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CHAPTER 52 TO 57 CONTENTS

LAYOUTChapters 52 to 57 have the same layout, which conforms to the definedpage block allocation system (PB 1 to 99-identification, PB 101 to199-allowable damage, PB 201 to 999-repairs). In addition, a tableof contents and a Service Bulletin (SB) list are provided at thebeginning of each chapter. Depending on the chapters, theModification/Service Bulletin list is to be found either at the chapterlevel, or main section level.


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CHAPTER 52 TO 57 CONTENTS (continued)

SERVICE BULLETIN LISTLocated at the beginning of each chapter, the SB list is a list of allservice bulletins referenced within the chapter.


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SERVICE BULLETIN LIST (CONT'D)The user can refer to this list to get more information regarding aspecific service bulletin in terms of:- revision status of the SB,- date of introduction within the SRM,- description.


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MODIFICATION/SERVICE BULLETIN LISTLocated at:- door level for chapter 52,- fuselage section level for chapter 53,- chapter level for chapter 54,- main assembly level for chapter 55,- wing section level for chapter 57, this list is provided to enable theuser to determine the effectivity of the modification/SB.


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MODIFICATION/SERVICE BULLETIN LIST (CONT'D)This list provides, for a given modification number, its associatedsuffix and the aircraft standard, and the effectivity expressed in MSN.


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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99)

EXAMPLE : COMPOSITE STRUCTURESIn the identification pages, the individual parts of the majorcomponents are illustrated and listed in tabular form. Eachidentification topic begins with an introduction page, which includesa general information paragraph. The item number is the key betweenthe illustration and the identification table. For composite structures,the illustration provides identification of individual layers, orientationand materials. A ply (layer) orientation reference is also given.


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IDENTIFICATION PAGE BLOCK (PAGES 01 TO 99)(continued)

EXAMPLE : METALLIC STRUCTURESFor metallic structure such as fuselage skin panels, the differentmaterial thicknesses are provided using letter codes or shaded areasas a key to the thickness tables. The associated identification tableprovides the additional material, part number modification statusinformation.


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IDENTIFICATION TABLE DETAILEDThe typical identification table associated with the illustration(s)contains different columns:- ITEM,- NOMENCLATURE,- SPECIFICATION AND/OR SECTION CODE,- THICKNESS IN MM (IN.) AND/OR PARTNUMBER,- IC: INTERCHANGEABILITY,- ACTION OR REPAIR,- STATUS (MOD/PROP) SB/RC (MOD/PROP:Modification/Proposal, RC: Recordable Concession),In addition: the relevant assembly drawings are listed at the end ofthe table (ASSY DRAWINGS:...........).


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IDENTIFICATION TABLE DETAILED (ITEM COLUMN)The item number is the link with the associated illustrations. Thiscolumn also indicates the different evolutions of the same item (witha suffix letter: 1A, 1B,..) compared to the basic version (withoutsuffix). Each evolution is linked to a production modification givenin the column ''status (MOD/PROP)''.


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IDENTIFICATION TABLE DETAILED(SPECIFICATION/SECTION CODE COLUMN)This column provides information related to the type of material used(material specification), and when possible/available, the appropriatestandards. To get more information regarding the materialspecifications the user can refer to the SRM chapter 51:- 51-31-00 for metallic materials,- 51-33-00 for non-metallic materials (composite materials).When the appropriate standard is shown, the user can refer to theAirbus S.A.S. Standard Manual(SM).


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CHAPTER 51-31-00 EXAMPLEThis example shows the way to proceed to find out the materialinformation within the chapter 51-31-00 for metallic structure, startingwith the material specification. A first list provides the table to beused (e.g. table 4 for 3.1364T42).


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CHAPTER 51-31-00 EXAMPLE (CONT'D)The material table provides, for a given material specification, theEuropean substitute (if used by the other European manufacturers andthe United States substitutes).


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STANDARD MANUAL EXTRACTThe standard manual provides the additional information linked to agiven standard number.


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IDENTIFICATION TABLE DETAILED(THICKNESS/PART NUMBER COLUMN)The Part Number (PN) corresponding to the structural component isshown in this column. Within the PN, the first nine characters givethe detailed drawing number of the component, which is an entry keyto the airbus drawing set. In general the "as drawn" parts are LeftHand (LH) and are provided with an even part number (e.g.: ...202).In the SRM, the identification table always states the LH part numberon the first line and the RH part number (when indicated) on thesecond line.


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IDENTIFICATION TABLE DETAILED (IC COLUMN)The IC column provides the interchangeability status of the parts:- 01: one way interchangeable (post-mod part to be used to replacepre-mod part),- 02: two ways interchangeable (post-mod or pre-mod parts can beused),- 03: no interchangeability.


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IDENTIFICATION TABLE DETAILED (ACTION ORREPAIR COLUMN)This column gives indication concerning the action or repair to beperformed. For existing general or specific repairs, thechapter/section/sub-section is inserted. For a recommendation toreplace the part, the word "REPLACE" is inserted. Where left blank,a case-by-case assessment has to be performed to determine therelevant corrective action (part replacement, repair as per SRM orspecific repair as per Airbus SAS definition).


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IDENTIFICATION TABLE DETAILED (STATUSMOD/PROP SB/RC COLUMN)Linked with the Item column, the STATUS MOD/PRP, SB/RC columngives the modification or service bulletin driving the differentevolutions of a structural component (listed in column Item). A prefixletter is used to identify the status BEFORE ("B" letter) or AFTER("A" letter) SB or Modification. The suffix letter (A, B, C, D...)indicated at the end of the MOD/PROP (and column "S" of theMOD/SB list) shows the different effectivity within the sameMOD/PROP number.


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IDENTIFICATION TABLE DETAILED (STATUSMOD/PROP SB/RC COLUMN (CONT'D))To find the relevant effectivity linked to a modification shown in theSTATUS column, the user must refer to the modification/servicebulletin list.

NOTE: the status before or after modification/SB and the relevantmodification solution (suffix letter) should not be forgotten.Within the modification/service bulletin list, the effectivitiesare expressed in MSN.to find the relationship between the customer version number(e.g.. AFR 01 0016) and the MSN, the user can refer to theairplane allocation list of the introduction chapter of theSRM.


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IDENTIFICATION TABLE DETAILED (STATUSMOD/PROP SB/RC COLUMN (CONT'D))When the modification is linked to a SB, the SB number is alsomentioned, below the modification number. The SB list located at thebeginning of the chapter can be used to get more information.


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CONFIGURATIONS

GENERALWhen modification or service bulletin are extensive and the precedingmethod of reflecting effectivity becomes cumbersome, thus distractingfrom the continuity of subject matter, additional page blocks areestablished. These additional page blocks are further identified by theintroduction of a configuration code (CONFIG-1, CONFIG-2), etc...)following the chapter/section/sub-section. Configuration codes arealways in ascending, sequential numerical order, i.e., CONFIG-1,CONFIG-2, CONFIG-3, etc... The first example shown belowillustrates the change in the A320 vertical stabilizer spar box designand manufacturing following modification 26500K4924H.


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CONFIGURATIONS (continued)

EXAMPLEHere is an example of different configurations. The modification ofthe design and production process of the carbon fiber vertical stabilizerspar box leads to two configurations within the chapter 55. Theconfiguration number is indicated at the bottom of the related pages.On this example, as highlighted, the configuration 1 is applicablebefore modification 26500K4924H. Configuration 2 is applicableafter modification 26500K4924H. It is of the operator duty to selectfirst in which configuration the concerned aircraft is, before goingfurther in the SRM investigation. This is done using the servicebulletin/modification list located at the beginning of the chapter orsection.


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ALLOWABLE DAMAGE PAGE BLOCK (PB 101)

GENERALThe information to be found within allowable damage page blockenables the operator to define whether a damaged airplane may bereturned into service without repair. An allowable damage permittedhas no significant effect on the strength or fatigue life of the structure,which must still be capable of fulfilling its function. Allowable damagemay require minimal rework such as cleanup or drilling of stop holes.Basically the allowable page block contains different page types:- general information pages,- damage criteria tables,- paragraph for each type of damage,- damage measurement procedure,- damage localization (zoning) figures,- allowable damage diagram.


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ALLOWABLE DAMAGE PAGE BLOCK (PB 101) (continued)

ALLOWABLE DAMAGE INFORMATIONAfter a careful reading of the first information page(s), the first stepin the allowable damage determination, consists in the use of thedamage criteria table.


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DAMAGE CRITERIA TABLEThe damage criteria table enables the operator to determine the relevantparagraph (e.g. 4A) that should be used, depending on the type ofdamage (e.g.: allowable rework), and the affected structure (e.g. skinplates).


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ALLOWABLE DAMAGE EFFECTIVITYFor some allowable damage (e.g. dent allowable damage), theallowable damage effectivity per weight variant must be checked.The table (e.g. table 102) at the beginning of the allowable damagerelevant paragraph shall be used. The allowable damage informationcontained within the related paragraph is justified and applicable forthe listed weight variant only. If not, contact Airbus.


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WEIGHT VARIANT DETERMINATIONTo define the "as delivered" weight variant of a concerned MSN, thefirst step is to refer to the airplane allocation list of the introductionchapter, which gives the corresponding type (e.g.: A320-232). Then,refer to the weight variant identification list, also in the introductionchapter.


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WEIGHT VARIANT DETERMINATION (CONT'D)The weight variant identification list gives, for a given aircraft type,the different possible "as delivered" weight variants. The next stepwill be to determine which one of these possible weight variants isapplicable to the concerned MSN.


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WEIGHT VARIANT DETERMINATION (CONT'D)To determine the applicable "as delivered" weight variant, the operatormust check which of the associated modification(s) is effective on theconcerned MSN. The Aircraft Inspection Report (AIR) can be usedfor this purpose. Once found, the corresponding weight variant isconsidered as the weight variant of concerned MSN at delivery.


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WEIGHT VARIANT DETERMINATION (CONT'D)The "as delivered" weight variant may change after delivery, followingthe embodiment of a service bulletin. It is the operator's responsibilityto check the embodiment of referenced SB, in order to determine therelevant weight variant for the affected MSN. The information is giveninto Table 2 "Service Bulletin/Weight Variant List".


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WEIGHT VARIANT DETERMINATION (CONT'D)The table 2-"Service Bulletin/Weight Variant List" provides for therelated aircraft type (e.g.: A320-232) the different service bulletinapplicable. For the given MSN, the operator must check whether oneof the possible SB has been embodied or not. If yes, the weight variantof the aircraft becomes the weight variant given by the table.


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WEIGHT VARIANT DETERMINATION (CONT'D)For aircraft which have been subject to weight variant change througha SB embodiment: the weight variant information to be used to identifythe effectivity of the given allowable damage (or repair information),is the heaviest weight variant that the subject aircraft has been operatedwith, since its delivery.


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ALLOWABLE DAMAGE DETERMINATION (CONT'D)The allowable damage diagram to be used, generally depends on thelocation of the damage on the concerned structure.Illustrations are used to locate the damage and thus to define therelevant diagram to be used (e.g. refer to diagram 102).


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ALLOWABLE DAMAGE DIAGRAM EXAMPLEAs a last step of the damage investigation, the use of the allowabledamage diagrams provides the necessary information concerning theactions to be performed, depending on damage extent.


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ALLOWABLE DAMAGE DIAGRAM EXAMPLE(CONT'D)Provided that no cracks have been detected, this first area of theallowable damage rework diagram defines typical allowable damagewithout any time limits. The surface protection needs to be restored.


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ALLOWABLE DAMAGE DIAGRAM EXAMPLE(CONT'D)This second area for reworks depth between 10 % and 25 % of thenominal thickness, also defines allowable damage but with a timelimit.In this example a repair has to be performed before 3000 flights atthe latest. The surface protection has to be restored.


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ALLOWABLE DAMAGE DIAGRAM EXAMPLE(CONT'D)This area is defined for reworks depth between 25 % and 40 % of thenominal thickness. It is still allowable damage, provided that no crackis detected. But a repair has to be performed before 50 flights at thelatest.


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ALLOWABLE DAMAGE DIAGRAM EXAMPLE(CONT'D)Stop drill and apply high-speed tape for one flight pressurized orunpressurized only, or do a temporary repair as per 53-00-11 figure210.Temporary repair has to be replaced by a final repair within 2500flights.


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ALLOWABLE DAMAGE DIAGRAM EXAMPLE(CONT'D)For damage located in this area, the damage shall be stop drilled beforea ferry flight without cabin pressure. Install high-speed tape beforethe ferry flight. A repair has to be performed.


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ALLOWABLE DAMAGE DIAGRAM EXAMPLE(CONT'D)For damage located in this area, a repair has to be performedimmediately or a ferry flight may be allowed upon manufacturer'sauthorization.


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REPAIRS PAGE BLOCK (PB 201)

GENERALThe Repairs Page Block (PB 201), contains necessary information tocarry out permissible repairs.Each of the repairs is described with illustrations and procedureinstructions, which includes repair applicability data and repairmaterial lists.


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REPAIRS PAGE BLOCK (PB 201) (continued)

LIST OF AVAILABLE REPAIR SCHEMESAt the beginning of each repair page block, a list of available repairschemes 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 withinthe related section when the repair is specific (e.g. paragraph 5A, Fig.201).


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REPAIR INSTRUCTIONSThe relevant paragraph of the repair instructions contain informationrelated to:- the applicability of the concerned repair,- general reminders linked to repair principles and rules.Note: an IIR appears within the repair instructions.


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REPAIR INSTRUCTIONS (CONT'D)All the materials (repair materials and consumable materials) are listedwithin the concerned repair instructions. Consumable materials arecall-up using their Consumable Material List (CML) code (e.g.09-013). For more information on these materials, the user can referto the SRM chapter 51-35-00 consumable materials and/or the CMLdocument.


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REPAIR INSTRUCTIONS (CONT'D)this SRM chapter(51-35-00) deals with consumable materials.


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REPAIR INSTRUCTIONS (CONT'D)Information concerning the consumable materials can also be foundin the CML.


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REPAIR INSTRUCTIONS (CONT'D)The repair instruction lists all the steps of the repair, with referencesto the standard processes and practices covered within the chapter 51when required.


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REPAIR ILLUSTRATIONSNote: an IIR appears within the repairs illustrations.


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WINDOW DAMAGE ASSESSMENT D/O (3)

GENERAL

The windows are:- the cockpit windows ( windshields, sliding windows, and aft fixedwindows),- the cabin windows and,- the passenger/crew door windows.To find the corrective actions for each type of defect, refer to the pageblock 6xx of the relevant AMM chapters:- AMM 56-11-11 for the windshields,- AMM 56-12-11 for the sliding windows,- AMM 56-11-12 for the aft fixed windows,- AMM 56-21-13 for the cabin windows,- AMM 56-31-00 for the passenger/crew door windows.


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GENERAL




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INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS

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




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INVESTIGATION OF DAMAGE ON COCKPIT WINDOWS(continued)

INVESTIGATION OF DAMAGE ON WINDSHIELDSRefer to AMM 56-11-11 pages block 6xx to get the corrective actionsfor the following types of defect on windshields:- cracks,- scratches,- chips,- delaminating,- discoloration,- interlayer micro flakes,- bubbles,- burn spot,- transparency,- rain repellent fluid residue,- damage on the soft liner (if supplied by SPS company only).




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INVESTIGATION OF DAMAGE ON SLIDING WINDOWSRefer to AMM 56-12-11 pages block 6xx to get the corrective actionsfor 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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INVESTIGATION OF DAMAGE ON AFT FIXEDWINDOWSRefer to AMM 56-11-12 pages block 6xx to get the corrective actionsfor 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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INVESTIGATION OF DAMAGE ON CABIN WINDOWSAND PASSENGER/CREW DOOR WINDOWS

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




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INVESTIGATION OF DAMAGE ON CABIN WINDOWSAND PASSENGER/CREW DOOR WINDOWS (continued)

INVESTIGATION OF DAMAGE ON CABIN WINDOWSRefer to AMM 56-21-13 page block 6xx to get the allowable damagesfor the following types of defect on cabin windows:- delaminating,- scratches,- pitting,- crazing,- crazing with bulging,- bulging,- orange peel effect,- chipping,- cracks,- vent hole damage.




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INVESTIGATION OF DAMAGE ON CABIN WINDOWSAND PASSENGER/CREW DOOR WINDOWS (continued)

INVESTIGATION OF DAMAGE ON PASSENGER/CREWDOOR WINDOWSRefer to AMM 56-31-00 page block 6xx to get the allowable damagefor the following types of defect on passenger/crew door windows:- scratches,- crazing,- bulging,- crazing with bulging,- delaminating.




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DAMAGE ASSESSMENT EXAMPLE 1 D/O (3)

INTRODUCTION

The purpose of this example is to present, the complete procedure to befollowed when a damage is discovered, from the damage mapping draftto the final structure damage assessment. This example was chosen as itrepresents one of the more usual types of damage on an A/C and givesan in depth investigation with all the different stages.


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INTRODUCTION




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IDENTIFICATION OF THE DAMAGE

The damage is located onto the fuselage skin, thus, all the informationregarding the identification of the part, allowable damage and repair, ifany, are to be found within the chapter 53 of the SRM. Informationconcerning the damage classification and reporting are to be found withinthe SRM chapter 51-11-00. The concerned damage is a dent with novisible crack. At this stage, take visual reference to facilitate damagelocation. Such as, a forward or aft passenger door, or a cargo door, aboveor below cabin floor level at stringer (Stgr) 23, close to a longitudinal orcircumferential joint, etc...). If the dent is close to a rivet row, an internalvisual inspection is required to determine whether the internal structure(frame, stringer, etc...) is also damaged or not.




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MAPPING

Using SRM 51-11-13 as a guide, the maximum information should betaken from the aircraft before starting any assessment (measurement andlocation of the maximum depth, distance of dent edges to nearest fastenerrows, existing closest skin joints or any other visible structure that willhelp in the detailed location of the damage, etc...).




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MAPPING (continued)

CONT'DUsing the data collected from the A/C, the mapping should becompleted by determining the exact location (in terms of framenumbers and stringer numbers). For this purpose, refer to the beginningof the chapter 53 (fuselage).




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MAPPING (continued)

CONT'DThe illustration of chapter 53-00-00 enables the operator to determinethe circumferential joint related frame numbers.




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MAPPING (continued)

CONT'DUsing the frame identification illustration of chapter 53-00-00, andthe data collected during the damage mapping, the frames surroundingthe damage can be determined. According to the mapping information,the damage is located between the first and the second frame after thecircumferential joint located at Fr 24. Consequently, the damage islocated between Fr 25 and 26.




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MAPPING (continued)

CONT'DTo complete the damage location, the stringers surrounding the damagealso need to be determined. For this purpose the "General panelidentification" illustrations, proposed within chapter 53-00-00 can beused. According to the data collected on the A//C and the location ofthe damage from the existing longitudinal skin joints, the affectedpanel can be determined. For this example the damage is located onpanel 7 - lower side shell - located between Stgr 18LH & 32LH, andFr 24 & 35.




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MAPPING (continued)

CONT'DThe information collected can be reported onto the damage mapping.The damage is located between Fr 25 and 26. The stringer numbercorresponds to the longitudinal skin joint from which the damage hasbeen located. Nevertheless, the exact stringer numbers surroundingthe damage need to be confirmed. For this purpose, the informationprovided in the identification page block of the concerned panel hasto be used.




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DETAILED IDENTIFICATION

The "fuselage section division" illustration of chapter 53-00-00 usedbefore enables the definition of the affected section: "Forward fuselage"- section 13/14 - chapter 53-20-00. The general illustration of 53-20-00identifies the main structural arrangement of the forward fuselage. Skinplates are part of the "MAIN STRUCTURE" covered by section 53-21-00.




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DETAILED IDENTIFICATION (continued)

CONT'DFollowing SRM 53-21-00 guidelines, the figure shows that the skinpanels (skin plates) are item number 1. The illustration associatednomenclature informs us that the full identification of the skin panels(skin plates) are covered by SRM 53-21-11.




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DETAILED IDENTIFICATION - CONT'D




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CONT'DAll the skin panels (plates) of the forward fuselage are listed withinthe nomenclature located at the front page of SRM 53-21-11. Usingthe information collected before (affected panel: lower side panel -left, between Fr 24 & 35 and Stgr 18 & 32), the nomenclature providesthe figure number we have to refer to: "Skin plates - LWR parts LHFr 24 to Fr 35: REFER TO Figure 1".




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CONT'DThe figure 1 identifies two different panels (view A and C). The viewA concerns the skin panel located from Stgr 18LH to 32LH. The viewC concerns the skin panel located from Stgr 32LH to 41LH. Accordingto the damage mapping, the view A is concerned. The damage hasbeen located on panel 7 (between Stgr 18LH and 32LH).




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CONT'DThe view A, identifies all the different items which are part of thepanel (e.g. crack stoppers, doublers...), and a view indication identifiesthe skin itself for more details (View D). The view D identifies thedifferent material thicknesses (letter codes), and all the stringerlocations.There are two different panel configurations illustrated, showing thebasic version of the panel and an other possible version effective afterthe embodiment of production modification(s). Modification numbersare indicated at the bottom of the page. The next step of theinvestigation is to define which of these panels is installed on theconcerned A/C.




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CONT'DTo identify the actual panel, the modification numbers indicated atthe bottom of the page have to be compared with the servicebulletin/modification list, located at the beginning of chapter 53-20-00.The purpose is to check their effectivity in terms of ManufacturerSerial Number (MSN).




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CONT'DThe view D of panel on figure 1, sheet 5 is valid after Modifications(MODs) 27117P5234 or 2729P5353. Checking the Modification / SBList at the beginning of chapter 53-20-00, MSN 2057 doesn't appearin this list of MSN proposed for each of the modification. So the panelinstalled on the A/C is a basic version, then, refer to view D figure 1,sheet 3.




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CONT'DThe damage is located between Fr 25 and 26, and is located betweenthe fourth and the fifth stringer from Stgr 18 LH (longitudinal skinjoint reference).This information can be reported onto the illustration and gives:- the material thickness of the area (code B, giving 1.4 mm (0.055in)),- the stringer location: the damage is located between Stgr 22LH and23LH.




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MAPPING (FINALIZATION)The damage mapping can now be completed with the stringer numbersand the nominal skin thickness in the dented area. The damageassessment using the allowable damage page block is the next step.




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DETAILED IDENTIFICATION - MAPPING (FINALIZATION)




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DAMAGE ASSESSMENT

GENERALTo start the damage assessment refer to the page block 101 of therelevant chapter/section (53-21-11), and start to read carefully theprocedure. A special attention shall be paid to the notes and cautions.




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DAMAGE ASSESSMENT - GENERAL




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DAMAGE ASSESSMENT (continued)

WEIGHT VARIANTA caution note indicates that the allowable damage effectivity perA/C weight variant may have to be verified. The weight variant is acriterion which is defined for each model of A/C and depending onits Maximum Take Off Weight (MTWO), Maximum Landing Weight(MLW), Maximum Zero Fuel Weight (MSFW). The allowable damagelimits are defined per weight variant and for a same model. The weightvariant can change, depending on the modification or Service Bulletin(SB) embodiment status. The actual weight variant of the affectedA/C has to be known before starting the assessment. Because of themodifications, which could be embodied on the A/C, only the airlineengineering department shall give you this information. The actualweight variant shall be compared with the data given in a table at thebeginning of allowable damage related paragraph.




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DAMAGE ASSESSMENT - WEIGHT VARIANT




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DAMAGE CRITERIAA second caution note indicates that in some cases, an inspection maybe required to check for crack, even if the damage is determined asbeing allowable.




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DAMAGE ASSESSMENT - DAMAGE CRITERIA




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DAMAGE CRITERIA (CONT'D)Check the applicability of the allowable damage for dents.




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DAMAGE ASSESSMENT - DAMAGE CRITERIA (CONT'D)




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DAMAGE CRITERIA (CONT'D)To keep on with the damage assessment procedure, a note asks theoperator to refer to the damage criteria table 101.




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DAMAGE CRITERIA (CONT'D)The allowable damage description/criteria table (101), shows twotypes of dents:- dent* referring to paragraph 4B,- dent ** referring to paragraph 4C.Note that an Inspection Instruction Reference (IIR) is indicated fordents*. The first step is to define which paragraph is applicable toreported dent. Dents are considered as fulfilling nearness/form criterionor out of nearness/form criterion, in accordance with their geometryand their proximity to the nearest adjacent internal structure elements.This must be determined according to the parameters defined in figure104.




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DAMAGE CRITERIA (CONT'D)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 measuredfrom the dent edge to any fastener row (frame, stringer) distance B.This distance should be minimum 15 mm (0.59 in),- the second criterion consists in comparing the depth of the dent (D)with the smallest distance measured from the deepest point of the dentto the closest adjacent structure (distance A).The depth of the dent should be maximum 10% of the distance A. Ifone of these criteria is not met, the dent **, and thus paragraph 4C(dent not fulfilling criteria) should be taken to keep on with paragraph4B.




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DAMAGE CRITERIA (CONT'D)The second criterion consists in comparing the depth of the dent (D)with the smallest distance, measured from the deepest point of thedent to the closest adjacent structure (distance A). If no access frominside, the measurement is taken from outside, from the deepest pointof the dent to the closest fastener row (distance X). The distance Awill become distance X - 15 mm, which is the average considerededge margin.




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DAMAGE CRITERIA (CONT'D)Check for the first criterion to be fulfilled. B distance : minimum 15mm. The smallest distance measured between the edge of the dentand the surrounding fastener rows is 29 mm, which is higher than 15mm. The first criterion is met.




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DAMAGE CRITERIA (CONT'D)Check for the second criterion to be fulfilled: the depth of the dentshould be maximum 10% of the distance A. The distance measurementhas been done from outside: the smallest distance between the deepestpoint of the dent and the surrounding fastener row is 66 mm. Sincemeasured from outside, distance A = 66 mm - 15 mm = 51 mm; 10%of A = 5.1 mm. The second criterion is also fulfilled since the depthof the dent (D = 4.5 mm) is smaller than 10% of A.




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DAMAGE CRITERIA (CONT'D)The dent fulfils "nearness/form criterion", then refer to paragraph 4.B.




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DAMAGE CRITERIA (CONT'D)As mentioned in a caution at the beginning of the allowable damagepages, the allowable damage applicability have to be checked, usingthe weight variant table (table 102) given at the beginning of theparagraph. The information coming from the airline-engineeringdepartment shows that the MSN 2057 is at weight variant 001.Checking the table 102, weight variant 001 is included in, and thusthe following allowable damage information can be used.




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DAMAGE CRITERIA (CONT'D)Compare the dents in accordance with diagram 103.




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ALLOWABLE DENT DIAGRAMThe skin thickness in the dented area, and the depth of the dent, arethe keys to get into to diagram. You must refer to the data collectedbefore (damage mapping).




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DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM




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ALLOWABLE DENT DIAGRAM (CONT'D)The skin thickness in the dented area is 1.4 mm (found in theidentification pages). The depth of the dent is 4.5mm (measured fromthe A/C damage mapping). These two values are plotted onto thediagram, which defines a point. The area where this point is locateddefines the subsequent actions to be performed. For the concerneddent, the actions to be performed are as follow: "check damage forcracks by detailed visual examination. If clear, repair within 3000FC". Provided that no crack is detected by detailed visual inspection,the dent is considered as an allowable damage with a time limit(temporary allowable damage). The A/C can be released. But a repairwill have to be done before 3000 Flight Cycles (FC).




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DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM (CONT'D)




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INTRODUCTION

The purpose of this example is to present you the complete procedure tobe followed when a damage is discovered, from the damage mappingdraft to the final structure damage assessment. This example was chosenas it represents one of the more usual types of damage on the A/C andenables to make an in depth investigation with all the different stages.




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INTRODUCTION




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The damage is located onto the fuselage skin, thus, all the informationregarding the identification of the part, allowable damage and repair, ifany, are to be found within the chapter 53 of the SRM. All the informationregarding the damage classification or the rework, if any, are to be foundwithin SRM chapter 51. The applicable damage is a scratch with novisible crack. At this stage: take visual reference to facilitate damagelocation. 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 alongitudinal or circumferential skin joint, etc...If the scratch is near a rivet row, an internal visual inspection is requiredto determine whether the internal structure (frame, stringer, etc...) is alsodamaged or not.




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MAPPING

Using the SRM 51-11-13 as a guide, the maximum information shouldbe taken from the A/C before starting any assessment. (measurement andlocation of the maximum depth, distance of rework edges to the nearestfastener rows, existing closest skin joints or any other visible structurethat will help in the detailed location of the damage, etc...).




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MAPPING




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MAPPING (continued)

CONT'DUsing the data collected from the A/C, the mapping should becompleted by determining the exact location (in terms of framenumbers and stringer numbers). For this purpose refer to the beginningof chapter 53 - fuselage.




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MAPPING - CONT'D




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MAPPING (continued)

CONT'DThe illustration of chapter 53-00-00 enables the operator to determinethe circumferential skin joints related frame numbers.




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MAPPING - CONT'D




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MAPPING (continued)

CONT'DUsing the frame identification illustration of chapter 53-00-00, andthe data collected during the damage mapping, the frames surroundingthe damage can be determined.According to the mapping information, the damage is located betweenthe first and the second frame after the circumferential joint locatedat frame (Fr) 64. Consequently the damage is located between Fr 62and 63.




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MAPPING - CONT'D




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MAPPING (continued)

CONT'DTo complete the damage location, the stringers surrounding the damagealso need to be determined. For this purpose, the "General PanelIdentification" illustrations, proposed within chapter 53-00-00 can beused. According to the data collected on the A/C and the location ofthe damage from the existing longitudinal skin joints, the affectedpanel can be determined. For this example, the damage is located onpanel 8 - lower side shell - located between Stgr 18 & 32, and Fr 47& 64.




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MAPPING - CONT'D




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MAPPING (continued)

CONT'DThe collected information can be reported onto the damage mapping.The damage is located between Fr 62 and 63. The stringer numbercorresponds to the longitudinal skin joint from which the damage hasbeen located. Nevertheless, the exact stringer numbers surroundingthe damage need to be confirmed. For this purpose, the informationprovided in the identification page block of the concerned panel hasto be used.




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MAPPING - CONT'D




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The "Fuselage Section Division" illustration of chapter 53-00-00 usedbefore, enables the definition of the affected section: rear fuselage -section 17 - chapter 53-40-00". The general illustration of 53-40-00identifies the main structural arrangement of the forward fuselage. Theskin plates are part of the main structure covered by the section 53-41-00.




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CONT'DFollowing SRM 53-41-00 guidelines, the figure shows that the skinpanels (skin plates) are item number 5. The illustration associatednomenclature informs us that the full identification of the skin panels(skin plates) are covered by SRM 53-41-11.




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CONT'DAll skin panels (plates) of the forward fuselage are listed within thenomenclature located at the front page of SRM 53-41-11. Using theinformation collected before (affected panel : lower side panel - left,between Fr 47 & 64 and Stgr 18 & 32), the nomenclature providesthe figure number we have to refer to: "Skin plates - LWR parts LHFr 47 to Fr 64: Refer To Figure 6".




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CONT'DThe figure 6, identifies two different panels configurations (view Aand 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, showingthe basic version of the panel and an other possible version effectiveafter the embodiment of production modification(s). The modificationnumbers are indicated at the bottom of the page. The next step of theinvestigation is to define which of these panels is installed on theconcerned A/C.




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CONT'DTo identify the actual panel, the modification numbers indicated atthe bottom of the page have to be compared with the servicebulletin/modification list (located at the beginning of chapter53-40-00). In this list, the effectivity in terms of Manufacturer SerialNumber (MSN) has to be checked.




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CONT'DThe view B of the panel on figure 1, sheet 5 is valid after Modifications(MODs) 21468K1489A, 22083K2232B, 24958K4082D or31012K7082. Checking the Modification/Service Bulletin (SB) Listat 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 toview B figure 6, sheet 2.




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CONT'DThe damage is located between Fr 62 and 63, and is located betweenthe sixth and the eighth stringer from Stgr 18 (longitudinal skin jointreference).This information can be reported onto the illustration and gives:- the material thickness of the area (code C, giving 1.6 mm (0.063 in),- the stringer location: the damage is located between Stgr 23LH and25LH.




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MAPPING (FINALIZATION)The damage mapping can now be completed with the stringer numbersand the nominal skin thickness in the scratched area. The damageassessment using the allowable damage page block is the next step.




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DAMAGE ASSESSMENT

To start the damage assessments refer to the page block 101 of the relevantchapter/section (53-41-11). And start to read carefully the procedure. Aspecial attention shall be paid to the notes and cautions.




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DAMAGE ASSESSMENT




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DAMAGE CRITERIARead carefully all the cautions, they could give you information onthe assessment. A second caution note indicates that in some cases,an inspection may be required to check for crack, even if the damageis determined as being allowable.




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DAMAGE CRITERIA (CONT'D)In the allowable damage description/criteria table (101), the paragraph4A has to be acknowledged for reworks.




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DAMAGE CRITERIA (CONT'D)Check the applicability of the allowable damage for reworks.




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WEIGHT VARIANTA caution note indicates that the allowable damage effectivity perA/C weight variant may have to be verified. The weight variant is acriterion, which is defined for each model of A/C and depending onits Maximum Take Off Weight (MTOW), Maximum Landing Weight(MLW), and Maximum Zero Fuel Weight (MZFW). The allowabledamage limits are defined per weight variant and for a same model.The weight variant can change depending on modification or SBembodiment status. The actual weight variant of the affected A/C hasto be known before starting the assessment. Because of themodifications, which could be embodied on the A/C, only theairline-engineering department shall give you this information. Theactual weight variant shall be compared with the data given in a tablein introduction pages.




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WEIGHT VARIANT (CONT'D)As mentioned in a caution at the beginning of the allowable damagepages, the allowable damage applicability has to be checked, usingthe weight variant exclusion table (table 4) given at the introductionof the SRM. The information coming from the airline-engineeringdepartment shows that the MSN 2042 is at weight variant 010.Checking the table 4, the weight variant 010 is included in and thus,the following allowable damage information can be used.




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DAMAGE ASSESSMENT - WEIGHT VARIANT (CONT'D)




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DAMAGE CRITERIA (CONT'D)Two diagrams are given, one for riveted areas and one for unrivetedareas.




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DAMAGE CRITERIA (CONT'D)The parameters of the rework have to be determined to be sure thatwe are dealing with the correct diagram.




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DAMAGE CRITERIA (CONT'D)To determine which diagram to use, we have to check if the damageis located in a riveted area. A riveted area extends from less than 15mm (0.590 in) all around a rivet. In the applicable damage, the rivetedarea and the unriveted area have to be acknowledged. The maximumdepths of the scratch in riveted and unriveted areas have to beacknowledged too.




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DAMAGE CRITERIA (CONT'D)To be allowable, the rework width has to be equal or longer than fortytimes the depth.




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DAMAGE CRITERIA (CONT'D)To complete the diagram, the maximum depth of the rework has tobe expressed as a percentage of the damaged skin thickness.Those two values (found before) are plotted onto the diagram, whichdefines a point. The area where this point is located defines thesubsequent actions to be performed.




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DAMAGE CRITERIA (CONT'D)Check for the first criterion to be fulfilled: the width of the damagemust be at least 40 x T. The depth of the depression in the riveted areais 0.2 mm; 40*0.2 = 8 mm. The width of the depression is 18.5 mm,which is higher than 8 mm. So, the first criterion is met.




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DAMAGE CRITERIA (CONT'D)To complete the diagram, the maximum depth of the rework in theriveted area has to be expressed as a percentage of the damaged skinthickness.




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ALLOWABLE REWORK DIAGRAMThe depth of the rework as percentage of the skin thickness inunriveted area, and the length of the rework, are the key to get into todiagram. You must refer to the data collected before. These two valuesare plotted onto the diagram, which defines a point. The area wherethis 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). Renewsurface protection and repair after 50 flights at the latest". Providedthat no crack is detected by detailed visual inspection, the rework isconsidered as an allowable damage with a time limit (temporaryallowable damage). The A/C can be released. But a repair will haveto be done before 50 flights.




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DAMAGE ASSESSMENT - ALLOWABLE REWORK DIAGRAM




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ALLOWABLE REWORK DIAGRAM (CONT'D)This diagram enables to determine if the damage is allowable, andthe condition of allowability.




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DAMAGE ASSESSMENT - ALLOWABLE REWORK DIAGRAM (CONT'D)




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DAMAGE ASSESSMENT - CONCLUSION

In the allowable example, two assessments have been done, the morerestrictive one has to be acknowledged. So, the damage has to be checkedfor cracks, damage up to depression depth has to be removed, the surfacehas to be renewed and the A/C has to be repaired after 50 flights at thelatest.




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DAMAGE ASSESSMENT - CONCLUSION




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A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3)

INTRODUCTION

The purpose of this example is to present you, the complete procedureto be followed when a damage is discovered, from the damage mappingdraft to the final structure damage assessment. This example was chosenas it represents one of the more usual types of damage on the A/C andenables to make an in depth investigation with all the different stages.


A318 DAMAGE ASSESSMENT EXAMPLE 3 D/O (3) May 10, 2006Page 582


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INTRODUCTION




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The damage is located onto the fuselage skin, thus, all the informationregarding the identification of the part, allowable damage and repair, ifany, are to be found within chapter 53 of the SRM. Informationconcerning the damage classification and reporting are to be found withinSRM chapter 51-11-00. The applicable damage is a scratch with no visiblecrack.At this stage: take visual reference to facilitate damage location.Such as, forward or an aft passengers door, or a cargo door, above orbelow cabin floor level at stringer (Stgr) 23, near a longitudinal orcircumferential joint, etc...).If the scratch is near a rivet row, an internal visual inspection is requiredto determine whether the internal structure (frame, stringer, etc...) is alsodamaged or not.




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MAPPING

DRAFTUsing the SRM 51-11-13 as a guide, the maximum information shouldbe taken from the A/C before starting any assessment (measurementand location of the maximum depth, distance from dent edges to thenearest fastener rows, existing closest skin joints or any other visiblestructure that will help in the detailed location of the damage, etc...).The damage is located onto the fuselage skin, thus, all the informationregarding the identification of the part, allowable damage and repair,if any, are to be found within the chapter 53 of the SRM.

NOTE: on the affected panel, there are no stringer rivet rows, thus,stringers, if any, should be welded onto the skin: it is notpossible to identify the stringer references.

As a consequence, it is necessary to measure the distance from alongitudinal skin joint to the dent maximum depth, in order to get areference for the location of the dent. This reference will be comparedwith the welded stringer references coming from the SRM, page blocks101 and 201 (see next pages).




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MAPPING - DRAFT




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MAPPING (continued)

CIRCUMFERENTIAL SKIN JOINT IDENTIFICATIONThe illustration of the chapter 53-00-00 enables the operator todetermine the circumferential joints related frame numbers. In thegiven example, the damage is located in the aft center fuselage section,with the circumferential skin joints at Fr 47/54 and Fr 64.




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MAPPING - CIRCUMFERENTIAL SKIN JOINT IDENTIFICATION




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MAPPING (continued)

CIRCUMFERENTIAL SKIN JOINT IDENTIFICATION(CONT'D)Using the frame identification illustration of chapter 53-00-00, andthe data collected during the damage mapping, the frames surroundingthe damage can be determined. According to the mapping information,the damage is located between the fourth and the fifth frame beforethe circumferential skin joint located at frame (Fr) 64. Consequentlythe damage is located between Fr 59 and 60.




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MAPPING - CIRCUMFERENTIAL SKIN JOINT IDENTIFICATION (CONT'D)




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MAPPING (continued)

LONGITUDINAL SKIN JOINT AND PANELIDENTIFICATION (CONT'D)To complete the damage location, the stringers surrounding the damagealso need to be determined. For this purpose, the "General PanelIdentification" illustrations, proposed within chapter 53-00-00, canbe used. According to the data collected on the A/C and the locationof the damage from the existing longitudinal skin joints, the affectedpanel can be determined. For this example, the damage is located onpanel 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 weldedskin/stringer panel, thus, refer to page block 101 to determine thestringers position.




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MAPPING - LONGITUDINAL SKIN JOINT AND PANEL IDENTIFICATION (CONT'D)




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MAPPING (continued)

WELDED STRINGERS POSITIONFirst, refer to the page block 101 of the relevant chapter/section(53-41-11) and start to read carefully the procedure. Refer to theallowable damage description/criteria table to find the concernedparagraph (4F): "Fuselage Skin Plates Fr 47 / 54 Thru Fr 64 BetweenStgr 32 LH and Stgr 41 RH (Welded Panel)".




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MAPPING - WELDED STRINGERS POSITION




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MAPPING (continued)

WELDED STRINGERS POSITION (CONT'D)Read the notes within the relevant paragraph to find information aboutthe definition and determination of undisturbed skin (unwelded andunriveted).




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MAPPING - WELDED STRINGERS POSITION (CONT'D)




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MAPPING (continued)

WELDED STRINGERS POSITION (CONT'D)This figure shows how unwelded and unriveted areas, welded areas,riveted areas and coupling areas are defined. Two methods ofmeasurement are given, we look at measurement from outside, a flagrefers to SRM chapter 53-41-11 page block 201 to get stringerpositions on a welded panel.




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MAPPING (continued)

WELDED STRINGERS POSITION (CONT'D)This diagram provides the distances from the lap joint 41RH to allwelded stringers, at each frame location. Therefore a new mapping isrequired. The distance from the lap joint (Stgr 41RH) to the dentmaximum depth (at Fr 60) becomes 825 mm.




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MAPPING (continued)

WELDED STRINGERS POSITION (CONT'D)This distance (825 mm) must be compared with the distances fromthe lap joint 41RH to the welded stringers, to locate the dent.




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MAPPING (continued)

WELDED STRINGERS POSITION (CONT'D)We conclude that the dent is located between Stgr 40 and 41LH andit is now possible to finalize the draft (see next page).




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The "fuselage section division" illustration of chapter 53-00-00 usedbefore enables the definition of the affected section: aft center fuselage,part of rear fuselage - section 16/17 - chapter 53-40-00. The generalillustration of the chapter 53-40-00 identifies the main structuralarrangement of the rear fuselage. The skin plates are part of the mainstructure covered by the section 53-41-00.




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DETAILED IDENTIFICATION (CONT'D)Following SRM 53-41-00 guidelines, the figure shows that the skinpanels (skin plates) are item number 5. The identification table informsus that the full identification of the skin panels (skin plates) are coveredby SRM 53-41-11.




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DETAILED IDENTIFICATION - DETAILED IDENTIFICATION (CONT'D)




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DETAILED IDENTIFICATION (CONT'D)The identification table of SRM 53-41-11 refers to the figure 3 sheet2 for the skin panel located between Fr 58A & Fr 64, and Stgr 41RH& Stgr 32LH.




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DETAILED IDENTIFICATION (CONT'D)The damaged panel is illustrated onto two sheets. According to thefigure 3 sheet 1, the damaged panel is item number 1.




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DETAILED IDENTIFICATION (CONT'D)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 nomenclaturetable so, the skin panel of MSN 2218 is the Part Number (P/N)D53479410202.




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DETERMINATION OF SKIN THICKNESS IN DENTEDAREAThe damage is located between Fr 59 & 60, and Stgr 40LH & 41LH.This information can be reported onto the illustration and gives theskin thickness in the dented area (code B, giving 1.6 mm (0.063 in)).




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DETAILED IDENTIFICATION - DETERMINATION OF SKIN THICKNESS IN DENTED AREA




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MAPPING (FINALIZATION)The damage mapping can now be completed with the stringer numbersand the nominal skin thickness in the dented area. As the mapping isat scale 1:1, we can measure the distance between Stgr 40LH and thedeepest point of the dent (59 mm), and the distance between Stgr 40and the edge of the dent (49 mm). The damage assessment using theallowable damage page block 101 is the next step.




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DAMAGE ASSESSMENT

GENERALTo start the damage assessments refer to the page block 101 of therelevant chapter/section (53-41-11), and start to read carefully theprocedure. A special attention shall be paid to the notes and cautions.




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DAMAGE ASSESSMENT - GENERAL




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WEIGHT VARIANTIn the relevant paragraph 4F (fuselage skin plates Fr 47/54 thru Fr 64between Stgr 32 LH and Stgr 41 RH (welded panel)), a caution noteindicates that the allowable damage effectivity per A/C weight variantmay have to be verified.




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WEIGHT VARIANT (CONT'D)The weight variant is a criterion which is defined for each model ofA/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 variantand, for a same model, the weight variant can change, depending onmodification or SB embodiment status. The current weight variant ofthe affected A/C has to be known before starting the assessment. Ifthe A/C weight variant is not within the table, a damage report has tobe sent to Airbus.Depending on SB/Mod since A/C delivery, only the airline-engineeringdepartment is able to give you the current A/C weight variant. Thecurrent weight variant shall be compared with the data given in a tableat the beginning of allowable damage related paragraph.




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DAMAGE ASSESSMENT - WEIGHT VARIANT (CONT'D)




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DAMAGE CRITERIAA second caution note indicates that in some cases, an inspectionprogram has to be followed.




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DAMAGE CRITERIA (CONT'D)Check the applicability of the allowable damage for dents.




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DAMAGE CRITERIA (CONT'D)Refer to paragraph 4F for dents. Dents are considered as fulfillingnearness/form criterion or out of nearness/form criterion in accordancewith their geometry and their proximity to the nearest adjacent internalstructure elements. This must be determined according to theparameters defined in figure 114 and diagram 102.




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AREA OF RIVETED SUBSTRUCTUREIn areas of riveted substructure, to define whether the dent fulfilsnearness/form criterion, two criteria have to be checked. Refer tofigure 114 sheet 1. B must be at least 15 mm, where B is the smallestdistance measured from the dent edge to any fastener row or anycutout. D must be maximum 10 % of A, where D is the maximumdepth of the dent and A is the smallest distance measured from Dpoint to the closest adjacent structure.




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DAMAGE ASSESSMENT - AREA OF RIVETED SUBSTRUCTURE




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AREA OF RIVETED SUBSTRUCTURE (CONT'D)Check whether the first criterion is fulfilled: B distance is minimum15 mm.The smallest distance measured between the edge of the dent and thesurrounding fastener rows is 95 mm, which is higher than 15 mm.The first criterion is met.




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DAMAGE ASSESSMENT - AREA OF RIVETED SUBSTRUCTURE (CONT'D)




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AREA OF RIVETED SUBSTRUCTURE (CONT'D)The second criterion consists in comparing the maximum depth ofthe dent (D) with the smallest distance measured from the deepestpoint of the dent to the closest adjacent structure (distance A). If noaccess from inside, the measurement is taken from outside. In thiscase, A is X-15 mm, where X is the distance between the deepestpoint and the closest fastener row.




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AREA OF RIVETED SUBSTRUCTURE (CONT'D)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 distanceA. The smallest distance between the deepest point of the dent andthe surrounding fastener rows is 221 mm. Since measured fromoutside, distance A = 221 mm - 15 mm = 206 mm. The secondcriterion is met: D = 3 mm is smaller than 10 % of A.




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AREA OF UNRIVETED SUBSTRUCTUREIn areas of unriveted substructure, to define whether the dent fulfilsthe nearness/form criterion, two criteria have to be checked. Refer tofigure 114 sheet 2. Dent should be out of the welded area. D ismaximum 10 % of A, where D is the maximum depth of the dent andA is the distance measured from D point to the boundary of the weldedarea. Figure 114 sheet 2 informs us to refer to figure 115 for weldedareas.




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DAMAGE ASSESSMENT - AREA OF UNRIVETED SUBSTRUCTURE




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DEFINITION OF AREAS (CONT'D)This figure defines that a welded area is delimited by 25 mm up anddown the stringer.




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DAMAGE ASSESSMENT - DEFINITION OF AREAS (CONT'D)




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AREA OF UNRIVETED SUBSTRUCTURE (CONT'D)The welded area of 50 mm width has been reported on the mapping;the first criterion is fulfilled as the dent is out of the welded area.Check whether the second criteria is fulfilled: D 10 % A. The dentfulfils both nearness form/criterion for unriveted area and riveted area,thus continue the damage assessment.




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DAMAGE ASSESSMENT - AREA OF UNRIVETED SUBSTRUCTURE (CONT'D)




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DAMAGE CRITERIA (CONT'D)As mentioned in a caution at the beginning of the allowable damagepages, the allowable damage applicability have to be checked, usingthe weight variant table (table 107) given at the beginning of theparagraph. The information coming from the airline-engineeringdepartment shows that the MSN 2218 is at weight variant 004. Sincethe weight variant 004 is within table 107, continue the damageassessment.




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DAMAGE CRITERIA (CONT'D)Before starting comparing the dents in accordance with diagram 102as 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 couplingareas.




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ALLOWABLE DENT DIAGRAMThe key to the allowable damage diagram is the skin thickness indented area and the dent depth. You must refer to the data collectedbefore (damage mapping). The diagram is associated to requirementsalready checked at an early stage.




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DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM




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ALLOWABLE DENT DIAGRAM (CONT'D)The skin thickness in the dented area is 1.6 mm (found in theidentification pages). The depth of the dent is 4 mm (measured fromthe A/C-damage mapping). These two values are plotted onto thediagram, which defines a point. The area where this point is locateddefines the subsequent actions to be performed. For the concerneddent read the note: "Check Damage For Cracks By Detailed VisualExamination. If Clear, repair Within 3000 FC". Provided that no crackis detected by detailed visual inspection, the dent is considered as anallowable damage with a time limit (temporary allowable damage).The A/C can be released. But a repair will have to be done before3000 Flight Cycles (FC). If cracked, contact Airbus or repair beforenext flight.




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DAMAGE ASSESSMENT - ALLOWABLE DENT DIAGRAM (CONT'D)




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STRUCTURE PROTECTIONS & AWARENESS D/O (3)

SOURCES OF DAMAGE

Throughout its operational life, the aircraft structure is subjected todifferent types of damage:- fatigue damage (cracking),- accidental damage (e.g. bird impact, ground handling,...),- deteriorations due to environmental and operating conditions (lightningstrike, corrosion, ...).


STRUCTURE PROTECTIONS & AWARENESS D/O (3) May 10, 2006Page 654


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SOURCES OF DAMAGE




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SOURCES OF DAMAGE (continued)

DAMAGE DETECTION/PREVENTIONConcerning fatigue damage, the aircraft is designed and justified, tobe free of significant fatigue cracking during its Design Service Goal(DSG). The scheduled structure inspection programs are prepared todetect any fatigue cracking before it reaches a critical length.Inspections for corrosion are also part of the scheduled maintenanceprogram. Nevertheless, the maximum protection schemes and attentionis paid to protect the aircraft structure against known environmentalaggressions. In addition, the basic protections should be kept in goodconditions and some basic precautions should also be considered.




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SURFACE PROTECTIONS

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




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SURFACE PROTECTIONS




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SURFACE PROTECTIONS (continued)

PROTECTIVE TREATMENT AREAS - FUSELAGE- difficult access, and/or high risk of accidental damage.




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SURFACE PROTECTIONS (continued)

TYPE OF PROTECTIVE TREATMENTS- Type 2 - heavy-duty corrosion preventive compound: grease-likecoatings containing corrosion inhibitors which protect against corrosiveagents.




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SURFACE PROTECTIONS - TYPE OF PROTECTIVE TREATMENTS




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SEALANTS

SEALING IN TYPICAL FUSELAGE AREASIn some specified areas of the aircraft, for example the lower shell, aprotective layer is put on the sealant. This layer makes sure that othermaterials (for example, fuel, hydraulic oil, engine oil and waste fluidsfrom the toilets and galleys) do not cause a deterioration of the sealant.




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SEALANTS (continued)

SEALING IN TYPICAL FUEL TANK AREASIn the fuel tanks, the sealant is used to prevent fuel leaks and corrosionof the fuel tank.




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SEALANTS - SEALING IN TYPICAL FUEL TANK AREAS




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DRAINAGE

- special drain valves installed in those parts of the fuselage and whichare pressurized in flight.The drain holes and drain valves are usually at the lowest part of thefuselage. It is important that any unwanted liquids get to the drain holesor valves. The structure of the lower fuselage is constructed so that a pathis given for these liquids. When you do a repair, make sure that you keepthis path free of unwanted materials.




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DRAINAGE




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COMPOSITE PARTS PROTECTION

COMPOSITE DAMAGESComposite structures can be damaged by lightning strikes or handlingoperations. The environmental conditions may be the source of damagelike rain, dust. The structure can also be affected by impact of foreignobjects or birds for example. At the design stage, the structure has themaximum protection against these different sources of damage.




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COMPOSITE PARTS PROTECTION - COMPOSITE DAMAGES




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COMPOSITE PARTS PROTECTION (continued)

LIGHTNING STRIKE PROTECTION- Zone 3: this zone includes all of the aircraft surfaces that are not inZone 1 and 2. In Zone 3, there is a low probability of attachment ofa lightning strike. However, high lightning currents can go throughZone 3 by direct conduction between two attachment points. Zone 3currents will also go into Zones 1 and 2.




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COMPOSITE PARTS PROTECTION - LIGHTNING STRIKE PROTECTION




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RADOMEThis AMM extract deals with an example of lightning strike protectionin Zone 1, the radome. The radome is a sandwich structure with quartzfiber skins; it is protected using copper straps on the external surface,and bonding braids connecting the aluminum alloy frame to thefuselage structure.




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COMPOSITE PARTS PROTECTION - RADOME




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ELEVATORS AND RUDDERThis second example shows the lightning strike protection of theelevators and rudder trailing edges and tip, which are also located inZone 1. The elevators and the rudder are basically carbon fiberstructures. Their trailing edges are made of an aluminum alloy profileand their tips are also made of aluminum alloy.




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COMPOSITE PARTS PROTECTION - ELEVATORS AND RUDDER




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ELECTRICAL CONTINUITYThe Nose Landing Gear doors are located in Zone 2. Their protectionand the electrical continuity is achieved using a metallic grid installedat the manufacturing stage on the top of the composite layers. Notethat in most cases, this grid should be restored when damaged, as perthe Structural Repair Manual (SRM) procedures.




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COMPOSITE PARTS PROTECTION - ELECTRICAL CONTINUITY




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HANDLING OF COMPOSITE STRUCTURESTo keep composite structures in good and serviceable conditions, theoperator should avoid any damage during handling and/or maintenanceoperations (such as chopped tools, take care of no step areas, ...).Chemical strippers are not authorized on composite structures (theresin system may be deteriorated). The protection like paint schemesand special layers (e.g. tedlar layers on inside surfaces) should be keptin good condition. The drying of composites is also essential beforehot bonding repair operations.




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COMPOSITE PARTS PROTECTION - HANDLING OF COMPOSITE STRUCTURES




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ENVIRONMENTAL & IMPACT PROTECTIONThe impact protection of the Trimmable Horizontal Stabilizer (THS)leading edge is achieved by a metallic cover plate.




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COMPOSITE PARTS PROTECTION - ENVIRONMENTAL & IMPACT PROTECTION




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ENVIRONMENTAL & IMPACT PROTECTION (CONT'D)- titanium fasteners.




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COMPOSITE PARTS PROTECTION - ENVIRONMENTAL & IMPACT PROTECTION (CONT'D)




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DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3)

SESSION OBJECTIVES

SESSION SET-UP

DAMAGE ASSESSMENT PROCEDURE


DETAILED IDENTIFICATION OF THE DAMAGEDPART

ALLOWABLE DAMAGE-GENERAL

DAMAGE CRITERIA

ALLOWABLE DENT DIAGRAM USAGE/FINALDECISION

CONCLUSION

DAMAGE LOCATION


DAMAGE ASSESSMENT EX. 1 OPERATIONAL SCENARIO (3) May 10, 2006Page 686


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SESSION OBJECTIVES ... DAMAGE LOCATION




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MAPPING

DRAFT




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MAPPING - DRAFT




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MAPPING (continued)

FINALIZATION




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MAPPING - FINALIZATION




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ALLOWABLE DENT DIAGRAM




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SESSION OBJECTIVES

SESSION SET-UP


DAMAGE IDENTIFICATION/LOCATION


ALLOWABLE DAMAGE - GENERAL

APPLICABLE ALLOWABLE DAMAGE DIAGRAM

ALLOWABLE SCRATCH DIAGRAM USAGE/FINALDECISION

CONCLUSION

DAMAGE LOCATION




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MAPPING

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MAPPING - DRAFT




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MAPPING (continued)

FINALIZATION




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MAPPING - FINALIZATION




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A318 DAMAGE ASSESSMENT EX. 3 OPERAT. SCENARIO(3)

SESSION OBJECTIVES

SESSION SET-UP


DAMAGE IDENTIFICATION

STRINGERS LOCATION


ALLOWABLE DAMAGE - GENERAL

DAMAGE CRITERIA

CONCLUSION

DAMAGE LOCATION


A318 DAMAGE ASSESSMENT EX. 3 OPERAT. SCENARIO(3) May 10, 2006Page 702


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MAPPING

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MAPPING - DRAFT




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MAPPING (continued)

RESULT




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MAPPING - RESULT




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ata-51 airbus

Documents

window damage assessment

window damage identification

b1b r i t i s h

y s e n g i n e e r

aluminum alloys

titanium alloys

secondary structure

technical publicationsmust