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1996...The Cessna Single Engine Structural Repair Manual

1. Replacement VERTICAL STABILIZER i rA d CAR enc Vertical Stabilizer and Negligible Reparable Danag diga ina g ikgrfercbrfrpvrir Replacement Negligible 2 2 2 Repairable Damage ei RE RERUM EPERERERERRRMENMER Replacement 55 00 00 Page 1 55 00 00 Page 1 55 10 00 Page 801 55 10 00 Page 801 55 10 00 Page 801 55 10 00 Page 801 55 10 00 Page 801 55 20 00 Page 801 55 20 00 Page 801 55 20 00 Page 801 55 20 00 Page 801 55 20 00 Page 801 55 30 00 Page 801 55 30 00 Page 801 55 30 00 Page 801 55 30 00 Page 801 55 30 00 Page 801 55 30 00 Page 801 55 40 00 Page 801 55 40 00 Page 801 55 40 00 Page 801 55 40 00 Page 801 55 40 00 Page 801 CONTENTS Page 1 of 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL HORIZONTAL AND VERTICAL STABILIZERS 1 General A Chapter 55 describes genera
2. 0582T1010 Wing Spar Repair Figure 801 Sheet 3 57 21 00 Page 804 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Filler 2024 0 Alclad 192 XI S Xo 11 i i 4 55 B 3 _ N WI Sg 77 548 4 I edge margin 0 375 inch typical S lt T4 Heat treat to 2024 Filler 2024 T3 Alclad Original parts Repair parts in cross section 2 Repair parts IEEE Filler 2024 T4 Alclad 0 75 inch rivet spacing typical Angle 2024 0 Alclad Heat treat to 2024 4 0582T1011 Jun 1 2005 57 21 00 Page 805 Wing Spar Repair Figure 801 Sheet 4 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL MEASURING WING TWIST INSPECTION CHECK 1 General A This section applies to the procedures required to perform a wing twist check and is applicable to both wing If damage has occurred to a wing it is advisable to check the wing twist washout B Wing twist washout for the Model 172 182 T182 and 206 T206 airplanes is 3 37 2 Tools Equipment and Materials A following equipment is required to accomplish the wing twist check procedure NAME NUMBER MANUFACTURER USE 32 Inch Straightedge Commercial
3. b Ge eee ees 57 20 00 Page 801 e MOL cto o D Pa PUE a ee 57 20 00 Page 801 Wing Rib Damage 57 20 00 Page 801 Wing Rib REPA 23 ie tee Saad 57 20 00 Page 801 WING SPARS x 2525 nsd no tud hoc ege My e rv oda 57 21 00 Page 801 Gere tals M C iMi p Nerio ers tox ed otf e uU 57 21 00 Page 801 Damage 57 21 00 801 Spar hah eL 57 21 00 Page 801 MEASURING WING TWIST 57 22 00 Page 601 General aa a ee etna o HUP APA 57 22 00 Page 601 Tools Equipment and Materials 57 22 00 Page 601 Model 172 Series Wing Twist Check 57 22 00 Page 601 Model 182 Series Wing Twist Check 57 22 00 Page 602 Model 206 T206 Series Wing Twist Check Procedure 57 22 00 Page 602 WING STRINGER 5 57 23 00 Page 801 Stringer Damage 57 23 00 Page 801 Stinget Repair ze Co
4. SULLA 51 40 00 1 eR RENE ERRARE NA SDN PA e T E tet RU ATE 51 40 00 Page 1 teu pete D UU 51 40 00 Page 1 Replacement Of Hi Shear Rivets 51 40 00 Page 1 Substitution Of Rivets Soesi t E vets Pelee han te 51 40 00 1 Rivet cso E 51 40 00 Page 2 Rivet RR UP Eee ee eee 51 40 00 Page 2 Solid Shank 51 40 00 2 Blind ANES Gh Gb ATN EAE EEEE EENEN RIRA RIRI RIIKE 51 40 00 Page 7 Spacing OF Rivets o cocto a a E NR E AA 51 40 00 Page 10 Threaded Fasteners Bolt Torques 51 40 00 Page 10 Rivets for Plastic or Composite Parts 51 40 00 Page 10 FLIGHT CONTROL SURFACE BALANCING 51 60 00 Page 1 Gerieralz o oe verde ec cece pice Dd ra CAIRO Ape T ee 51 60 00 Page 1 Tools and 53 ee Eee baat uum 51 60 00 Page 1 Procedures for Balancing Control 51 60 00 Page 1 Balancing Definitions
5. nee 51 60 00 Page 1 Control Surface Balance Requirements 51 60 00 Page 7 REPAIRS GENERAL 2 5 Prbbesbtebba Tish Gish ese Re exea DE 51 70 00 Page 801 Introductio E atl ated anaes ang duped eed eet aie 51 70 00 Page 801 USAGE iR AS RIRI RIOED 51 70 00 Page 801 Preparation for 51 70 00 Page 801 RIVETED ALUMINUM STRUCTURE 51 71 00 801 Preparing Riveted Aluminum Structure For Repair 51 71 00 Page 801 GLASS FABRIC REPAIR 51 73 00 Page 801 General ARR Aaa t Dette Defect 51 73 00 Page 801 Tools and Maternal ster rece pp er E P P RO 51 73 00 Page 801 Repair Of Glass Fabric 51 73 00 Page 801 REPAIR OF THERMO FORMED THERMO PLASTIC COMPONENTS 51 73 01 Page 801 Thermo formed Thermo Plastic 51 73 01 Page 801 Temporary Repas vori e dope nix A 51 73 01 Page 801 CONTENTS Page 1 of 2 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL TYPIGAL SKIN REPAIRS
6. p d Stop drilled Stop drilled Surface patch Round hole Beveled edge Surface patch Surface patch for crack Beveled edge Surface patch for round holes Avoid sharp corners vi Trim damaged area and round all corners Beveled edge Surface patch for irregular shaped damage Typical Windshield and Windows Repair Figure 801 Sheet 1 56 10 00 Page 802 Cessna Aircraft Company Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 184 Correct gt gt ELS gt 4 E n Incorrect A lt lt Sanding repair Patch should be thicker Patches c lt Patch tapered on sharper angle than material Patch and hole should be trimmed with tapered edges During cementing pressure Heat edges of patch need be applied only on until soft and force top surface Taper assures it into hole Hold it in equal pressure on all sides place until cool and hard to assure perfect fit Then remove patch for cementing bath 22 After cement has hardened sand or file edges level with surfaces 2682T1043 Typical Windshield and Windows Repair Figure 801 Sheet 2 56 10 00 Page 803 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 2 Make plug patch slightly thicker than the material being repaired and similarly bevel the edges 3 Install plug
7. 7 Hinge point 57 10 Aileron aqui ie NOTE Level surface ws inches DETAIL D 100 50 NOTE NOTE Alternate method before making trailing edge measurement make sure trailing edge of aileron is Do maU RECON 58 5 SS inches straight in this area Piano hinge Balancing mandrel 0582T1003 Balancing Control Surfaces Figure 2 Sheet 4 51 60 00 Page 6 Cessna Aircraft Company Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B159 Hinge line Chord line Control Surface Overbalance Figure 3 B Underbalance refer to Figure 4 is defined as the condition that exists when surface is trailing edge heavy and is defined by symbol If the balance beam uses a sliding weight the weight must be on the leading edge side of the hinge line to balance the control surface is considered to be under balanced M Hinge line Chord line Control Surface Underbalance Figure 4 5 Control Surface Balance Requirements NOTE Approved Flight must never be exceeded when the surface is in its final configuration for flight A Referto Tables 1 2 and 3 for balance limits of the various airplane control surfaces These approved flight limits must take into account all items which may be attached and or applied to the various control surfaces static wicks trim tabs paint decorative trim stripes and so forth Table 1 Model 172 Static Balance Limits CONTROL SURFAC
8. Usage A Typical repairs may be accomplished individually or combined with other repairs for a major repair Technique and material variation is permissible only so far as to facilitate fabrication and ensure the original strength and usefulness of the affected component Preparation for Repair A The airplane should be located in an area where once positioned minimum movement or relocation is required The airplane should be leveled and supported as necessary Refer to appropriate Maintenance Manual Chapter 7 Jacking Maintenance Practices and Chapter 8 Leveling Maintenance Practices 51 70 00 Page 801 Cessna Aircraft Company Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL RIVETED ALUMINUM STRUCTURE REPAIR 1 Preparing Riveted Aluminum Structure For Repair A To prepare an area for repair examine and classify the damage Make a thorough check before beginning repairs In some cases a damaged part may be classified as needing replacement however after removal closer inspection indicates the part may be repaired Remove all ragged edges dents tears cracks punctures and similar damages Stop drill all cracks using a No 30 0 128 inch drill Leave edges after removal of damaged area parallel to any square or rectangular edges of the unit Round all corners Smooth out abrasions and dents Deburr all edges of repair and ensure that no nicks or scratches remain Brush a
9. CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL to an 8 by 11 inch loose leaf page and represents a frame Horizontal rows are lettered from A at the top to L at the bottom Vertical columns are numbered 1 to 24 The combination of a letter and a number identifies a frame image in the aerofiche card The List of Chapters provides a quick reference to information contained in the aerofiche 5 Using the Structural Repair Manual or Aerofiche A Division of Subject Matter 1 Structural repair information is divided into chapters in accordance with Air Transport Association Specification 100 Each Chapter is further subdivided to provide individual or related structural member presentation 2 Chapter 51 provides general structural information required to perform a repair Also included in Chapter 51 are general repair procedures that may be accomplished in noncritical areas Effectivity Page 1 Alistof effective pages is provided with each chapter All pages listed are active and shall appear in sequence as recorded in the Effectivity Page 2 The Effectivity Page contains tabular listing of ATA number page and date of each page in that chapter A change in the chapter requires a revision to the chapter s Effectivity Page The date corresponds to the date that appears on the individual page which defines when that page was issued Page Numbering System 1 The Structural Repair Manual or corresponding aerofich
10. Channel Repair Figure 801 for repair illustrations 57 23 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL AUXILIARY SPAR REPAIRS General A auxiliary spar is constructed of formed sheet metal and is behind the trailing edge ribs from approximately WS 100 50 to 208 00 The auxiliary spar is attached to upper skins lower skins and other wing structure using rivets Auxiliary Spar Damage A Damage to the auxiliary spar can be divided into three major categories and is detailed in Wing Damage Classification Auxiliary Spar Repair A Repairs to the auxiliary spar are illustrated in Figure 801 57 24 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B195 Doubler 2024 T4 Alclad Clean out damaged Filler 2024 4 Alag 0 25 inch edge margin 20 rivets each side of damaged area MS20470AD4 rivets Original parts Repair parts Repair parts in cross section 0582T1012 Auxiliary Spar Repair Figure 801 Sheet 1 57 24 00 Page 802 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LEADING EDGE REPAIRS Leading Edge Damage Classification A Damage to the leading edge can be divided into three major categories
11. Remove the minimum material required to clean up the damage b Make sure the diameter of the axle attachment holes is 0 383 inches maximum for 3 8 inch bolts c Make sure the diameter of the axle attachment holes is 0 321 inches maximum for 5 16 inch bolts d If reaming to the maximum dimension does not remove all signs of corrosion discard the landing gear spring 51 11 00 Pages Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL REPAIR MATERIALS General A This section provides information covering the materials used for repairs Repair Materials A In general materials used in the airplane include 2024 and 7075 aluminum alloys Sheet material requiring little or no forming will generally be of 2024 T3 clad aluminum Formed parts such as ribs bulkheads etc will be of 2024 T42 clad aluminum Forgings are of 7075 173 Materials used in repairs should be where possible of the same material and heat treated to the same temper The thickness should be equal to or greater than the material being repaired unless otherwise noted If the type of material cannot be readily determined and the forming required is not severe 2024 T3 may be used generally since the strength of T3 is greater than that of T4 or T42 T4 and T42 may be used interchangeably but they may not be substituted for T3 When it is necessary to form a part with a smaller bend radius than the
12. ated oed ie 51 75 00 Page 801 Cn cM 51 75 00 801 Guidelines for Corrugated Skin Crack Repairs 51 75 00 Page 801 CONTROL SURFACE 51 76 00 Page 801 Gerleral zs deoa ha ale hale ater sa 51 76 00 801 5 Page 2 of 2 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL STANDARD PRACTICES AND STRUCTURES GENERAL General A Chapter 51 describes general repair practices materials and procedures which are applicable throughout the subsequent chapters This chapter also provides general information for performing any structural repairs Unless otherwise specified all dimensions are in inches forces are in pounds and torques are in inch pounds The airplanes are of an all metal semimonocoque construction with the skin carrying a portion of all structural loads To obtain information covering dimensions areas and stations diagrams refer to current appropriate Model 172 Model 182 or Model 206 Maintenance Manual Chapter 6 Dimensions and Areas For information covering leveling and weighing refer to current appropriate Model 172 Model 182 or Model 206 Maintenance Manual Chapter 8 Leveling and Weighing Description A B The fuselage is of conventional semimonocoque construction
13. iL a 9 0 P 9 a E 9 8 8 8 aes B 3 882 CN T CN oo oS lt Figure 801 Sheet 1 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL NN NS 2959 c Na uM Jy S o Y 5 e 8 a ANUN 5 x Cy 5 sE y UUET o8 8 9 g 7 2h NS 5 E 55 5 0 S338 V Btg lt lt 25 B T sy 7 159 2772 5 5 og Wwy 9 gt 9 EN 7 2 RRC WWW 35 s ai 1 4 aou NN N UU Ay 2 508 As 8 55o5 tes STB9 di AU WD 3 gg 9 9 5 e NW ad 4558 9 4 28 5 8 57 o Eo 5 x 1 4555 5 886 d cog N o 0 5 5 E o Sr Noo 80 2 xo N CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 191 0 75 inch rivet spacing Filler 2024 T4 Alclad Clean out damaged area 0 25 inch edge margin Spar Ms y VIEW Angle 2024 T4 Alclad MS20470AD4 A rivets N Original parts Repair parts E Repair parts in cross section
14. over the higher strength base alloy surface Since pure aluminum has relatively greater corrosion resistance than the stronger alloy it is imperative the clad surface be maintained intact to the maximum extent possible and to avoid unnecessary mechanical removal of the protective coating In addition aluminum parts receive a chemical conversion coating and are then epoxy primed Clean area to be reworked Strip paint as required To determine the extent of corrosion damage refer to Corrosion Damage Limits Remove light corrosion by light hand sanding Mechanically remove moderate or severe corrosion by hand scraping with a carbide tipped scraper or fine fluted rotary file Remove residual corrosion by hand sanding Select appropriate abrasive from Figure 2 Blend into surrounding surface any depressions resulting from rework and surface finish with 400 grit abrasive paper Clean reworked area Determine depth of faired depressions to ensure that rework limits have not been exceeded Chemically conversion coat rework area Restore original finish epoxy prime IAIWIN 10 o Co N gre Steel a Unlike some other metal oxides the red oxide of steel rust will not protect the underlying base metal The presence of rust actually promotes additional attack by attracting moisture from the air and acting as a catalyst in causing additional corrosion to take place Light red rust on bolt heads hold down nuts and other
15. the flap actuator in the left hand wing electrical wiring and wiring disconnect points the wing portion of the ventilation system strut attach fittings and the inside of the fuel cell B Refer to applicable Maintenance Manual Chapter 6 Dimensions and Areas for wing station diagrams C If questions arise concerning approved repairs or for repairs not shown in this section contact Cessna Propeller Aircraft Product Support Box 7706 Wichita KS 67277 316 517 5800 Facsimile 316 942 9006 Tools Equipment and Materials A Referto Figure 1 foran illustration of wing and fuselage support stands which may by fabricated locally and used during structural repair Installation of Access Holes NOTE In some instances it may be advantageous to create access holes in the wing skin to facilitate wing repair Refer to the following steps and Figure 2 for an illustration of access holes WARNING The following procedures are not applicable to the integral fuel cell skins A Precautions and Notes Add the minimum number of access holes necessary Any circular or rectangular access hole which is used with approved optional equipment installations may be added in lieu of the access hole illustrated Do not add access holes at outboard end of wing remove wing tip instead Locate new access holes near the center of a bay spanwise Locate new access holes forward of the front spars as close to the front spar as practical
16. 0 50 2 222222 NX UK gt CESSNA AIRCRAFT COMPANY SINGLE ENGINE 9 a 5 9 54 gt a gt STRUCTURAL REPAIR MANUAL 5 51 76 00 Page 802 Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 169 Stop drill hole diameter 0 094 inch minimum Crack or hole Mechanically expanded rivet Original parts Z Repair parts Repair parts cross section Typical Control Surface Trailing Edge Repair Figure 802 Sheet 1 51 76 00 Cessna Aircraft Company 2682T1026 Page 803 Jun 1 2005 DOORS CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT PAGE DATE 52 Title 52 List of Effective Pages 52 Record of Temporary Revisions 52 Table of Contents 52 00 00 Page 1 Jun 1 2005 52 10 00 Page 101 Jun 1 2005 52 LIST OF EFFECTIVE PAGES Page 1 of 1 Cessna Aircraft Company Jun 1 2005 RECORD TEMPORARY REVISIONS Temporary Revision Number Page Number Issue Date By Date Removed CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CONTE
17. 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL DAMAGE INVESTIGATION AND CLASSIFICATION 1 General A C For the purposes of this manual damage is considered to be a deviation from the original configuration of a structural part that compromises its structural integrity by significantly reducing its strength significantly decreasing its resistance to fatigue significantly increasing its susceptibility to corrosion significantly altering its flutter characteristics or adversely affecting the flight characteristics of the airplane This can include but is not limited to scratches dents dings gouges cracks drill starts double drilled holes plastic deformation reduction in cross sectional areas changes in component center of gravity missing or inadequate fasteners corrosion dissimilar metal contact work hardening temper change due to excessive heat and so forth Use good judgment in determining the type of significant change to flat stock structural material The terms dent crease abrasion gouge nick scratch crack and corrosion referred to elsewhere in the manual are defined below as a guide for this determination particularly with respect to the external skin of the airplane 1 Dent A dentis normally a damaged area which is depressed with respect to its normal contour There is no cross sectional area change in the material Area boundaries are smooth Its form is generally the
18. 26 00 Page 801 57 27 00 Page 801 57 27 00 Page 801 57 27 00 Page 801 57 40 00 Page 801 57 40 00 Page 801 57 40 00 Page 801 57 40 00 Page 801 57 50 00 Page 801 57 50 00 Page 801 57 50 00 Page 801 CONTENTS Page 2 of 2 Jun 1 2005 1 General CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL WINGS GENERAL A Description of Wing Assemblies 1 2 3 4 5 The wing assemblies are a semicantilever type employing semimonocoque type of structure The internal structure consists of a built up front spar a formed inboard front fuel spar a rear spar and a formed auxiliary spar assembly in the aileron attach area Ribs are formed sheet metal and consist of nose intermediate and trailing edge assemblies On the 172 series airplanes stressed skin is riveted to the rib and spar assemblies to complete the rigid structure On 182 and 206 series airplanes the skin is bonded to the leading edge ribs and riveted at other locations The inboard section of the wing is sealed to form an integral fuel cell The sealed area runs from the wing root outboard toward the strut attach and from the front fuel spar to the rear spar NOTE On the 172 series airplanes the fuel closeout rib is located approximately 7 inches outboard from the wing root Access openings hand holes with removable cover plates are located in the wing These openings afford access to flap and aileron bellcranks and control systems
19. 50 0 55 0 094 0 125 0 125 0 125 0 125 0 156 0 156 N es gt pus 5 C 3o Ten 5 o Qo Co o Be cH 9 no 5 LL 3 0 o bm ors C 55 Q 2 0 016 0 020 0 025 0 032 0 040 0 051 0 063 268211029 802 1 2005 51 75 00 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Rivet diameter Hole size D B B163 Doubler Carry existing rivet pattern through patch Section through assembled patch Original parts in cross section Ba Repair parts EXE Repair parts Existing skin T42 Doubler 2024 Edge margin is equal to 2 times rivet diameter same gage as skin T42 same gage as skin Patch 2024 RIVET TABLE 1 16 NOTE 0 82 NOTE 0 94 NOTE RIVET SPACING 0 84 NOTE 0 96 NOTE 0 90 NOTE 0 70 0 96 NOTE 0 58 0 94 NOTE 0 50 0 55 0 094 0 125 0 125 0 125 0 125 0 156 GAGE DIAMETER MS20426AD MS20470AD 0 156 SKIN RIVET 0 016 0 020 0 025 0 032 0 040 0 051 0 063 vicinity of flush and protruding head rivets Patch 2682T1030 Skin Repair Figure 801 Sheet 2 NOTE Spacing is for a double row of rivets Jun 1 2005 51 75 00 Page 803 Cessna Aircraft Company
20. Construction consists of formed bulkheads longitudinal stringers reinforcing channels and skin panels The wings are of an all metal strut braced semimonocoque construction utilizing two spars Each wing consists of a wing panel with an integral fuel bay an aileron and a flap The empennage group is of a fully cantilevered design and consists of a conventional rudder and elevator configuration The horizontal stabilizer is of one piece construction consisting of spars ribs and skins Elevators are constructed of spars ribs and skin panels The skin panels are riveted to the ribs and spars A balance weight is located in the outboard end of each elevator forward of the hinge line An elevator trim tab is attached to the right hand elevator and is constructed of a spar ribs and skin riveted together The vertical stabilizer is constructed of a forward and aft spar ribs and skin The rudder is constructed of spars ribs and skin panels The main landing gear consists of 6150M alloy spring steel cantilevered with attaching parts of high strength 7075 T73 aluminum alloy forgings Nose gear components are 4130 alloy steel and 7075 T73 aluminum alloy forgings The engine mount is constructed of welded 4130 steel tubing on the 172 and 182 The 206 has a built up aluminum sheet metal engine mount The removable engine cowling is made of 2024 Alclad secured with quarter turn fasteners 51 00 00 Page 1 Cessna Aircraft Company Jun
21. Degrees bo Spirit level protractor Sliding Trailing edge weight e support Qu line N Balancing Elevator mandrel Leveled surface 2682T1023 Balancing Control Surfaces Figure 2 Sheet 2 51 60 00 4 Cessna Aircraft Company Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B157 Ailerons Hinge line 2 Horizontal plane 0 850 inches balance aileron inverted with trailing edge at point opposite cutout for middle hinge 0 850 inches below hinge line horizontal plane viEW A A 2682T1024 Balancing Control Surfaces Figure 2 Sheet 3 51 60 00 Page 5 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 158 6 0 mandrel 1 16 inch slot C inches 3 4 inch deep to fit Per d aileron hinge Knife edge 22 ES DETAIL B inch amp 18 0 inches 5 4 0 P inches After locating trailing edge support balance DETAIL A 6 5 by adding washers and inches or nuts DETAIL C TUN Spirit level Trailing edge i Balancing g p protractor support mandrel Sliding weight Chord line A at aileron midspan WS 154 00 e
22. Distribution contains Type Class B 1 2 and Type VIII Class B 12 access sealants with the proper quantity of accelerator for each sealant WARNING The accelerators contain heavy metal peroxides Keep them away from heat and flame Use only in well ventilated areas Avoid skin and eye contact Wear eye shields In case of eye contact flush generously with water and get prompt medical attention Mixing Sealant A Use all the accelerator and sealant in the container when mixing to ensure the proper ratio of accelerator to sealant Stir the accelerator to absorb all floating liquid before it is mixed with the sealant The accelerator can then be poured into the container of sealant for mixing otherwise a wax free container must be used Stir accelerator and sealant until they become a uniform mixture Do not stir air into mixture so it forms bubbles if bubbles appear they must be removed CAUTION Protect drain holes and fuel outlet screens when applying sealants NOTE Work life of sealants contained SK210 56 is 2 hours from the start of mixing Work life of sealants contained in SK210 101 is one half hour from the start of mixing This is based on a standard condition of 77 F 256 and 50 percent relative humidi ty An increase in either temperature or humidity will shorten the work life of the sealants Applying Sealant A Use the following procedures as the best method for applying sealant 1 Apply fay surface sealant t
23. Exercise care while accomplishing this operation to prevent unseating the rivet by too much pressure For the first few blows the bucking bar should be held lightly against the rivet shank so it will receive the impact of the blow through the rivet The bucking bar must be held square with the rivet to produce uniform upsets As few blows as possible should be struck to properly upset rivet Blows must be as uniform as possible Loose Or Working Solid Shank Rivets 1 Rivets which appear to be loose shall be checked with a 0 002 inch feeler gauge by inserting the gauge around the head of the rivet in question If the feeler gauge can be inserted to the shank of the rivet it shall be classified as a loose rivet and it shall be replaced If the feeler gauge can be inserted approximately halfway to the shank for less than 30 percent of the circumference of the rivet head it shall not be classified as a loose rivet The feeler gauge shall be used to check the shear section between the riveted members such as skin to spar or different sections of skins in a similar manner to that used around the rivet head If the skin around the brazier head or countersunk rivet can be moved by depressing the skin with finger pressure around the rivet the rivet shall be replaced If a rivet is found which turns by applying a rotating load to the head of the rivet it should be replaced In areas where exterior paint has been applied to rivet heads the paint may h
24. Notes 1 2 5 Refer to Notes 1 Refer to Notes 1 5 6 Refer to Notes 1 6 Refer to Notes 1 6 6 Refer to Notes 1 Refer to Notes 1 6 Refer to Notes 1 6 Refer to Notes 1 4 Refer to Notes 1 3 4 Refer to Notes 1 2 5 Cessna Aircraft Company Fastener WITH HL18 NAS1054 NAS14XX NAS529 NAS1146 NAS7034 NAS464 NAS1103 NAS1116 NAS1303 NAS1316 NAS6203 NAS6216 NAS6603 NAS6616 AN173 NAS14XX 5529 51446 Collar HL70 HL82 NAS179 NAS528 NAS1080C NAS1080E NAS1080G NAS1080AG NAS528 NAS179 NAS1080C NAS1080E NAS1080G NAS1080AG NAS1080K MIL S 7742 MIL S 7742 MIL S 7742 MIL S 7742 MIL S 7742 AN305 MS20305 MS21044 MS21045 NAS1080C NAS1080E NAS1080G NAS1080AG NAS528 NAS179 NAS1080C NAS1080E NAS1080G NAS1080AG 51 40 00 Page 4 Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Table 2 Approved Fastener Substitutions continued Fastener Collar DIAMETER Fastener Collar REPLACE WITH Refer to Notes 1 57034 NAS1080K 5 Refer to Notes 1 NAS464 Refer to Note 8 6 Refer to Notes 1 NAS1103 Refer to Note 8 6 NAS1106 Refer to Notes 1 NAS1303 Refer to Note 8 6 NAS1306 Refer to Notes 1 NAS6203 Refer to Note 8 6 NAS6206 Refer to Notes 1 NAS6603 Refer to Note 8 6 NAS6606 NOTE 1 Refer to appropriate tables for nominal diameters av
25. Page 1 Cessna Aircraft Company Jun 1 2005 170 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL FS FS 142 00 178 00 Es rs S 12400 8 12 44 00 FS FS 108 00 56 70 MODEL 172R 172S FUSELAGE FS FS FS FS FS 8 12 56 70 92 00 140 00 172 00 FS FS FS 79 00 110 00 45 00 MODEL 182S 182T T182T FUSELAGE Fuselage Stations Figure 1 Sheet 1 Cessna Aircraft Company 05101008 0710T1008 53 00 00 Page 2 Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 171 FS FS FS FS FS FS 18 40 65 30 112 00 152 20 180 60 208 00 B FS FS 138 00 FS FS FS 100 00 FS FS 494 80 FS 8 10 59 70 Di 166 40 230 10 FS FS 68 30 0 00 FS MODEL 206H T206H FUSELAGE FS FS 54 80 rs 9000 5 Fs 112 00 65 30 39 30 Aee FS FS 124 00 FS 68 30 FS 100 00 18 40 FS 59 70 CARGO DOORS RIGHT SIDE ONLY MODEL 206H T206H 121071007 121071007 Fuselage Stations Figure 1 Sheet 2 53 00 00 Page 3 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL FUSELAGE DAMAGE CLASSIFICATION General A Damage to the fuselage can be divided into three major categories negligible damage repairable damage and major replacement damage The categories are provided to assist in determining the e
26. Patch and doubler 2024 T4 same gage as skin CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL unless otherwise noted NOTE All dimensions are in inches WwW MS20470AD4 B164 rivets Patch repair for 3 00 diameter hole 24 required Existing skin UU N amp Doubler 7 50 diameter Patch 2 Zz 4 S Section through assembled patch 1 lj d WS diameter 3 00 hole Patch repair for 2 00 diameter hole Existing skin Doubler E 4 00 diameter 3 00 diameter 5 00 diameter Section through assembled patch E V NS N MS20470AD4 rivets 16 required N NN i 42 N Patch repair for 1 00 diameter hole 2 00 gt x 1 75 diameter Doubler 2 diameter hole Na Filler K 2 50 diameter MS20470AD4 rivets 9 gt G 5 c 5 o Original parts Repair parts Repair parts in cross section diameter 1 00 hole 8 required 268211025 Overlapping circular patch Skin Repair Figure 801 Sheet 3 Jun 1 2005 51 75 00 Page 804 Cessna Aircraft Company 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Doubler 2024 T42 same gage as skin Repair parts 2 Patch 2024 42 R
27. REPAIR MANUAL CONTENTS WINDOWS GENERALE 5 che bur wenn uae 56 00 00 Page 1 Gerieral uico RR Oe bte C Wik nade KORR n RA 56 00 00 Page 1 PLASTIC WINDOW SURFACE REPAIR 56 10 00 Page 801 Repair of Plastic Window Surfaces 56 10 00 Page 801 Tools and Materials ect cnp pb hbbb etse PEE pP eee eet ee ENE EEREN 56 10 00 Page 801 Stop MINING eso ehem mcr ADU RET Ae nea dees San Deak 56 10 00 Page 801 Surface ua e n aa ebb has 56 10 00 Page 801 Insert Plug Patch ec care sees esate dd ava er RO eee AER AE oa ee 56 10 00 Page 801 Minor SCralcli8s sedo Re A ee ee seeds 56 10 00 Page 804 Ol aning Plastic cono ono E 56 10 00 804 CONTENTS Page 1 of 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL WINDOWS GENERAL 1 General A This chapter provides repair information applicable to windshields and windows used on the 1996 and On single engine airplanes These repairs may be utilized without removing components from the airplane For windshield window removal or replacement refer to the various model Maintenance Manuals Chapter 56 Windows 56 00 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STR
28. Repair After Hard 0 STRINGER AND CHANNEL General FIREWALL REPAIR General Material Repairing the Firewall 1 CONTENTS Cessna Aircraft Company 53 00 00 Page 1 53 00 00 Page 1 53 00 00 Page 1 53 10 00 Page 1 53 10 00 Page 1 53 10 00 Page 1 53 10 00 Page 1 53 10 00 Page 2 53 20 00 Page 801 53 20 00 Page 801 53 20 00 Page 801 53 20 00 Page 801 53 20 00 Page 801 53 20 00 Page 801 53 30 00 Page 801 53 30 00 Page 801 53 40 00 Page 801 53 40 00 Page 801 53 40 00 Page 801 53 40 00 Page 801 Page 1 of 1 Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL FUSELAGE GENERAL 1 General A B Chapter 53 describes general repair practices materials and procedures which are applicable to the Fuselage and Fuselage Structure Refer to Figure 1 for illustrations of fuselage stations For repairs beyond the scope of this chapter refer to Chapter 51 Typical Skin Repairs 2 Fuselage A B The fuselage is of semimonocoque construction and consists of formed bulkheads longitudinal stringers reinforcing channels and skin panels If questions arise concerning approved repairs or for repairs not shown in this section contact Cessna Propeller Aircraft Product Support 53 00 00
29. black or dark gray stain is found adjacent to or around the fastener head Generally it takes the form of a dirt or oily streak aft of the loose rivet d Mark ared line across the fastener head and the adjacent material Check the line at the next inspection Any loosening of the fastener will break the line as indicated in Figure 3 9 Spacing Of Rivets A There are no specific rules which are applicable to every case or type of riveting There are however certain general rules which should be understood and followed Edge distance of rivets should not be less than two diameters of the rivet measured from the edge of the sheet or plate to the center of the rivet hole Spacing between rivets when in rows depends upon several factors principally the thickness of the sheet the diameter of the rivets and the manner in which the sheet will be stressed This spacing is seldom less than four diameters of the rivet measured between the centers of the rivet holes Rivets spaced four diameters apart are found in certain seams of semimonocoque fuselages webs or built up spars and various plates or fittings Where there are two rows of rivets they are usually staggered The transverse pitch or distance between rows should be slightly less than the pitch of the rivets with 75 percent of the rivet pitch being the usual practice An average spacing or pitch of rivets in the cover or skin of most structures except at highly stressed points will be
30. drill slipping off and tracking across the metal While holding the drill at a 90 angle drill the rivet to the depth of its head Be careful not to drill too deep because the rivet 51 40 00 Page 2 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL shank will turn with the drill and cause a tear The rivet head will often break away and climb the drill which is a good signal to withdraw the drill If the rivet head does not come lose of its own accord insert a drift punch into the hole and twist slightly to either side until the head comes off 4 Drive out the shank of the rivet with a drift punch slightly smaller than the diameter of the shank On thin metal or unsupported structures support the sheet with a bucking bar while driving out the shank If the shank is exceptionally tight after the rivet head is removed drill the rivet about two thirds of the way through the thickness of the material and then drive out the remainder of the rivet with a drift punch b The removal of flush rivets is the same as that just described except that no filing of the manufactured head is required before center punching Be very careful to avoid elongation of the dimpled or the countersunk holes The rivet head should be drilled to approximately one half the thickness of the top sheet Table 1 Approved Replacement Fasteners Chart REPLACE Inch thickness or thicker WITH MS20470AD3 0 025 NAS139
31. for opposite wing WARNING Purge fuel bays with an inert gas argon or carbon dioxide prior to repairing fuel leaks to preclude possibility of explosions Insert inert gas supply hose into fuel filler opening Allow gas to flow into bay for several minutes to remove all fuel vapors Since argon or carbon dioxide are heavier than air these gasses will remain in bay during repair Non sparking tools shall be used to make repairs air motors plastic scrapers etc NOTE Portable vapor detectors are available to determine presence of explosive mixtures and are calibrated for leaded fuel The detectors can be used to determine when it is safe to make repairs NOTE During structural repair parts must be predrilled countersunk or dimpled and cleaned before being sealed and positioned for final installation Remove all existing sealant from area to be sealed leaving a taper to the remaining sealant The taper will allow a scarf bond and a continuous seal when the new sealant is applied NOTE The best method of removing sealant is with a chisel like tool made of hard fiber or plexiglass Remaining sealant can be removed with aluminum wool Steel wool or sandpaper must not be used 57 11 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Stop drill cracks using a No 30 0 128 inch drill Remove all ragged edges dents tears cracks and punctures After remova
32. from 6 to 12 diameters of the rivet The best practice in repair is to make pitch of rivets equal to those in the original structure 10 Threaded Fasteners Bolt Torques A The importance of correct application cannot be overemphasized Refer to appropriate Maintenance Manual Chapter 20 Torque Data Maintenance Practices for additional information covering torque values Under torque can result in unnecessary wear of nuts and bolts as well as parts they are holding together When insufficient pressures are applied uneven loads will be transmitted throughout assembly which may result in excessive wear or premature failure due to fatigue Over torque can be equally damaging because of failure of a bolt or nut from overstressing threaded areas There are a few simple but very important procedures that should be followed to assure that correct torque is applied 1 Calibrate torque wrench periodically to assure accuracy and recheck frequently 2 Be sure that bolt and nut threads are clean and dry unless otherwise specified 3 Run nut down to near contact with washer or bearing surface and check friction drag torque required to turn nut 4 Add friction drag torque to desired torque recommended Refer to appropriate Maintenance Manual Chapter 20 Torque Data Maintenance Practices to obtain complete torque calculating procedures This is referred to as final torque which should register on indicator or setting for a snap over ty
33. hae 71 20 00 Page 801 dus 71 20 00 Page 801 CONTENTS Page 1 of 1 Jun 1 2005 1 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL POWERPLANT GENERAL General A Single engine airplanes produced from 1996 and On use Lycoming powerplants These powerplants are attached to the fuselage by dynafocal mounts 172R 1725 1825 182 and T182T or by sheet metal bed mounts 206H and T206H This chapter covers structural repair to the cowlings 172R 1725 1825 182 and 182 and structural repair to the welded engine mounts 172R 1725 1825 182 182 For repair information not covered in this manual contact Cessna Propeller Aircraft Product Support P O Box 7706 Wichita KS 67277 Telephone 316 517 5800 or Facsimile 316 942 9006 71 00 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL ENGINE COWLING REPAIRS General A This section provides repair procedures for the cowl skins and reinforcement angles Repair of Cowling Skins A Cowl halves are made of formed aluminum skin If extensively damaged complete sections of cowling must be replaced Standard insert type skin patches however may be used if repair parts are formed to fit Small cracks may be stop drilled and dents straightened if they are reinforced on the inner side with a doubler of the same material Repair of Reinforcement Angles A Due to t
34. no evidence of skin tears cracks or skin penetrations which are not stress wrinkles and do not interfere with internal structure of mechanism constitute negligible damage and rework is considered cosmetic Repairable damage Dents or dings deeper and or larger than specified above must be repaired Skin tears cracks or penetrations must be repaired Dings that include understructure ribs frames and spars must be repaired by reforming or removal and replacement of the damaged member or damaged are Reevaluation of the skin after repair of the understructure will determine if the skin damage is negligible repairable or requires replacement Damage Necessitating Replacement Of Parts If a skin is badly damaged repair must be made by replacing an entire skin panel from one structural member to the next Repair seams must be made to lie along structural members and each seam must be made exactly the same in regard to rivet size spacing and pattern as the manufactured seams at the edges of the original sheet If the manufactured seams are different the stronger must be copied If the repair ends at a structural member where no seam is used enough repair panel must be used to allow an extra row of staggered rivets with sufficient edge margin to be installed Wing Stringer Damage Criteria A B Negligible Damage Minor Scratches or abrasions are the only form of damage considered negligible to wing stringers Repairable damage Den
35. nonstructural hardware is generally not dangerous However it is indicative of a general lack of maintenance and possible attack in more critical areas such as highly stressed steel landing gear components and flight control surface actuating components When paint failures occur or mechanical damage exposes highly stressed steel surfaces to the atmosphere even small amounts of rusting are potentially dangerous and must be removed The most practical means of controlling corrosion of steel is the complete removal of the corrosion products by mechanical means Except on highly stressed steel surfaces the use of abrasive papers small power buffers and buffing compounds and wire brushes are acceptable for clean up procedures However residual rust usually remains in the bottom of small pits and crevices Clean area to be reworked Strip paint as required Remove all degrees of corrosion from steel parts using a stainless steel hand brush or hand operated power tool Alternatively use dry abrasive blasting process WIN 51 11 00 Page5 Cessna Aircraft Company Jun 1 2005 150 CESSNA AIRCRAFT COMPANY STRUCTURAL REPAIR MANUAL SINGLE ENGINE Metals Restrictions Operation Abrasive Paper or Cloth Abrasive Alum Stain Abrasive Matenalg Alum Silicon Garnet Fabric finum less to be inum Carbide or Pad Steel Y
36. patch as illustrated in Figure 801 4 Heat plug patch until it is soft press into the hole without plastic solvent adhesive and allow to cool to make a perfect fit 5 Remove plug patch coat surfaces to be bonded with plastic solvent acrylic chips dissolved in methylene chloride and insert plug patch in the hole 6 Maintain a firm light pressure until the plastic solvent adhesive has set 7 Sand or file edges level with surface buff and polish Do not attempt hand polishing until surface is clean A soft open type cotton wheel is suggested NOTE Acrylic and cellulose plastics are thermoplastic Friction created by buffing or polishing for too long a time in one spot can generate sufficient heat to soften the surface This will produce visual distortion and is to be guarded against Minor Scratches A following procedure should be used when repairing minor scratches 1 Remove minor scratches by vigorously rubbing the affected area by hand using a soft clean cloth dampened with Novus 2 plastic polish and finish by polishing with Novus 1 Remove polish with a soft dry cloth NOTE Plastics should not be rubbed with a dry cloth since this is likely to cause scratches and also builds up an electrostatic charge which attracts dust particles to the surface If after removing dirt and grease no great amount of scratching is visible finish the plastic with a good grade of commercial wax Apply the wax in a thin even co
37. sealed on the fuel side Fay surface sealing is applying sealant to one mating part before assembly Enough sealant must be applied so it will squeeze out completely around joint when the parts are fastened together The fillet seal is applied after the joint is fay surface sealed and fastened Sealer is fillet applied to the edge of all riveted joints joggles bend reliefs voids rivets or fasteners All boundaries and any other place that could become a fuel leak are sealed The fay sealant need not be cured before applying the fillet sealer however the fay sealant must be free of dirt or other contaminants before applying fillet seal Fillets laid on intersecting joints shall be joined together to produce a continuous seal Sealant must be pressed into the joint to displace any entrapped air bubbles Use an extrusion gun to lay a bead along joint and work out all entrapped air with a small paddle to eliminate bubbles Integral Fuel Bay Sealant A Two types of sealants are used one to seal the bay and the other to seal access doors fuel quantity transmitters fuel inlet assemblies and fuel test receptacle The access door sealant is more pliable and will not adhere to metal as firmly as the bay sealant This permits access doors fuel quantity transmitter etc to be removed without damage Service Kit SK210 56 available from Cessna Parts Distribution contains Type Class B 2 and Type VIII Class B 2 access sealants with Cessna Parts
38. skin holes because of a permanent set in the stringer If this is apparent replacement of the stringer will usually restore the original strength characteristics of the area NOTE Wrinkles occurring in the skin of the main landing gear bulkhead areas must not be considered negligible The skin panel must be opened sufficiently to permit a thorough examination of the lower portion of the landing gear bulkhead and its tie in structure Wrinkles occurring in open areas which disappear when the rivets at the edge of the sheet are removed or a wrinkle which is hand removable may often be repaired by a 1 2 inch x 1 2 inch x 0 050 inch 2024 T42 extruded angle or a heavy J section The angle should be inserted fore and aft across the center of the wrinkle and should extend to within 1 16 inch to 1 8 inch of the fuselage bulkheads comprising the end of the bay Rivet pattern should be similar to existing manufactured seam at edge of sheet Negligible damage to stringers formed skin flanges bulkhead channel and like parts is similar to that for the wing skin Refer to Chapter 57 Wing Damage Classification for a definition of negligible damage to these components Repairable Damage A If a skin is badly damaged repair must be made by replacing an entire skin panel from one structural member to the next Repair seams must be made to lie along structural members and each seam must be made exactly the same in regard to rivet size spacing an
39. the beam and keep the beam 90 degrees to the hinge line of control surface Paint is a considerable weight factor In order to keep balance weight to a minimum it is recommended that existing paint be removed before adding paint to a control surface Increase in balance weight will also be limited by the amount of space available and clearance with adjacent parts Good workmanship and standard repair practices should not result in unreasonable balance weight The approximate amount of weight needed may be determined by taping loose weight at the balance weight area Lighten balance weight by drilling off part of weight Make balance weight heavier by fusing bar stock solder to weight after removal from control surface The ailerons should have balance weight increased by ordering additional weight and gang channel listed in applicable Parts Catalog and installing next to existing inboard weight the minimum length necessary for correct balance except that a length which contains at least two attaching screws must be used If necessary lighten new weight or existing weights for correct balance 4 Balancing Definitions A Overbalance refer to Figure 3 is defined as the condition that exists when surface is leading edge heavy and is defined by symbol If the balance beam uses a sliding weight the weight must be on the trailing edge side of the hinge line to balance the control surface the control surface is considered to be overbalan
40. the manufacturers standard sheets NA eO NOTE When installing a blind rivet pull type rivet in a hole where the previous blind rivet was removed by drilling and punching the rivet out inspect the drilled hole to assure all metal sheets are in place and not separated prior to pulling rivet It may be necessary to insert a stiff wire in adjacent hole to hold metal in position while pulling rivet 9 When placing pulling head on rivet stem hold riveter and pulling head in line with axis of rivet while holding tool in a light and flexible manner 0 When tool is actuated pulling head will pull down and seat against rivet head 11 Clamping action will pull sheets together and seat rivet when tool is actuated 12 When tool is actuated action of rivet will automatically assist in bringing tool and pulling head into proper alignment with rivet axis NOTE Pressing down with force will not allow rivet and tool to align themselves with hole and could limit head setting of rivet however enough force to seat the head against the skin is necessary 13 Hold tool in line with rivet as accurately as possible and allow a steady but light pressure pull trigger and let the rivet align itself 14 When rivet is completely installed release trigger and pulling head will eject pulling portion of stem through forward end 15 Rivet must break within these limits Faser Desmumbe SemFuswess NAS1738 or NAS1739 All 40 010 or 0 0
41. wing station and place bolt A on mark c Check to assure that protractor bubble is still centered If proper twist is present the protractor readings will be the same parallel NOTE Forward or aft bolt may be lowered from wing 0 10 inch maximum to attain level indication 57 22 00 605 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 359 Bolt Bolt B Bolt A Bolt dimension dimension dimension Wing station inches inches inches or B 2 00 1 00 29 50 39 00 2 00 1 00 29 50 100 00 0 66 1 00 20 00 207 00 d FWD eg ne rte Y Wing tp WS 207 00 WS 100 50 WS 39 00 NOTE All wing twist occurs between WS 100 50 and the tip rib View looking down on wing 262072001 Measuring Model 206 T206 Series Wing Twist Figure 603 Sheet 1 57 22 00 Page 606 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL WING STRINGER REPAIRS 1 Stringer Damage Classification A Damage to the wing stringers can be divided into three major categories and is detailed in Wing Damage Classification 2 Stringer Repair A Repairs to wing stringer are similar to repairs to fuselage stringers Refer to Chapter 52 Stringer and
42. 0 800 0 063 0 160 5 1969 1 2024 T3511 0 625 0 813 0 050 0 050 0 075 0 075 0 050 0 625 0 813 0 063 0 160 Extrusions and Formed Sections Figure 1 Sheet 1 51 30 00 Page 2 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B2096 B R r x Loc EM 4 A C EXTRUDED ANGLE ORIGINAL MATERIAL SUBSTITUTION STD 2024 SHEET SHAPE MATERIAL A B C D r 5 109 2024 T3511 0 500 0 500 0 050 0 050 0 050 0 500 0 500 0 050 0 120 e ANGLE b EXTRUDED ANGLE ORIGINAL MATERIAL SUBSTITUTION STD 2014 SHEET SHAPE MATERIAL B D 5 S 214 2024 3511 1 000 1 000 0 125 0 125 0 125 0 094 95 45 1 000 1 000 0 125 0 125 95 45 Extrusions and Formed Sections Figure 1 Sheet 2 51 30 00 Page 3 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B2097 EXTRUDED BULB BAR ORIGINAL MATERIAL SUBSTITUTION STD 2024 SHEET SHAPE MATERIAL A B C R b r S 211 2024 T3511 1 450 0 250 0 125 0 125 0 125 1 450 0 250 0 125 0 125 Re ROLLED J SECTIONS DIMENSIONS INCHES MATERIAL 2024 T3 OR T4 CLAD ALUMINUM ALLOY STD AREA SHAPE T L H R SQ INCHES 5 49 0 032 0 750 0 620 0 125 0 125 0 0500 S 282 0 032 0 620 0 620 0 120 0 120 0 0456 5 1113 1 0 020 0 750 0 620 0 125 0 125 0 0314 Extrusions and Formed Sections Figure 1 Sheet 3 51 30 00 4 Cessna Aircraft
43. 10 00 Page 1 Generals 4 ue t TE ad VER DRE DR UP ERA Rd 51 10 00 Page 1 Damage Investigation 51 10 00 Page 1 Damage 51 10 00 Page 2 Refinishing Damaged Areas Following Repairs 51 10 00 Page 2 CORROSION AND CORROSION CONTROL GENERAL 51 11 00 Page 1 General cx e E eos e a a e c LOK RR Aces REM uta 51 11 00 Page 1 Types of Corrosion cec nae DI ed Po 51 11 00 Page 1 Typical Corrosion Areas 1 51 11 00 Page 3 Corrosion DeteCON re mane QUE ee OPERAR 51 11 00 Page 4 Corrosion Damage 51 11 00 Page 4 Corrosion Removal gt er Ra a d EPI OP E DA A 51 11 00 Page 5 Control of Corrosion on Landing Gear 05 51 11 00 Page 7 REPAIR MATERIALS ee ERG 51 30 00 1 EE ETE PE PE E E etate at EEE a 51 30 00 Page 1 Repair Materials y AI oL eh alas 51 30 00 Page 1 Extrusions and Formed Sections 51 30 00 Page 1 FASTENERS
44. 2 00 Page 803 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL WING RIB General A Flanged upper and lower edges of all ribs serve as cap strips in addition to providing rigidity to the rib The skin riveted or bonded directly to each rib flange provides the cellular strength for each successive bay The nose center and trailing segments are riveted together through the front and rear spars to form the basic airfoil section Spanwise Alclad stringers stiffen the skin between ribs Wing Rib Damage Classification A Damage to the wing rib can be divided into three major categories and is detailed in Wing Damage Classification Wing Rib Repair A Repairs to the wing rib are illustrated in Figure 801 57 20 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL X N gt NS MS20470AD4 Rivets B187 9 Ve 29 o eo X d qA gt Oo ND 2 o og Eg g D 2 on 9 Doe 8209 Pose 5 5 5 5 ES 0 25 Inch edge margin 0 75 Inch rivet spacing Original parts 222 Repair parts view A A Repair parts cross section 268211049 Typical Rib Repair Figure 801 Sheet 1 P
45. 20 inch Cherry Max 0 010 or 0 015 inch Cherry Max 0 010 or 0 020 inch 16 Protruding stems usually indicate incorrect grip length or oversize holes D Loose or Working Blind Rivets 1 Blind rivets which are found to be loose or show evidence of working must be replaced with rivets of like size and type In some instances it may be necessary to use the next larger size rivet Loose fasteners may be indicated by the following situation a fastened material moves relative to the fastener Skin deflection is evident 51 40 00 Pages Cessna Aircraft Company Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 152 i CK Xa KKK do hu d Wrong Right Obstruction BK minimum blind clearance lt lt KX d Co E Misaligned hole U CY Wrong Right Eject stem Pulling head 2682T1019 Installation of Blind Rivets Figure 2 Sheet 1 51 40 00 Page9 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL b Tipping of the fastener head may indicate its looseness or slippage Rivet head periphery rolled upward also indicates looseness c A
46. 802 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B200 0 031 inch maximum Square cut end gap Splice member of original tube es D PP S gt L e S D L 1 00 inch or 1 25D or B C B orB whichever is greater Cut out damaged portion of tube prepare splice member and cut sleeves lt gt EP FB Slip sleeves over Slip sleeves over joints and weld joints and weld in place in place Sleeves may be rotated to Completed splice using diagonally cut sleeves suit conditions and to 0 031 inch provide maximum maximum gap reinforcement 2 orB Completed splice using fishmouth cut sleeves 2682T1052 Typical Engine Mount Repairs Figure 801 Sheet 2 71 20 00 Page 803 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B201 0 25 of tube circumference maximum with no wall penetration 45 to 70 Less 2A than 2A A A Area of damage chafing smooth notches Fillet weld and smooth gouges all around 2682T1053 Typical Engine Mount Repairs Figure 801 Sheet 3 71 20 00 Page 804 Cessna Aircraft Company Jun 1 2005
47. 8B4 NAS1398D4 0 020 NAS1738B4 NAS1738D4 MS20470AD4 0 050 NAS1398B4 NAS1398D5 0 040 NAS1398B5 NAS1398D5 NAS9301B5 NAS1738B4 NAS1738E4 NAS1738D4 NAS9301B4 0 032 NAS1738B5 NAS1738E5 NAS1738D5 NAS9301B5 MS20470AD5 0 063 NAS1398B5 NAS1398D5 0 050 NAS1398B6 NAS1398D6 NAS1738B5 NAS1738E5 CR3213 5 0 040 NAS1738B6 NAS1738E6 NAS1738D5 CR3213 6 MS20470AD6 0 080 NAS1398B6 NAS1398D6 0 071 NAS1398D6 0 063 NAS1738B6 NAS1738E6 NAS1738D CR3213 6 MS20426AD3 Countersunk 0 063 NAS1398B4 NAS1399D4 Refer to Note 1 0 040 NAS1739D4 MS20426AD4 Countersunk 0 080 NAS1399B4 NAS1399D4 CR3213 4 0 050 NAS1739D4 MS20426AD4 Dimpled 0 063 NAS1739B4 NAS1739E4 51 40 00 Page3 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Table 1 Approved Replacement Fasteners Chart continued REPLACE Inch thickness or thicker WITH MS20426AD5 Countersunk 0 063 NAS1739D5 NAS1739B5 NAS1739E5 0 050 CR3242 5 MS20426AD5 Dimpled 0 071 NAS1739B5 NAS1739E5 NOTE 1 Rework Required Countersink oversize to accommodate oversize rivet NOTE2 GENERAL NOTE Do not use blind rivets in any portion of the engine air induction system structure Table 2 Approved Fastener Substitutions Fastener Collar REPLACE NAS178 NAS179 NAS1054 NAS179 NAS528 DIAMETER Refer to Notes 1 2 6 7 Refer to Notes 1 4 Refer to Notes 1 4 Refer to Notes 1 3 4 Refer to
48. 99 Processed or Oxide Finer Ferrous Does Not Corrosion 150 180 Fine to Alloys Apply to Removal Grit Grit Ultra Steel Heat or or or Fine Treated to Fairing Finer Finer X x x Strengths to 220 000 psi and ere Above Finishing 400 X X X Aluminum Not Use Corrosion 150 7 0 Grit Very Alloys Silicon Removal Grit or Finer Fine Except Carbide or or and X X X Clad Abrasive Fairing Finer Ultra Aluminum Fine Finishing 400 X X Clad Sanding Corrosion 240 7 0 Grit Very Aluminum Limited Removal Grit orFiner Fine to the or or and x Removal Fairing Finer Ultra x of Minor Fine Scratches Finishing 400 X Magnesium Corrosion 240 Very Alloys Removal Grit Fine or or and Fairing Finer Ultra Fine Finishing 400 X X Abrasives for Corrosion Removal Figure 2 Sheet 1 51 11 00 Pages Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 4 Remove residual corrosion by hand sanding 5 After removing all corrosion visible through a magnifying glass fair depression resulting from rework and finish with 400 grit abrasive paper 6 Clean reworked area 7 Determine depth of rework area to ensure rework limits are not exceeded 8 Prime using rust inhibitive primer within one hour of rework 9 Reapply finish topcoat if required 7 Control of Corrosion on Landing Gear Springs A General 1 2 3 4 main landing gear sp
49. Airplane Identification INTRODUCTION Page 1 Aerofiche microfiche INTRODUCTION Page 1 Using the Structural Repair Manual or Aerofiche INTRODUCTION Page 2 Revision Manual errer usa gid ie eae eee INTRODUCTION Page 2 Identifying Revised Material 0 cece cece eee eee eee ttt INTRODUCTION Page 3 LISTCOPREVISIONS 2 iexuch RE nes bes a nei LIST REVISIONS Page 1 General tench dali Mee tate LIST REVISIONS Page 1 EST OF GHAPTERS LIST OF CHAPTERS Page 1 CONTENTS Page 1 of 1 Cessna Aircraft Company Jun 1 2005 1 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL INTRODUCTION General A The information in this publication is based on data available at the time of publication and is updated supplemented and automatically amended by all information issued in Service News Letters Service Bulletins Supplier Service Notices Publication Changes Revisions Reissues and Temporary Revisions All such amendments become part of and are specifically incorporated within this publication Users are urged to keep abreast of the latest anendments to this publication through information available at Cessna Author
50. Bulkhead Repair Figure 801 Sheet 3 Page 804 Jun 1 2005 53 20 00 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B175 Bulkhead Bad rivet hole or damaged stringer Fillers same thickness as stringers extend beyond doubler and pick up as many rivets as doubler does oO 8 Doubler Bulkhead 0 2 Z L Stringer Rivets same type and diameter as original N Original parts 222 Z ZZ Repair parts Ee Repair parts in cross section 268271028 Typical Rib Repair Figure 802 Sheet 1 53 20 00 Page 805 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL STRINGER AND CHANNEL REPAIR 1 General A Damageto the stringers or channels can be repairable Refer to Figure 801 for an illustration of typical stringer and channel repairs 53 30 00 Page 801 Cessna Aircraft Company Jun 1 2005 Pick up existing rivet spacing CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 0 25 inch edge margin B176 T42 same Doubler 2024 gage as stringer Five rivets each side of damaged area Clean out Filler 2024 T42 same gage as stringer view A A MS20470AD4 rivets Original parts ZB Repair parts Orig
51. Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B R TYPICAL 4 H B2008 R TYPICAL TYPE 4 E tL ROLLED HAT SECTIONS DIMENSIONS INCHES MATERIAL 2024 T3 OR T4 CLAD ALUMINUM ALLOY STD AREA SHAPE TYPE A B H R T SQ INCHES S 370 A 7 000 2 000 0 650 0 150 0 032 0 0889 5 1039 1 B 0 500 1 500 0 380 0 060 0 032 0 0661 5 1126 B 1 050 2 200 0 500 0 090 0 025 0 0744 5 1541 1 B 1 060 2 200 0 750 0 090 0 032 0 1106 R TYPICAL zT C 195 ROLLED HAT SECTIONS DIMENSIONS INCHES MATERIAL 2024 T4 CLAD ALUMINUM ALLOY STD AREA SHAPE A B C R T SQ INCHES S 1162 0 650 0 700 0 600 0 150 0 025 0 0691 Extrusions and Formed Sections Figure 1 Sheet 4 51 30 00 Page 5 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL FASTENERS General A Fasteners used in the airplane are generally solid aluminum rivets blind rivets and steel threaded fasteners Usage of each is primarily a function of the loads to be carried accessibility and frequency of removal Rivets used in airplane construction are usually fabricated from aluminum alloys special cases monel corrosion resistant steel and mild steel copper and iron rivets are used Rivets A Standard solid shank MS rivets are those generally used in airplane construction They are fabricated in the following head types roundh
52. E STATIC BALANCE LIMITS APPROVED FOR FLIGHT CONFIGURATION INCH LBS AILERON 0 0 TO 11 31 RUDDER 0 0 TO 9 0 LEFT ELEVATOR 0 0 TO 18 5 RIGHT ELEVATOR 0 0 TO 24 5 51 60 00 Page 7 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Table 2 Model 182 Static Balance Limits CONTROL SURFACE STATIC BALANCE LIMITS APPROVED FOR FLIGHT CONFIGURATION INCH LBS AILERON 0 0 TO 9 64 RUDDER 0 0 TO 6 0 LEFT ELEVATOR 0 0 TO 20 47 RIGHT ELEVATOR 0 0 TO 20 47 Table 3 Model 206 Static Balance Limits CONTROL SURFACE STATIC BALANCE LIMITS APPROVED FOR FLIGHT CONFIGURATION INCH LBS AILERON 0 0 TO 3 0 RUDDER Landplane 4 0 TO 3 0 LEFT ELEVATOR 0 0 TO 12 1 RIGHT ELEVATOR 0 0 TO 12 1 51 60 00 Page 8 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL REPAIRS GENERAL Introduction A Many components of the airframe structure are similar in design and fabrication Examples of such items are sheet metal webs formed structural shapes and extrusions Typical repairs to these and other items have been compiled in this section to eliminate the duplication of repairs under each applicable component Repairs in this section apply to the member shown regardless of location on the airplane structure except as limited and will include only those parts or members necessary to show the typical situation
53. ESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Soft A 1100 aluminum shank rivets or B 5056 aluminum shank rivets are also used to install non metallic parts Original equipment soft rivets will be either red or green colored under the paint If the butt or driven end of the rivet is adjacent to the non metallic part it is preferable to install a washer over the shank to prevent the rivet shank which swells during driving from overloading the non metallic hole The hole in the washer should match the specified installation hole for the fastener If the tail end of the rivet is installed through metal substructure the washer is not necessary Take care when driving rivets through non metal to not overdrive the rivet If the rivet is overdriven the shank will swell even with the washer in place The rivet butt should be driven to no more than necessary to retain the part never more than 1 4 times the shank diameter If the original equipment rivet provided connection between metal parts as well as non metallic parts it may be a standard AD rivet Original equipment AD rivets are colored gold or uncolored Replace original equipment AD rivets with AD rivets 51 40 00 Page 12 Cessna Aircraft Company Jun 1 2005 1 General A B CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL FLIGHT CONTROL SURFACE BALANCING This section applies to the balancing of the ailerons elevators and rudder Co
54. GLE ENGINE STRUCTURAL REPAIR MANUAL Add doubler B172 N Aw amp Existing structure 2682T1014 88 5 ES Typical Cabin Bulkhead Repair Figure 801 Sheet 1 Original parts Repair parts in cross section Repair parts p Page 802 Jun 1 2005 53 20 00 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B173 WS NN Flatten lightening hole as required N N JA N NEA NY WK ASN ANN CRM Add angles and cap as shown view Repair of cracked bulkhead Original parts NOTE Add repair parts of next larger gage material ZI Repair parts N Refer to FAA advisory circular AC43 13 1A Repair parts in cross section Figure 2 28 for required gage thickness and number of rivets 2682T1015 Typical Cabin Bulkhead Repair Figure 801 Sheet 2 Jun 1 2005 53 20 00 Page 803 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B174 NOTE Add bulkhead segment length Original parts to extend to adjacent stringer or approximately seven N Repair parts inches each side of defect 2682T1016 view Make from bulkhead section with center section removed Repair parts in cross section Typical Cabin
55. Locate new access holes aft of the front spar between the first and second stringers aft of the spar When installing the doubler rotate it so the two straight edges are closest to the stringers Alternate bays with new access holes staggered forward and aft of the front spar are preferable A maximum of five new access holes in each wing is permissible If more are required contact Cessna Propeller Aircraft Product Support B Access Hole Installation Refer to Figure 2 1 2 Establish exact location for inspection cover and inscribe centerlines Determine position of doubler on wing skin and center over centerlines Mark the ten rivet hole locations and drill to size shown 57 00 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 3 Cut out access hole using dimension shown 4 Flex doubler and insert through access hole and rivet in place 5 Position cover and secure using screws as shown Cessna Aircraft Company 57 00 00 Page 2 Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 185 12 inch wide heavy canvas 1x12x48 1 12 11 1 12 30 75 30 75 2 4 20 1 12 8 5 inch wide cotton NOTE All dimensions shown in inches or nominal lumber size Wing and Fuselage Support Stands Figure 1 Sheet 1 0582T1005 57 00 00 Page 3 Cessna Aircr
56. NTS DOORS GENERAL 41 25 53 5 eb noord visas Phen Rae deter seeds ends 52 00 00 Page 1 Gerieral LR OR KK RR IRA 52 00 00 Page 1 DOOR DAMAGE 52 10 00 Page 101 Repairable 1 52 10 00 Page 101 CONTENTS Page 1 of 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL DOORS GENERAL 1 General A B Chapter 52 describes general repair practices materials and procedures which are applicable to the doors and door structure If questions arise concerning approved repairs or for repairs not shown in this section contact Cessna Propeller Aircraft Product Support 52 00 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL DOOR DAMAGE CLASSIFICATION 1 Repairable Damage A Bonded doors may be repaired by the same methods used for riveted structure Rivets are a satisfactory substitute for bonded seams on these assemblies The strength of the bonded seams in doors may be replaced by a single 3 32 2117 AD rivet per running inch of bond seam The standard repair procedures outlined in AC43 13 1b are also applicable to bonded doors 52 10 00 Page 101 Cessna Aircraft Company Jun 1 2005 FUSELAGE CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCT
57. P O Box 7706 Wichita KS 67277 USA Telephone 316 517 5800 or Facsimile 316 942 9006 71 20 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B199 Original tube Welds Splice member ES M Pulling d Internal snug trimmed and fitting sleeves welded o gt ze Internal snug fitting sleeves Internal sleeve 0 625 inch sleeve 0 125 inch gap At Number 40 hole D MEL D L 1 00 inch or 1 25D whichever is greater D Sleeve will end here L Mark cut and ___ End A drill tube end B a H H 1D 1 25D End B Push 0 063 inch welding rod thru hole and out end of tube Drill number 40 Weld end A of weldin hole slanted 58 pee de a 4252 9 mark around 0 063 inch center of sleeve welding rod Chamfer end to aid sliding and mark sleeve and weld on end of sleeve pulling wire 0 125 inch gap ee Splice Original tube Splice member Original tube Y L 9 y y a N Pull end B Internal sleeve Internal sleeve Trim sleeve pulling wire and weld Insert sleeve into splice hole after completion of repair member and align tubes 7 Pull welding rod until center marks line up 268211051 Typical Engine Mount Repairs Figure 801 Sheet 1 71 20 00 Page
58. RAL REPAIR MANUAL LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT PAGE DATE 71 Title 71 List of Effective Pages 71 Record of Temporary Revisions 71 Table of Contents 71 00 00 Page 1 Jun 1 2005 71 10 00 Page 801 Jun 1 2005 71 20 00 Pages 801 804 Jun 1 2005 71 LIST OF EFFECTIVE PAGES Page 1 of 1 Cessna Aircraft Company Jun 1 2005 RECORD TEMPORARY REVISIONS Temporary Revision Number Page Number Issue Date By Date Removed CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CONTENTS POWERPLANT GENERAL 0 0 00000 cece cece eee eee eee General s eI Sek ENGINE COWLING REPAIRS General Pe huele erem AA Repair of Cowling Repair of Reinforcement 1 DYNAFOCAL TYPE ENGINE MOUNT REPAIRS Getieral 225 beer n ees tensa e EAE ES eaa Engine Mount Cessna Aircraft Company 71 00 00 1 p EL 71 00 00 Page 1 LS emus 71 10 00 Page 801 wade ae ondas 71 10 00 Page 801 TET 71 10 00 Page 801 mE 71 10 00 Page 801 Lii Na fus 71 20 00 Page 801 a
59. Rivet length is based on the grip Solid Shank Rivets A Removal of Solid Shank Rivets Refer to Figure 1 1 When it becomes necessary to replace a rivet extreme care should be taken in its removal so that the rivet hole will retain its original size and replacement with a larger size rivet will not be necessary If the rivet is not removed properly the strength of the joint may be weakened and the replacement of rivets made more difficult 2 When removing a rivet work on the manufactured head It is more symmetrical about the shank than the shop head and there will be less chance of damaging the rivet hole or the material around it To remove rivets use hand tools a power drill or a combination of both The preferred method is to drill through the rivet head and drive out the remainder of the rivet with a drift punch First file a flat area on the head of any round or brazier head rivet and center punch the flat surface for drilling On thin metal back up the rivet on the shop head when center punching to avoid depressing the metal The dimple in 2117 T3 rivets usually eliminates the necessity of filing and center punching the rivet 3 Select a drill one size smaller than the rivet shank and drill out the rivet head When using a power drill set the drill on the rivet and rotate the chuck several revolutions by hand before turning on the power This procedure helps the drill cut a good starting spot and eliminates the chance of the
60. Textron Company Maintenance Manual SINGLE ENGINE MODELS 172 182 1182 206 AND T206 1996 And On 2 DECEMBER 1996 REVISION 4 1 JUNE 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT PAGE 00 Title 00 List of Effective Pages 00 Record of Revisions 00 Record of Temporary Revisions 00 Table of Contents DATE INTRODUCTION LIST OF REVISIONS LIST OF CHAPTER Cessna Aircraft Company Pages 1 3 Jun 1 2005 Page 1 Jun 1 2005 Page 1 Jun 1 2005 00 LIST OF EFFECTIVE PAGES Page 1 of 1 Jun 1 2005 Revision Number Date Inserted Date Removed Page Number Revision Number Date Inserted Date Removed Page Number RECORD TEMPORARY REVISIONS Temporary Revision Number Page Number Issue Date By Date Removed CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CONTENTS INTRODUCTION 2 532 rent ementi dba INTRODUCTION Page 1 e s o ecu oe Del o tag aia avin wie INTRODUCTION Page 1 Seal Sr de at hae INTRODUCTION Page 1
61. Title 55 List of Effective Pages 55 Record of Temporary Revisions 55 Table of Contents 55 00 00 55 10 00 55 20 00 55 30 00 55 40 00 Cessna Aircraft Company PAGE Page 1 Page 801 Page 801 Page 801 Page 801 DATE Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 55 LIST OF EFFECTIVE PAGES Page 1 of 1 Jun 1 2005 RECORD TEMPORARY REVISIONS Temporary Revision Number Page Number Issue Date By Date Removed CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CONTENTS HORIZONTAL AND VERTICAL 5 General I s neti nels HORIZONTAL 2 Horizontal 2 Negligible hope ede pee een ds Repairable Replacement ELEVATORS 4 ace te a at meteo d a dte itt vk enc p Negligible Repairable
62. UCTURAL REPAIR MANUAL PLASTIC WINDOW SURFACE REPAIR 1 Repair of Plastic Window Surfaces A Damaged window panels and the windshield on the airplane are normally removed and replaced if the damage is extensive However certain repairs as described in the following paragraphs can be accomplished without removing the damaged part from the airplane Three types of temporary repairs for cracked plastic are possible No repairs of any kind are recommended on highly stressed or compound curves or where the repair would be likely to affect the pilots or copilot s field of vision during normal flight or landing operations Curved areas are more difficult to repair than flat areas and any repaired area is both structurally and optically inferior to the original surface Refer to Figure 801 for an illustration of typical windshield and window repair NOTE If temporary repairs are made operations should be kept to a minimum until replacement of window can be made 2 Tools and Materials NAME NUMBER MANUFACTURER USE Novus 1 Novus Co To polish scratches Minneapolis MN 55435 out of windows Novus 2 Novus Co To polish scratches out of windows Methylene Chloride Commercially Available Solvent for repair of windows 3 Stop Drilling A following procedure should be used when stop drilling 1 When a crack appears in a panel drill a hole at the end of the crack to prevent further spreading The hole should be approximately 1 8 i
63. URAL REPAIR MANUAL LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT 53 Title 53 List of Effective Pages 53 Record of Temporary Revisions 53 Table of Contents 53 00 00 53 10 00 53 20 00 53 30 00 53 40 00 Cessna Aircraft Company PAGE Pages 1 3 Pages 1 2 Pages 801 805 Pages 801 806 Pages 801 803 DATE Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 53 LIST OF EFFECTIVE PAGES Page 1 of 1 Jun 1 2005 RECORD TEMPORARY REVISIONS Temporary Revision Number Page Number Issue Date By Date Removed FUSELAGE GENERAL General Fuselage CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CONTENTS FUSELAGE DAMAGE CLASSIFICATION General Negligible Repairable Replacement CABIN BULKHEAD REPAIR 2 2 General Repair of Webs or Repair ot Channels 2 c Rene CO RETE RRERTe dae oe Landing Gear Bulkheads
64. a 5 lt lt gt gt 25 Inch edge marg 22 e a 5 A e 2 7 N N Doubler 2024 T42 same gage as channel Filler 2024 T42 MS20470AD4 Rivets same gage as channel Original parts Repair parts BEER Repair parts in cross section Typical Stringer and Channel Repair Figure 801 Sheet 4 Jun 1 2005 53 30 00 Page 805 Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B180 4 4 ia E 1 view Original parts Ba Repair parts margin typical edge 0 75 inch rivet spacing Page 806 Jun 1 2005 0582T1002 53 30 00 Figure 801 Sheet 5 Typical Stringer and Channel Repair Cessna Aircraft Company Repair parts in cross section CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL FIREWALL REPAIR 1 General A firewall is constructed of 0 016 inch 18 8 corrosion resistant annealed stainless steel sheet 1 A typical firewall patch is illustrated in Figure 801 2 typical repair to the interior firewall angle is illustrated in Figure 802 2 Material NAME NUMBER MANUFACTURER USE Firewall Sealant Pro Seal 700 A Loctite Aerospace Firewall sealant AMS 3374 Bay Point CA 94565 3 Repairing the Firewall Assembly A Firewall sheets may be repaired by removing damaged mater
65. ad acid type battery electrolyte If baking soda is not available flood the area with water Stainless Steel control cables a Checking for corrosion on control cables is normally accomplished during the preventative maintenance check During preventative maintenance broken wire and wear of the control cable is also checked 51 11 00 Page3 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL b If the surface of the cable is corroded carefully force the cable open by reverse twisting and visually inspect the interior Corrosion on the interior strands of the cable constitutes failure and the cable must be replaced If no internal corrosion is detected remove loose external rust and corrosion with a clean dry coarse weave rag or fiber brush NOTE Donotuse metallic wools or solvents to clean installed cables Use of metallic wool will embed dissimilar metal particles in the cables and create further corrosion Solvents will remove internal cable lubricant allowing cable strands to abrade and further corrode c After thorough cleaning of the exterior cable surface apply a light coat of lubricant VV L 800 to the external cable surface 4 Corrosion Detection A The primary means of corrosion detection is visual but in situations where visual inspection is not feasible other techniques must be used The use of liquid dye penetrants magnetic particle X ray and ultr
66. aft Company Jun 1 2005 Number 40 0 098 inch diameter hole 10 required Lower wing skin reference CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 1443 1 doubler 1 required 5 062 inch diameter hole 1 required MS20426AD53 rivets 10 required 225 4F cover 1 required Screw 6 required DETAIL View looking down from inside wing at top of skin 058271006 Access Hole Installation Figure 2 Sheet 1 57 00 00 Page 4 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL WING DAMAGE CLASSIFICATION Damage Classification A B Damage to the wing and its component assemblies can be divided into three major categories negligible damage repairable damage and damage necessitating replacement of parts These categories are intended to provide the mechanic with some general guidelines to use in determining the extent and criticalness of any damage Obviously there will be some overlapping between categories and common sense should be used in determining the final action to be taken with regard to any damage For an illustration of various wing component repairs refer to applicable sections within this chapter Wing Skin Damage Criteria A Negligible damage Any smooth dents in the wing skin that are not more than 0 030 inch below contour and can be circumscribed with a 2 inch diameter circle that have
67. age 802 Jun 1 2005 57 20 00 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B188 0 75 inch rivet spacing damaged area Clean out Filler 2024 T4 Alclad around damaged One row rivets area 5 6 N ANN Doubler 2024 T3 Alclad n b N N D lt 5 o ps t Q S ot 2 Repair parts 0582T1007 Jun 1 2005 57 20 00 Page 803 view A A Typical Rib Repair Figure 801 Sheet 2 Cessna Aircraft Company Repair parts cross section CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL WING SPARS General A Front and rear spars are of riveted construction Damage Classification A Damage to the wing spar can be divided into three major categories and is detailed in Wing Damage Classification Spar Repair A Repairs to the wing spar are illustrated in Figure 801 57 21 00 Page 801 Cessna Aircraft Company Jun 1 2005 6 4 4 22 2 Doubler 2024 4 Alclad VIEW A A NW Filler 2024 4 Alclad MS20470AD4 rivets Clean out damaged edge margin area 0 375 inch typical ri 2 NS N lt Jun 1 2005 058211008 57 21 00 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B189 F N N
68. ailable NOTE 2 NOTE 3 NOTE 4 NOTE 5 NOTE 6 NOTE 7 NOTE 8 NOTE 9 Available in oversize for repair of elongated holes Ream holes to provide a 0 001 inch interference fit NAS529 4 thru 12 take NAS528 same dash number NAS529 14 thru 20 take NAS179 Steel shank fastener designated for drive on collars Choose protruding head only Steel shank fastener designated for squeeze on collars Installation requires sufficient space for the tool and extended shank of the fastener Choose protruding head only Threaded fastener Preferred substitute fastener When you substitute a threaded fastener for a high strength steel shank rivet use one of these steel nuts AN365 MS20365 517825 MS21044 MS21045 551943 or NAS1079 Approval of the use of these nuts in this application does not constitute a general approval to use these nut on high strength bolts GENERAL NOTE These fastener substitutions address shear strength and hole tolerances only The specific application may not allow all of these substitutions because of space considerations The United States Department of Defense no longer maintains MS and NAS standards Identical parts may have MS NASM or AIA NAS part numbers EXAMPLE 520470 04 6 rivets may also be identified as NASM20470AD4 6 NAS1738M4 4 rivets may be identified as AIA NAS1738M4 4 Installation of Solid Shank Rivets 1 Alarge percentage of riveting of airplane structure is accomplished
69. aileron skin Original part Repair patch in cross section 0582T1001 Corrugated Skin Repair Figure 802 Sheet 1 51 75 00 Page 807 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Any control surface that has a crack that progresses past a stop drilled hole shall be repaired or replaced A control surface that has any of the following conditions shall have a repair made as soon as practical a Acrack that is longer than 2 inches b Acrack that does not originate from the trailing edge or a trailing edge rivet c Cracks in more than six trailing edge rivet locations per skin Affected control surfaces with corrugated skins and having a stop drilled crack that does not extend past the stop drilled hole may remain in service without additional repair Refer to Figure 802 as applicable for repair information 51 75 00 Page 808 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CONTROL SURFACE REPAIR 1 General A Damage which would involve a control surface repair After the repair is completed the control surface balance must be checked as described in Flight Control Surface Balancing Refer to Figures 801 and 802 which illustrate typical control surface repairs 51 76 00 Page 801 Cessna Aircraft Company Jun 1 2005 dimensions shown in inches
70. also cannot be considered negligible Repairable Damage A Skin patches may be used to repair skin damage These patches are illustrated in Chapter 51 Typical Skin Repairs Figure 801 For skin damage which includes corrugations refer to Chapter 51 Typical Skin Repairs Figure 802 Flight control surfaces must be balanced after repair or painting in accordance with balancing procedures outlined in Chapter 51 Flight Control Surface Balancing Replacement Damage A Warped and cracked skin ribs and hinge brackets are replaceable items Where damage is extensive replacement of the entire assembly is recommended 55 20 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL VERTICAL STABILIZER General A The vertical stabilizer is of conventional aluminum construction utilizing spars ribs and skins Vertical Stabilizer and Dorsal A The vertical stabilizer and dorsal are constructed jointly to form a single unit Negligible Damage A The same criteria which is used to define negligible damage to the fuselage may be applied to the vertical stabilizer Refer to Chapter 53 Fuselage Damage Classification for a complete description of negligible damage Repairable Damage A Skin damage exceeding that considered negligible that can be repaired as illustrated in Chapter 51 Typical Skin Repairs Figure 801 For skin damage which includes c
71. and is detailed in Wing Damage Classification Leading Edge Repairs A Repairs to the leading edge are illustrated in Figure 801 Notes and Repair Limits A following notes and repair limits are applicable to lading edge repairs 1 2 3 4 5 6 Dimple leading edge skin and filler material counter sink the doubler Use MS20426AD4 rivets to install filler except where bucking is impossible Use blind rivets where regular rivets cannot be bucked Contour must be maintained After repair has been completed use epoxy filler as necessary and sand smooth before painting Vertical size of patch is limited by ability to install doubler clear of front spar Lateral size is limited to seven inches across trimmed out area Number of repairs is limited to one per bay 57 25 00 Page 801 Cessna Aircraft Company Jun 1 2005 1 00 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 196 E o 5 P 2 a o a 5 S o BE o 5 a 1 x 98 cx NO gt gt Oo no 2t 8 9 5 1 Yo A 2a O amp v 2 v 0 Los ges vL S305 Qm 22523 gt 2 Original parts ZA Repair parts Page 802 Jun 1 2005 0582T1013 57 25 00 Leading Edge Repair Fi
72. arden due to aging processes and show hairline cracks around the edge of the rivet heads This should not be used as a basis for determining whether or not the rivet is loose The hardened paint may crack at times and collect dirt or exhaust fumes which will appear as discoloration It is not possible to detect loose rivets visually Replacement rivets should be of like size and type In some instances however it will be necessary to use the next size larger diameter For general repair practices the spacing between the centerlines of adjacent rivet holes shall be four diameters or greater In some areas where the spacing between rivets prohibits the use of the next larger rivets special repair instructions and procedures shall be followed Contact Cessna Single Engine Support 8 Blind Rivets A General 1 2 Blind rivets are intended for use where access is available to only one side of the work Replacement of solid rivets with blind rivets should only be accomplished within the guidelines of Table 1 when the installation of a solid shank rivet is not possible Blind rivets do not have the same resistance to corrosion and fatigue as solid shank rivets and should not be considered a universal replacement for solid shank rivets Removal of Blind Rivets CAUTION Do not drill completely through the rivet sleeve This method of removing 1 2 a rivet will tend to enlarge the hole Use a small center drill to provide a gui
73. asonic devices can be used but most of these sophisticated techniques are intended for the detection of physical flaws within metal objects rather than the detection of corrosion 1 2 Visual Inspection a A visual check of the metal surface can reveal the signs of corrosive attack the most obvious of which is a corrosive deposit Corrosion deposits of aluminum or magnesium are generally a white or grayish white powder while the color of ferrous compounds varies from red to dark reddish brown 1 The indications of corrosive attack are small localized discoloration of the metal surface Surfaces protected by paint or plating may only exhibit indications of more advanced corrosive attack by the presence of blisters or bulges in the protective film Bulges in lap joints are indications of corrosive buildup which is well advanced In may cases because the inspection area is obscured by structural members equipment installations or for other reasons it is awkward to check visually In such cases mirrors boroscopes or like devices must be used to inspect the obscured areas Any means which allows a thorough inspection can be used Magnifying glasses are valuable aids for determining whether or not all corrosion products have been removed during cleanup operations Liquid Dye Penetrant Inspection a Inspection for large stress corrosion or corrosion fatigue cracks on nonporous or nonferrous metals may be accomplished using dye penetra
74. at and bring to a high polish by rubbing lightly with a soft cloth Cleaning Plastic A Thefollowing procedure is the recommended method for cleaning plastic windows 1 Clean the plastic by washing with plenty of water and mild soap using a clean soft grit free cloth sponge or bare hands CAUTION Do not use gasoline alcohol benzene acetone carbon tetrachloride fire extinguisher or deicing fluids lacquer thinners or window cleaning sprays because they will soften the plastic and cause crazing 56 10 00 Page 804 Cessna Aircraft Company Jun 1 2005 WINGS CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT PAGE 57 Title 57 List of Effective Pages 57 Record of Temporary Revisions 57 Table of Contents 57 00 00 Pages 1 4 57 10 00 Pages 1 4 57 11 00 Pages 801 802 57 12 00 Pages 801 803 57 20 00 Pages 801 803 57 21 00 Pages 801 805 57 22 00 Pages 601 606 57 23 00 Page 801 57 24 00 Pages 801 802 57 25 00 Pages 801 802 57 26 00 Pages 801 802 57 27 00 Pages 801 802 57 40 00 Page 801 57 50 00 Page 801 DATE Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 57 LIST OF EFFECTIVE PAGES Cessna Aircraft Company Page 1 of 1 Jun 1 2005 RECORD TEMPORARY REVISIONS Temporary Revision Number Pa
75. ave no evidence of tears cracks or penetrations which are not stress wrinkles and do not change Oil can or pop in and out with internal pressure are considered negligible damage Repairable damage Dents or bends in the wing fuel spar and ribs may be repaired by reforming or by replacing a section of the structure Since aluminum work hardens it is much more likely to crack when reformed and should be carefully inspected for such cracks after rework Removal and replacement of a damaged section is preferred to reformation Damage Necessitating Replacement Of Parts Due to the amount of fuel bay sealant which must be removed from fuel bay components to facilitate repair individual parts are not available to replace fuel bay spars or ribs The entire fuel bay area must be replaced as a unit Wing Leading Edge Damage Criteria A Negligible damage Any smooth dents in the wing leading edge skin that are not more than 0 030 inch 0 76 mm below contour and circumscribable with not more than a 1 5 inch 38 mm diameter circle that has no evidence of skin tears cracks or skin penetrations which are not stress wrinkles and do not interfere with internal structure constitute negligible damage However because of the critical nature of the wing leading edge this cosmetic repair should be completed Repairable damage Dents or dings deeper and or larger than specified above must be repaired Skin tears cracks or penetrations must be repai
76. bearing area 3 Holes must not be elongated and the Hi Shear substituted must be a smooth push fit B Field replacement of main landing gear forgings on bulkheads may be accomplished by using the following hardware 1 NAS464P NAS436P and either NAS1103 through NAS1120 NAS1303 through NAS623 or NAS6203 through NAS6220 bolt and either a 521042 nut and AN960 NAS1149 washers in place of Hi Shear rivets for forgings with machined flat surfaces around the attachment holes b ESNA2935 mating base washer and ESNA RM52LH2935 self aligning nut with forgings with a draft angle of up to a maximum of eight degrees without machined flat surfaces around the attachment holes Substitution Of Rivets A When adapting the typical repairs shown in this manual to suit actual conditions it may be necessary to use different fasteners than those originally used This may be due to non availability of a particular fastener restricted access or other difficulties When replacing rivets it is desirable to use rivets identical to the type of rivet removed Countersunk head rivets are to be replaced by rivets of the same type and degree of countersink When rivet holes become enlarged deformed or otherwise damaged several options are available 1 The simplest solution is to install a 1 32 inch 0 032 inch larger size rivet as a replacement This solution uses the designed repairability of the structure and is the quickest repair 2 Repair rivet
77. ced 51 60 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 154 5180002 12 weight assembly washer and bolt 5180002 5 sliding weight N 5180002 2 beam assembly mandrels NOTE Included in 5180002 1 flight control surface balancing fixture kit 6280T2006 Flight Control Surface Balancing Fixture Kit Figure 1 Sheet 1 51 60 00 Page 2 Cessna Aircraft Company Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 155 Centerline beam must be aligned with control Beam assembly surface hinge centerline Conirol surface chord line Hanger Hinge centerline assembly Add washers as necessary to fine balance the beam assembly e Adjustable weight Hanger assembly Mandrel to be in proper position Read control surface moment at center Sliding weight of weight Beam assembly Control surface chord line Mancrel Flat surface 2682T1022 Balancing Control Surfaces Figure 2 Sheet 1 51 60 00 Page 3 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 156 A balance in this range is underbalance A balance in this range is overbalance Balancing mandrel Rudder E 1 LJ LJ 90
78. ck that does not extend past the stop drilled hole may remain in service without additional repair 7 Refer to Figure 802 as applicable for repair information C Corrugated Elevator Skin Repair 1 It is permissible to stop drill crack s that originate at the trailing edge of the control surface provided the crack s is are not more than 2 inches in length 2 Stop dill crack s using a Number 30 0 128 inch diameter drill 3 crack may only be stop dilled once NOTE A crack that passes through a trailing edge rivet and does not extend to the trailing edge of the skin may be stop drilled at both ends of the crack 51 75 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Patch to be same thickness as skin Doubler to be one gage heavier Both 2024 T3 or stronger B162 Doubler Patch Doubler Existing skin view A A Original parts W HZ Repair parts N Repair parts in cross section Carry existing rivet pattern through patch Skin Edge margin is Use flush rivets in vicinity of flush and protruding Flush patch at equal to 2 times rivet diameter RIVET TABLE RIVET SPACING GAGE DIAMETER MS20426AD MS20470AD SKIN RIVET 1 16 NOTE 0 56 0 82 NOTE 0 94 NOTE 0 82 0 90 NOTE 0 70 0 96 NOTE 0 58 0 84 NOTE 0 96 NOTE 0 94 NOTE 0
79. cordance with procedures outlined in Chapter 51 Flight Control Surface Balancing Damage Criteria A Damage to the flaps and ailerons can be divided into three major categories and is detailed in Wing Damage Classification Flap and Aileron Repair A Skin damage exceeding that considered negligible that can be repaired with minor patches can be considered repairable Flush skin patches are illustrated in Chapter 51 Typical Skin Repairs Figure 801 A typical rib repair is illustrated in Chapter 51 Control Surface Repair Figure 801 trailing edge repair in Chapter 51 Control Surface Repair Figure 802 are typical flap and aileron repairs B Flight control surfaces which have been repaired or replaced must be balanced in accordance with procedures outlined in Chapter 51 Flight Control Surface Balancing 57 40 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL WING LIFT STRUTS 1 General A wing lift struts consist of 6061 T6 tube stock formed into an aerodynamic shape Attach fittings are machined from 7075 T73 bar stock and attached to the strut tubes 2 Wing Strut Damage Classification A Damage to the wing lift strut can be divided into three major categories and is detailed in Wing Damage Classification 57 50 00 Page 801 Cessna Aircraft Company Jun 1 2005 POWER PLANT CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTU
80. curs between WS 100 50 View looking down wing and the tip rib 268271041 Measuring Model 172 Series Wing Twist Figure 601 Sheet 1 57 22 00 603 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 194 UT LINZ Bolt A Bolt Bolt A Bolt dimension dimension dimension Wing station inches inches inches or B 2 00 1 00 29 50 39 00 2 00 1 00 29 50 100 50 0 45 1 00 24 00 207 00 dh FWD Wing tip WS 207 00 WS 100 50 WS 39 00 NOTE All wing twist occurs between WS 100 50 gt 3 View looking down wing and the tip rib 2682T1041 Measuring Model 182 Series Wing Twist Figure 602 Sheet 1 57 22 00 Page 604 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 2 At WS 100 00 a Hold straightedge parallel to wing station b Place bolt A on mark set protractor head against lower edge of straightedge and verify bubble in protractor head indicates level 3 At WS 207 00 a Remove bolt A from straightedge Grind another bolt A to a dimension of 0 66 inch Place this bolt 20 00 inches from bolt B and secure to straightedge b Hold straightedge parallel to
81. d pattern as the manufactured seams at the edges of the original sheet If the manufactured seams are different the stronger must be copied If the repair ends at a structural member where no seam is used enough repair panel must be used to allow an extra row of staggered rivets with sufficient edge margin to be installed Typical methods of repair for skins bulkheads stringers and channels are illustrated in Chapter 51 Typical Skin Repairs Before repairs are attempted all cracks or deep scratches must be stop drilled with a No 30 0 128 inch drill and all sharp corners and ragged edges must be trimmed away and deburred 53 10 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 4 Replacement Damage A All forgings and castings of any material and structural parts made of steel must be replaced if damaged Structural members of a complicated nature that have been distorted or wrenched should be replaced Seat rails serve as structural parts of the fuselage and must be replaced if damaged 53 10 00 Page 2 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CABIN BULKHEAD REPAIR General A Bulkheads are comprised of formed C channel sections The principal material of construction is 2024 0 Alclad aluminum alloy which after forming is heat treated to a 2024 T42 condition All bulkheads in the fuselage a
82. de for a larger drill on top of the rivet stem and drill away the tapered portion of the stem to destroy the lock Pry the remainder of the locking collar out of the rivet head with a drift punch 51 40 00 Page 7 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Drill nearly through the head of the rivet using a drill the same size as the rivet shank Break off rivet head using drift pin inserted into the drilled hole as a pry Drive out remaining rivet shank with a pin having a diameter equal to the rivet shank S 8 4 5 Installation of Blind Rivets 1 Refer to Figure 2 for an illustration of installation procedures 2 Check that rivet hole size and rivet are compatible 3 Check that proper pulling head is installed on rivet gun 4 Adjustment of pulling head must be made in accordance with manufacturers instructions 5 Check that proper operating air pressure is available to rivet gun NOTE Blind rivets may be installed using pneumatic or mechanical guns whichever is available Check that holes in parts to be fastened are properly aligned In blind clearance applications check the minimum blind clearance BK dimension if the manufactured head of blind rivet is protruding above the top sheet The rivet will pull down the sheet as the stem is pulled if the BK dimension is met or exceeded 8 The minimum blind clearance is the BK dimension and is listed in
83. e 57 10 00 Page 2 Wing Strut Damage 57 10 00 Page 3 Aileron Damage Criteria Corrugated Skin 57 10 00 Page 3 Aileron Damage Criteria Model 206 Aileron 57 10 00 Page 3 Wing Flap Damage Criteria Corrugated Skin 57 10 00 Page 4 Wing Flap Damage Criteria Model 206 57 10 00 Page 4 WING FUEL BAY 5 57 11 00 Page 801 Preparing Damaged Area In Wing Fuel Bay for Repair 57 11 00 Page 801 FUEL BAY SEALING DURING STRUCTURAL 57 12 00 Page 801 General oeste Eee Ves Sed E cas TR e e a a e e iH 57 12 00 Page 801 Integral Fuel Bay 57 12 00 Page 801 Mixing Sealant sec ret mut Ree Gece Mareen LI RES ROSA RETE D 57 12 00 Page 801 Applying Sealant dae ed 57 12 00 801 Sealing Fuel 57 12 00 802 gt 57 12 00 802 Testing Integral Fuel 57 12 00 Page 802
84. e page numbering system consists of the Air Transport Association Specification 100 three element numbers separated by dashes The page number and date are printed immediately to the right of the three element number The three element number is assigned to a component with the first set of numbers corresponding to the ATA 100 assigned chapter number 2 The page number complies with Air Transport Association Specification 100 for subdividing a Structural Repair Manual Blocks of sequential page numbers are used to identify Pages 1 Through 100 Structural Identification Pages 101 Through 199 Allowable Damage Pages 201 through 999 Repair Procedures 3 date which appears below the page number signifies when the page was issued If no revisions to that page have occurred the date signifies original date 4 Illustrations use the same figure numbering as the page block in which they appear For example Figure 202 would be the second figure in a repair procedure 6 Revision Manual A Regular Revision 1 Pages to be removed or inserted in the Structural Repair Manual are controlled by the Effectivity Page Pages are listed in sequence by the three element number and then by page number When two pages display the same three element number and page number the page displaying the most recent Date of Page Issue shall be inserted into the Structural Repair Manual The date column on the corresponding chapter Effectivity Page shall ver
85. e repaired Skin tears cracks or penetrations must be repaired Dings that include understructure ribs must be repaired by reforming or removal and replacement of the rib Revaluation of the skin after the repair of the understructure will determine if the skin damage is negligible repairable or replacement damage Special care must be taken to minimize added weight since the surface must be rebalanced after rework Damage Necessitating Replacement Of Parts Because of the balance requirements multiple areas of damage may require replacement of skins to allow balance limits to be attained 57 10 00 Page3 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 13 Wing Flap Damage Criteria Corrugated Skin Flap A Negligible damage Any smooth dents in the flap skin that are not more than 0 050 inch 1 27 mm below contour and circumscribable with not more than a 1 5 inch 88 1 mm diameter circle that have no evidence of skin tears cracks or skin penetrations and which do not include corrugations constitute negligible damage Repairable damage Dents or dings deeper and or larger than specified may be repaired Skin tears cracks or penetration must be repaired Dings that include corrugations are unlikely to be reworkable but may be repaired by replacing the damaged area Corrugated skin material is available from Cessna Damage Necessitating Replacement Of Parts Multiple re
86. e so low below 40 percent that sealant curing will be retarded If necessary the relative humidity may be increased by the use of water containing less than 100 parts per million total solids and less the 10 parts per million chlorides 7 Testing Integral Fuel Bay A The fuel system consists of two vented integral fuel tanks one in each wing The following procedures are for testing integral fuel bay 1 Remove vent line from vent fitting and cap fitting 2 Disconnect fuel lines from bay 3 To one of the bay fittings attach a water manometer capable of measuring 20 inches of water 4 To the other bay fitting connect a well regulated supply of air 1 2 psi maximum or 13 8 inches of water Nitrogen may be used where the bay might be exposed to temperature changes while testing 5 Make sure filler cap is installed and sealed CAUTION Do not attempt to apply pressure to the bay without a good regulator and a positive shutoff in the supply line Do not pressurize the fuel bay to more than one half psi or damage may occur Apply pressure slowly until one half psi is obtained Apply a soap solution as required Allow 15 to 30 minutes for pressure to stabilize If bay holds for 15 minutes without pressure loss bay is acceptable 57 12 00 Page 802 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 10 Reseal and retest if any leaks are found 57 1
87. ead flathead countersunk head and universal head Flathead rivets are generally used in the airplane interior where head clearance is required MS20426 countersunk head rivets are used on the exterior surfaces of the airplane to minimize turbulent airflow MS20470 universal head rivets are used on the exterior surfaces of the airplane where strength requirements necessitate a stronger rivet head than that of the countersunk head rivet Hi Shear rivets are special patented rivets having a high shear strength equivalent to that of standard NAS bolts They are used in special cases in locations where high shear loads are present such as in spars wings and in heavy bulkhead ribs This rivet consists of a cadmium plated pin of alloy steel Some have a collar of aluminum alloy Some of these rivets can be readily identified by the presence of the attached collar in place of the formed head on standard rivets Blind rivets are used where strength requirements permit where one side of the structure is inaccessible making it impossible or impractical to drive standard solid shank rivets Replacement Of Hi Shear Rivets A Replacement of Hi Shear rivets with close tolerance bolts or other commercial fasteners of equivalent strength properties is permissible 1 hardware used for the Hi Shear rivets is determined according to the size of the holes and the grip lengths required 2 Bolt grip length should be chosen so that no threads remain in the
88. epair parts in cross section Original parts Edge margin is equal to 2 times rivet diameter same gage as skin Patch lt o z ao a hole size 7000 A C 05 UC m x PE 5 B165 2682T1032 Jun 1 2005 51 75 00 Page 805 Skin Repair Figure 801 Sheet 4 Cessna Aircraft Company viEW Section through assembled patch B CESSNA AIRCRAFT COMPANY INGLE ENGINE STRUCTURAL REPAIR MANUAL Edge margin equal to 2 times rivet diameter B166 See sheet 1 for rivet spacing Patch 2024 T42 same gage as skin Overlapping rectangular patch Clean out damaged area equal to 2 times rivet diameter 0 50 inch patch Edge margin equal to 2 times On firewall sheet repair rivet diameter use MIL S 5059 corrosion resistant steel and MS204 rivets Doubler Doubler 2024 T42 same gage as skin B hole size 2 4 Original parts 242 B Repair parts ZZ 22 Repair parts EM 0 25 B cross section view Section through assembled patch 268211033 Skin Repair Figure 801 Sheet 5 51 75 00 Page 806 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B167 0 25 inch 6 4 mm minimum edge margin Use existing rivet pattern and rivet size damaged Patch may overlap or be inserted under existing
89. es will display the current revision date in the Date of Page Issue location When extensive technical changes are made to text in an existing section that requires extensive revision revision bars will appear the full length of text When art is revised or added a change bar will appear on the full length of the page INTRODUCTION Page3 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LIST OF REVISIONS 1 General A This Structural Repair Manual includes the original issue and the following listed revisions To make sure that information in this manual is current and the latest maintenance and inspections procedures are available revisions must be incorporated in the manual as they are issued Table 1 Original Issue 2 December 1996 Revision Number Date Writer Revision Number Date Writer 1 16 May 1997 2 16 July 1999 3 15 January 2001 4 1 June 2005 jmk LIST OF REVISIONS Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LIST OF CHAPTERS CHAPTER Jun 1 2005 FICHE FRAME 51 Standard Practices Structures Jun 1 2005 1 A10 52 Doors Jun 1 2005 1 D2 53 Fuselage Jun 1 2005 1 D8 55 Stabilizers Jun 1 2005 1 E7 56 Windows Jun 1 2005 1 E16 57 Wings Jun 1 2005 1 F2 71 Powerplant Jun 1 2005 1 H2 NOTE 1 Represents date of page one of each chapter s List of Effective Pages which is applicable to Manual revisi
90. esult of fatigue failure 8 Corrosion Corrosion due to a complex electrochemical action is a damaged area of any size and depth which results in a cross sectional area change Depth of such pitting damage must be determined by a cleanup operation Damage of this type may occur on surfaces of structural elements Refer to Corrosion and Corrosion Control Section 51 11 00 Use good sense and proper visual measurement in the determination of significant cross sectional area changes of both depth and length of any type or combinations of damage mentioned above 2 Damage Investigation A After a thorough cleaning of the damaged area all structural parts should be carefully examined to determine the extent of damage Frequently the force causing the initial damage is transmitted from one member to the next causing strains and distortions Abnormal stresses incurred by shock or impact forces on a rib bulkhead or similar structure may be transmitted to the extremity of the structural member resulting in secondary damage such as sheared or stretched rivets elongated bolt holes or canned skins or bulkheads Points of attachment should be examined carefully for distortion and security of fastenings in the primary and secondary damaged areas at locations beyond the local 51 10 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL damage This is particularly true with wing
91. et diameter Clean out damaged area Doubler stainless steel same gage as firewall Edge margin equal to 2 times rivet diameter Edge margin equal to 2 times rivet diameter Crack repair method Rivet spacing equal to 6 times rivet diameter Stop drill hole diameter 0 09 B inch minimum B Hole diameter cm 0 25B Seal with AMS 3374 Seal with AMS 3374 Type Ill firewall sealant Type Ill firewall sealant Monel rivets typical view A A view B B SECTION THROUGH ASSEMBLED PATCH SECTION THROUGH ASSEMBLED PATCH 2682T1048 Typical Firewall Repair Figure 801 Sheet 1 53 40 00 Page 802 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B182 0 25 inch edge margin Clean out damaged area 9 Angle 2024 4 Alclad 10 rivets each side Firewall angle Filler 2024 T4 Alclad Firewall MS20470AD4 rivets Fuselage skin Eee Original parts Repair parts view A A Repair parts cross section 0582T1004 Firewall Angle Repair Figure 802 Sheet 1 53 40 00 Page 803 Cessna Aircraft Company Jun 1 2005 STABILIZERS CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT 55
92. ets along the rear spar and ribs and springing back the skin By using the proper bucking bars through holes in spar web skins may by closed with a minimum of blind rivets Replacement Damage A If the damaged area would require a repair which could not be made between adjacent ribs or the repair would be located in an area with compound curves compete skin panels must be replaced Ribs and spars may be repaired but replacement is generally preferable Where damage is extensive replacement of the entire assembly is recommended 55 10 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL ELEVATOR General A The elevator assembly consists of a left and right section bolted together near the airplane centerline by torque tubes Each section consists of a front and a rear spar ribs skins and a trim tab assembly A balance weight is bolted to the outboard tip leading edge Negligible Damage A Any smooth skin dents that are free from cracks abrasions and sharp corners and which are not stress wrinkles and do not interfere with any internal structure or mechanism may be considered as negligible damage Exception to negligible damage on elevator surfaces is the front spar cracks appearing in web of hinge fitting or in tip rib which supports overhanging balance weight Cracks in overhanging tip rib in the area at the front spar intersection with web of the rib
93. f passive materials such as alloys of aluminum nickel and chromium It is first noticeable as a white or gray powdery deposit similar to dust which blotches the surface When the deposits are cleaned away tiny pits can be seen in the surface Dissimilar Metal Corrosion 1 When two dissimilar metals are in contact and are connected by an electrolyte continuous liquid or gas path accelerated corrosion of one of the metals occurs The most easily oxidized surface becomes the anode and corrodes The less active member of the couple becomes the cathode 51 11 00 Page Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B149 C Continuous liquid path electrolyte Current flow gt Cathodic area b Electron conductor metal Electron fow Simplified Corrosion Cell Unbroken paint film No contact between electrolyte and anode and cathode Continuous liquid path electrolyte Electron conductor metal Corrosion eliminated by application of organic film 26821017 Corrosion Identification Figure 1 Sheet 1 51 11 00 Page 2 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL of the galvanic cell The degree of attack depends on the relative activity of the two surfaces the greater the difference in activity the more
94. ge Number Issue Date By Date Removed CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CONTENTS WINGS ea 57 00 00 1 Gerieral coxa A toe RAUS ana Mega ERE E wea ee EA 57 00 00 Page 1 Tools Equipment and Materials 57 00 00 Page 1 Installation of Access 57 00 00 Page 1 WING DAMAGE 57 10 00 Page 1 Damage 57 10 00 Page 1 Wing Skin Damage Criteria 57 10 00 Page 1 Wing Stringer Damage 57 10 00 Page 1 Wing Auxiliary Spar Damage Criteria 57 10 00 Page 1 Wing Rib Damage Criteria 57 10 00 Page 1 Wing Spar Damage 57 10 00 Page 2 Wing Fuel Bay Spars Rib Damage Criteria 57 10 00 Page 2 Wing Leading Edge Damage Criteria 57 10 00 Page 2 Bonded Leading Edge Damag
95. gure 801 Sheet 1 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL BONDED LEADING EDGE REPAIR General A Bonded leading edges are used on the Model 182 and Model 206 T206 series of airplanes The following repairs apply to these airplanes only Bonded Leading Edge Damage Classification A Damage to the bonded leading edge can be divided into three major categories and is detailed in Wing Damage Classification Bonded Leading Edge Repair A Repairs to the bonded leading edge are illustrated in Figure 801 57 26 00 Page 801 Cessna Aircraft Company Jun 1 2005 197 1 2 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Use rivet pattern at wing stations 23 65 for repair from wing station 23 65 to wing station 85 87 Use rivet pattern at wing station 100 50 for lap splice patterns from wing station 100 50 to wing station 190 00 Use rivet spacing similar to the pattern at wing station 100 50 at leading edge ribs between Number of flush rivets in dimpled skin lap splices required in replacement leading edge skin Select number of flush rivets to be Wing Solid Blind used at each wing station leading edge rib from table SOR MS20426 4 CR2248 4 Ribs and stringers Blind rivets may be substituted for solid rivets in proportionally increased numbers in accordance with the table Spars Blind rivets may be installed i
96. hat is longer than 2 inches 50 mm 2 Cracks in more than 10 percent of the attach fastener locations per fairing 2 Fairings with a stop drilled crack that does not extend past the stop drilled hole may remain in service until the next 100 hour or equivalent inspection 51 73 01 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL TYPICAL SKIN REPAIRS 1 General A Damage which would involve a typical skin repair can be described as damage that requires modification such as material replacement or patching Skin damage in the form of dents scratches or punctures requires a patch Refer to Figure 801 for an illustration of typical skin repairs Refer to Figure 802 for corrugated skin repairs 2 Guidelines for Corrugated Skin Crack Repairs A Corrugated Aileron Skin Repair 1 It is permissible to stop drill crack s that originate at the trailing edge of the control surface provided the crack s is are not more than 2 inches in length 2 Stop dill crack s using a Number 30 0 128 inch diameter drill 3 crack may only be stop dilled once NOTE A crack that passes through a trailing edge rivet and does not extend to the trailing edge of the skin may be stop drilled at both ends of the crack 4 Any control surface that has a crack that progresses past a stop drilled hole shall be repaired or replaced 5 A control surface that has any of the follo
97. heir small size cowl reinforcement angles should be replaced rather than repaired if they become damaged 71 10 00 Page 801 Cessna Aircraft Company Jun 1 2005 1 General A CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL DYNAFOCAL TYPE ENGINE MOUNT REPAIRS The engine mount is fabricated from 4130 chrome molybdenum steel tubing The mount attaches to the firewall at four points and to the engine using rubber isolation mounts at four points NOTE Repair by gas welding is acceptable 2 Engine Mount Repairs A The following procedures are to be used when making repairs to the engine mount Refer to Figure 801 1 All welding on the engine mounts should be of the highest quality since the tendency of vibration will accentuate any minor defect present and cause fatigue cracks Engine mount members are preferably repaired by using larger diameter replacement tube welds However reinforced 30 degree scarf welds in place of the fishmouth welds are considered satisfactory for engine mount repair work Minor damage such as a crack adjacent to an engine attaching lug may be repaired by rewelding the tube and extending a gusset past the damaged area Extensively damaged parts must be replaced Engine mounting lugs and engine mount to fuselage attach fittings should be replaced not repaired For information on damage beyond the scope of these repairs consult Cessna Propeller Aircraft Product Support
98. ial and splicing in a new section The splice must be lapped over the old material sealed and secured with steel rivets 1 Patches splices and joints must be repaired using MS20450 steel rivets B Following any repair to the firewall assembly seal the damaged areas as follows 1 Clean area on surface to be sealed with methyl propyl ketone 2 Mix one part of catalyst thoroughly with 100 parts of Pro Seal No 700 base NOTE Sealant should be mixed by weight It is important that accelerator be completely and uniformly dispersed throughout the base compound 3 Using a spatula caulking gun or flow gun apply a fillet of sealer along cracks seams joints and rows of rivets NOTE Ifthe sealant is applied before the parts are mated use enough sealing compound to completely fill the joint and wipe away excess after parts are mated NOTE Ifthe sealant is applied with a brush or a brush flow gun more than one coat of sealant will be necessary on very porous material Sealant should be allowed to air dry 10 minutes between coats 53 40 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B181 2 Patch and doublers Original parts 0 50 inch 302 stainless steel on firewall 2 Repair parts al Repair parts in cross section Edge margin equal to 2 times rivet diameter Hole repair method Rivet spacing equal to 6 times riv
99. ify the active page Temporary Revision 1 For paper publications a Temporary revision pages are filed in the Structural Repair Manual by replacing existing pages in the manual File the temporary revision cover page according to the filing instructions on the Temporary Revision Cover Page 2 For aerofiche publications a Draw a line through any aerofiche frame page affected by the Temporary Revision with a permanent red ink marker This will be a visual identifier that the information on the frame page is no longer valid and the Temporary Revision should be referenced for added pages in a temporary Revision draw a vertical line between the applicable frames which is wide enough to show on the edges of the pages Temporary Revisions should be collected and maintained in a notebook or binder near the aerofiche library for quick reference INTRODUCTION 2 Cessna Aircraft Company Jun 1 2005 7 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Identifying Revised Material A B Additions or revisions to text in an existing section will be identified by a revision bar in the left margin of the page and adjacent to the change When additions or revisions are made to text in an existing section all pages displaying the same three element number shall also display the same Date of Page Issue The date column on the corresponding chapter Effectivity Page shall verify the active page These pag
100. inal parts in cross section 268211044 Typical Stringer and Channel Repair Figure 801 Sheet 1 Page 802 Jun 1 2005 53 30 00 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL SY Repair parts in cross section ve NV 2 Ky 25 AI S i P 2 Ay s WW 7 E 0885 e lo e re 7 ce 27 2 5 2 o amp ED E amp E 5 8 2682T1045 Jun 1 2005 view 53 30 00 Page 803 Typical Stringer and Channel Repair Figure 801 Sheet 2 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Doubler 2024 T42 Same gage as channel Two rows of rivets outboard of lightening holes AN NS _ y 5 Stop drill hole diameter 0 09 inch minimum J 2 o A af 5 2 x 22 Repair parts VIEW A A 268211046 Skin Repair parts in cross section Typical Stringer and Channel Repair Figure 801 Sheet 3 Jun 1 2005 53 30 00 Page 804 Aircraft Company Cessna CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Clean out damaged area Two rows of rivets outboard of lightening holes Filler
101. ip pen approximately 0 50inch aft of lateral row of rivets in wing leading edge spar flange 3 Locate outboard WS 207 00 Make a mark with a felt tip pen approximately 0 50inch aft of lateral row of rivets in wing leading edge spar flange Measure Wing Twist at Each Wing Station Refer to Figure 603 NOTE While performing the following procedure stay as far away as possible from the canned areas of the wing 1 At WS 39 00 a bolt A to a dimension of 2 00 inches Grind bolt B to a dimension of 1 00 inch Place these bolts 29 50 inches from each other on the upper edge of the straightedge and secure using tape b Secure protractor to bottom of the straightedge c Hold straightedge parallel to wing station and place bolt A on mark d Set bubble in protractor to center and lock protractor to hold this reading 57 22 00 Page 602 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B193 T LINZ Bolt A Bolt B Bolt A Bolt dimension dimension dimension Wing station inches inches inches or B 2 00 1 00 29 50 39 00 2 00 1 00 29 50 100 50 0 45 1 00 24 00 207 00 FWD Wing tip WS 207 00 WS 100 50 WS 39 00 NOTE All wing twist oc
102. ized Service Stations or through the Cessna Product Support subscription services Cessna Service Stations have also been supplied with a group of supplier publications which provide disassembly overhaul and parts breakdowns for some of the various supplier issued revisions and service information which may be reissued by Cessna s Authorized Service Stations and or through Cessna s subscription services WARNING All inspection intervals replacement time limits overhaul time limits the method of inspection life limits cycle limits etc recommended by Cessna are solely based on the use of new remanufactured or overhauled Cessna approved parts If parts are designed manufactured remanufactured overhauled purchased and or approved by entities other than Cessna then the data in Cessna s maintenance service manuals and parts catalogs are no longer applicable and the purchaser is warned not to rely on such data for non Cessna parts All inspection intervals replacement time limits overhaul time limits the method of inspection life limits cycle limits etc for such non Cessna parts must be obtained from the manufacturer and or seller of such non Cessna parts Coverage A Cessna Single Engine Structural Repair Manual is prepared in accordance with the Air Transport Association Specification 2200 for Manufacturers Technical Data B Structural Repair Manual contains material identification for structure subject t
103. justified cutting away the damaged portion and inserting a trimmed portion of the original section adequately reinforced by splice plates or doublers will prove satisfactory This is knownas an insertion type patch Landing Gear Bulkheads A Landing gear bulkheads are highly stressed members irregularly formed to provide clearance for control cables fuel and brake lines Patch type repairs on these bulkheads are for the most part impractical Minor damage consisting of small nicks or scratches may be repaired by dressing out the damaged area or by replacement of fasteners Any other damage must be repaired by replacing the landing gear support assembly as an aligned unit Repair After Hard Landing A Buckled skin or floor boards and loose or sheared rivets in the area of the main gear support are indications of damage to structure from an extremely hard landing When such evidence is present the entire support structure must be examined and all support forgings must be checked for cracks 1 Use fluorescent dye penetrant and magnification to examine for cracks B Bulkheads in the damaged area must be checked for alignment Deformation of bulkhead webs must be checked using a straightedge C Damaged support structure buckled floorboards and skins and damaged or questionable forgings must be replaced 53 20 00 Page 801 Cessna Aircraft Company Jun 1 2005 view CESSNA AIRCRAFT COMPANY SIN
104. k set protractor head against lower edge of straightedge and verify bubble in protractor head indicates level At WS 207 00 a Remove bolt A from straightedge Grind another bolt A to a dimension of 0 45 inch Place this bolt 24 00 inches from bolt B and secure to straightedge b Hold straightedge parallel to wing station and place bolt A on mark c Check to assure that protractor bubble is still centered If proper twist is present the protractor readings will be the same parallel NOTE Forward or aft bolt may be lowered from wing 0 10 inch maximum to attain level indication 57 22 00 Page 601 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 4 Model 182 Series Wing Twist Check Procedure A Mark Wing Station Reference Points Refer to Figure 602 1 Locate WS 39 00 Make a mark with a felt tip pen approximately 0 50inch aft of lateral row of rivets in wing leading edge spar flange 2 Locate WS 100 50 Make a mark with a felt tip pen approximately 0 50 inch aft of lateral row of rivets in wing leading edge spar flange 8 Locate outboard WS 207 00 Make a mark with a felt tip pen approximately 0 50 inch aft of lateral row of rivets in wing leading edge spar flange Measure Wing Twist at Each Wing Station Refer to Figure 602 NOTE While performing the following procedure stay as far away as possible from the canned areas of the
105. kness 51 11 00 Page4 Cessna Aircraft Company Jun 1 2005 4 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Severe Corrosion a General appearance may be similar to moderate corrosion with severe blistering exfoliation and scaling or flaking Pitting depths may be as deep as 15 percent of the material thickness This type of damage is normally repaired by complete part replacement but patches or other types of repair may be available Contact Cessna Propeller Aircraft Product Support P O Box 7706 Wichita KS 67277 USA for assistance 6 Corrosion Removal A following methods are provided as an aid in determining the correct method for corrosion removal Standard Methods 1 2 3 Several standard methods are available for corrosion removal The method normally used to remove corrosion are chemical treatments hand sanding with aluminum oxide or metal wool that is of similar material to the surface being treated and mechanical sanding or buffing with abrasive mats or grinding mats The method used depends on the metal and the degree of corrosion Select appropriate materials from the abrasives chart as illustrated in Figure 2 Aluminum and Aluminum Alloys Most formed aluminum parts and skins of this airplane consist of various gauges of sheet 2024 T3 and 2024 T42 Alclad Alclad is formed by laminating a thin layer of relatively pure aluminum one to five mils thick
106. l of damaged area leave edges parallel to any square or rectangular edge of the unit NOTE Damage adjacent to a previous repair requires removal of old repair and inclusion of the entire area in the new repair Round all corners Smooth out abrasions Vacuum thoroughly to remove all chips filings dirt etc from bay area All surfaces to be sealed after repair should be thoroughly cleaned by wiping with a clean cloth dampened with methyl propyl ketone acetone or similar solvent and dried with a clean cloth before allowing solvent to evaporate Always pour the solvent on the cloth to prevent contaminating solvent Do not allow cloth to drip Never use contaminated solvent Any repair that breaks the fuel bay seal will require resealing that bay area refer to applicable Maintenance Manual Chapter 28 Fuel Tank Sealing Maintenance Practices for sealing materials and procedures 57 11 00 Page 802 Cessna Aircraft Company Jun 1 2005 1 2 3 4 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL FUEL BAY SEALING DURING STRUCTURAL REPAIR General A Any repair that breaks the fuel bay seal will necessitate resealing that bay area Repair parts that need sealing must be installed during the sealing operations All joints within the boundary of the bay but which do not provide a direct fuel path out of the bay such as fuel spar flanges and rib flanges must be fay surface sealed and fillet
107. l repair practices materials and procedures which are applicable to the Horizontal and Vertical Stabilizers horizontal and vertical stabilizers are of all metal fully cantilever semimonocoque design consisting of spars stringers ribs and skins Skins are riveted to supporting structure with conventional MS20470AD rivets C If questions arise concerning approved repairs or for repairs not shown in this section contact Cessna Propeller Aircraft Product Support 55 00 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL HORIZONTAL STABILIZER Horizontal Stabilizer A horizontal stabilizer is constructed from spars ribs stringers doublers and skins Refer to applicable Maintenance Manual Chapter 6 Dimensions and Areas for horizontal stabilizer station diagram Negligible Damage A same criteria which is used to define negligible damage to the fuselage may be applied to the horizontal stabilizer Refer to Chapter 53 Fuselage Damage Classification for a complete description of negligible damage Repairable Damage A Skin patches may be used to repair skin damage These patches are illustrated in Chapter 51 Typical Skin Repairs Figure 801 For skin damage which includes corrugations Refer to Chapter 51 Typical Skin Repairs Figure 802 B Access to the internal stabilizer structure may be gained by removing a portion of the riv
108. ll aluminum parts having rough edges with a solution of Iridite or alodine mixed in a ratio of one ounce of lridite or alodine to one gallon of water and rinse thoroughly To restore original paint and corrosion protectant properties to factory standards refer to appropriate Maintenance Manual Chapter 20 Exterior Finish Cleaning Painting for refinishing procedures and required materials NOTE Damage adjacent to a previous repair requires removal of the old repair and inclusion of the entire area in the new repair 51 71 00 Page 801 Cessna Aircraft Company Jun 1 2005 3 General CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL GLASS FABRIC REPAIR A following procedures are for parts which are constructed of epoxy prepreg glass fabric Tools and Materials NOTE Equivalent substitutes may be used for the following NAME Fiberglass Polyethylene sheet Adhesive Adhesive Adhesive Methyl Propyl Ketone Sandpaper Rubber sheet A procedures listed below are for repairing of glass fabric parts Refer to Figure 801 for an illustration NUMBER 181 weight EA9394 EA9396 Epon 815 Various grits Repair Of Glass Fabric Parts of a typical glass fabric repair 1 Cut and trim area immediately beyond damage If parts were painted remove paint and sand MANUFACTURER Hexcel Commercially available Loctite Aerospace Bay Point CA 94565 Loctite Aerospace Loctite Ae
109. lustrated Parts Catalogs do not identify the standard shape from which parts are fabricated Detailed measurements of damaged areas are required to determine the standard section from which parts are fabricated 51 30 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B2095 E A E asl 69 f acct EXTRUDED ANGLE ORIGINAL MATERIAL SUBSTITUTION STD 2024 SHEET SHAPE MATERIAL B D R2 5 5 85 2024 3511 0 625 0625 0 045 0 045 0060 0 045 0045 0 625 0 625 0040 0 050 5 97 2024 3511 0875 0875 0063 0 063 0 094 0 063 0 063 0875 0875 0 063 0 125 5 107 2024 3511 1 000 0 750 0 062 0 062 0 060 0031 0031 1 000 0 750 0 063 0 075 5 111 2024 3511 0750 0750 0 094 0 094 0 125 0 094 0094 0 750 0750 0 090 0 125 5 125 2024 3511 0750 0625 0 062 0 062 0062 0 031 0 031 0 750 0 625 0 063 0 100 174 2024 13511 1 000 1 000 0 062 0 062 0 062 0 031 0 031 1 000 1 000 0 063 0 100 5 376 2024 3511 0875 0 750 0063 0063 0 125 0 063 0063 0875 0 750 0 063 0 100 8 1111 2024 3511 0750 0750 0050 0 050 0 050 0 025 0 025 0750 0 750 0 050 0 050 EXTRUDED ANGLE ORIGINAL MATERIAL SUBSTITUTION STD 2024 SHEET SHAPE MATERIAL A B D Ra a b E S 81 2024 T3511 0 600 0 650 0 045 0 045 0 060 0 060 0 000 0 600 0 650 0 050 0 120 5 86 2024 T3511 0 750 0 562 0 051 0 051 0 060 0 060 0 000 0 750 0 562 0 063 0 160 5 95 2024 T3511 0 800 0 800 0 050 0 050 0 062 0 060 0 000 0 800
110. ly Available Used to aid in determining wing twist Protractor Head with Commercially Available To ensure wings are level Bubble Level Bolts Machined to Fabricate Locally Used to determine wing Specific Lengths twist 3 Model 172 Series Wing Twist Check Procedure A Mark Wing Station Reference Points Refer to Figure 601 1 Locate WS 39 00 Make a mark with a felt tip pen approximately 0 50 inch aft of lateral row of rivets in wing leading edge spar flange 2 Locate WS 100 50 Make a mark with a felt tip pen approximately 0 50 inch aft of lateral row of rivets in wing leading edge spar flange 3 Locate outboard WS 207 00 Make a mark with a felt tip pen approximately 0 50 inch aft of lateral row of rivets in wing leading edge spar flange B Measure Wing Twist at Each Wing Station Refer to Figure 601 NOTE While performing the following procedure stay as far away as possible from the canned 1 areas of the wing At WS 39 00 a bolt A to a dimension of 2 00 inches Grind bolt B to a dimension of 1 00 inch Place these bolts 29 50 inches from each other on the upper edge of the straightedge and secure using tape b Secure protractor to bottom of the straightedge c Hold straightedge parallel to wing station and place bolt A on mark d Set bubble in protractor to center and lock protractor to hold this reading At WS 100 50 Hold straightedge parallel to wing station Place bolt A on mar
111. mage A Minor skin dents and nicks are considered negligible and should be worked out by burnishing Repairable Damage A Skin damage exceeding that considered negligible damage be repaired by patching Typical repairs are illustrated in Chapter 51 Typical Skin Repair and Control Surface Repair A flight control surface which has been repaired or replaced must be balanced in accordance with the procedures outlined in Chapter 51 Flight Control Surface Balancing Replacement Damage A Assemblies that have been twisted or warped beyond usable limits and parts with extensive corrosion damage are considered replaceable Small parts which may be easily fabricated from materials available locally should be replaced 55 40 00 Page 801 Cessna Aircraft Company Jun 1 2005 WINDOWS CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT PAGE DATE 56 Title 56 List of Effective Pages 56 Record of Temporary Revisions 56 Table of Contents 56 00 00 Page 1 Jun 1 2005 56 10 00 Pages 801 804 Jun 1 2005 56 LIST OF EFFECTIVE PAGES Page 1 of 1 Cessna Aircraft Company Jun 1 2005 RECORD TEMPORARY REVISIONS Temporary Revision Number Page Number Issue Date By Date Removed CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL
112. n wing spars only in those locations where blind rivets were used during original manufacture i e fuel bay area of front spars on airplanes with integral fuel bays Existing tack rivet Existing rivet pattern Typical leading edge section pee Bonded Leading Edge Repair Figure 801 Sheet 1 57 26 00 Page 802 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL FLAP LEADING EDGE REPAIR 1 Flap Leading Edge Damage Classification A Damage to the wing flap can be divided into three major categories and is detailed in Wing Damage Classification 2 Flap Repairs A Repairs to the flap leading edge are illustrated in Figure 801 Repairs to the corrugated skin are illustrated in Chapter 51 Typical Skin Repairs Figure 802 57 27 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 57 27 00 802 Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL FLAPS AND AILERONS General A Each wing flap assembly is constructed of a spar ribs upper and lower skins and leading edge skin B Each aileron assembly is constructed of a single spar ribs upper and lower skin Balance weights are installed in the lower inboard leading edge and are retained with screws Flight control surfaces which have been repaired or replaced must be balanced in ac
113. nch in diameter depending on the length of the crack and the thickness of the material This is a temporary repair NOTE If temporary repairs are made operations should be kept to a minimum until replacement of window or windshield can be made 4 Surface Patch A following procedure should be used when preparing a surface patch 1 2 3 4 5 Trim away damaged area round all corners Cut a piece of plastic of sufficient size to cover the damaged area and extend 34 inch on each side of crack or hole Bevel edges as shown in Figure 801 NOTE If section to be repaired is curved shape surface patch to the same contour by heating it in an oil bath at a temperature of 248 F to 302 or it may be heated ona hotplate until soft Boiling water should not be used for heating Coat surfaces to be bonded evenly with plastic solvent adhesive acrylic chips dissolved in methylene chloride and place immediately over the hole Maintain a uniform pressure of 5 to 10 pounds per square inch on the surface patch for a minimum of 3 hours Allow surface to dry 24 to 36 hours before sanding or polishing is attempted 5 Insert Plug Patch A following procedure should be used when preparing a plug patch 1 Trim hole to a perfect circle or oval and bevel edges slightly 56 10 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B183 2
114. ness Aileron Damage Criteria Corrugated Skin Aileron A Negligible damage Any smooth dents in the aileron skin that are not more than 0 050 inch 1 27 mm below contour and circumscribable with not more than a 1 5 inch 38 1 mm diameter circle that have no evidence of skin tears cracks or skin penetrations and which do not include a corrugation constitute negligible damage Repairable damage Dents or dings deeper and or larger than specified must be repaired Skin tears cracks or penetrations must be repaired Dings that include corrugations are unlikely to be reworkable but may be repaired by replacing the damaged area Corrugated skin material is available from Cessna Special care must be taken to minimize added weight since the surface must be rebalanced after rework Damage Necessitating Replacement Of Parts Because of the balance requirements multiple areas of damage may require replacement of skins to allow balance limits to be attained Aileron Damage Criteria Model 206 Aileron A Negligible damage Any smooth dents in the aileron skin that are not more than 0 030 inch 0 76 mm below contour and circumscribable with not more than 1 5 inch 38 1 mm diameter circle that have no evidence of skin tears cracks or skin penetrations which are not stress wrinkles and do not interfere with internal structure constitute negligible damage Repairable damage Dents or dings deeper and or larger than specified must b
115. ng heat to the damaged area using a heat gun Replacement ribs may be installed using protruding head or dimpled flush rivets Damage Necessitating Replacement Of Parts Where extreme damage has occurred complete leading edge skin panels should be replaced Wing Strut Damage Criteria A Negligible damage Any smooth dents in the strut that are not more than 0 090 inch 2 03 mm below contour and circumscribable with not more than a 3 0 inch 76 2 mm diameter circle is negligible damage Minor scratches which do not involve removal or displacement of strut material is negligible damage Because of the critical nature of the strut any non cosmetic scratches must be reworked Repairable damage For grooves in the strut caused by fairings strut may be repaired if groove is less than 0 020 inch and is more than 0 75 inch from a rivet center For lower trailing edge strut damage typically caused by door hitting strut strut may be repaired if groove depth is less than 50 of original material thickness Damage Necessitating Replacement Of Parts For grooves in the strut caused by fairings strut must be replaced if groove exceeds 0 010 inch in depth and is less than 0 75 inch from a rivet center AND OR if groove exceeds 0 020 inch in depth and is more than 0 75 inch from a rivet center For lower trailing edge strut damage typically caused by door hitting strut strut must be replaced if groove is deeper than 50 of the original material thick
116. nside of the bay in the area of the leak while soap bubble solution is applied to the outside of the bay After the leak source has been found proceed as follows 1 Remove existing sealant in the area of the leak as described in Chapter 57 Wing Fuel Repairs 2 Clean the area and apply a fillet seal Press sealant into leaking area with a small paddle working out all air bubbles 3 If leakage occurs around a rivet or bolt restrike the rivet or loosen bolt retorque and reseal around nutplate 4 Apply fay surface door sealant to access doors fuel quantity transmitters etc if removed and install 5 Test fuel bay for leakage as outlined in Testing Integral Fuel Bay 6 Curing Time A B Class B 2 sealant has a maximum tack free time of 40 hours and a maximum cure time of 72 hours These values are based on a standard condition of 77 F 25 C and 50 percent relative humidity Class B 1 2 sealant has a maximum tack free time of 10 hours and a maximum cure time of 30 hours These values are based on a standard condition of 77 F 256 and 50 percent relative humidity The cure of sealants can be accelerated by an increase in temperature and or relative humidity Warm circulating air at a temperature not to exceed 140 606 may be used to ac celerate cure Heat lamps may be used if the surface temperature of the sealant does not exceed 140 606 At temperatures above 120 496 the relative humidity wil normally b
117. nt processes The dye applied to a clean metallic surface will enter small openings or cracks by capillary action After the dye has an opportunity to be absorbed by any surface discontinuities the excess dye is removed and a developer is applied to the surface The developer acts like a blotter to draw the dye from cracks or fissures back to the surface giving visible indication of any fault that is present on the surface The magnitude of the fault is indicated by the quantity of dye brought back to the surface by the developer 5 Corrosion Damage Limits A Following cleaning and inspection of the corroded area the actual extent of the damage may be evaluated using the following general guidelines and sound maintenance judgement 1 Determine the degree of corrosion damage light moderate or severe with a dial type depth gage if accessibility permits If the area is inaccessible clay impressions or any other means which will give accurate results should be used In the event the corrosion damage is severe or worse contact Cessna Propeller Aircraft Product Support P O Box 7706 Wichita KS 67277 USA for assistance Light Corrosion a Characterized by discoloration or pitting to a depth of approximately 0 001 inch maximum Moderate Corrosion a Appears similar to light corrosion except there may be blistering or some evidence of scaling or flaking Pitting depths may be as deep as 10 percent of the material thic
118. ntrol surface balance must be verified after repair or painting Proper balance of control surfaces is critical to prevent flutter during normal operating conditions 2 Tools and Equipment NAME NUMBER MANUFACTURER USE Control Surface 5180002 1 Cessna Aircraft Co Cessna Balance elevator and Balance Fixture Kit Part Distribution aileron 5800 E Pawnee P O Box 1521 Wichita KS 67218 Scale 0 10 Pounds Commercially Available Balance rudder in 0 01 Pound increments 3 Procedures for Balancing Control Surfaces A The flight control surface balancing fixture kit part number 5180002 1 is shown in Figure 1 1 2 3 Balance of control surfaces must accomplished a draft free room or area Place hinge bolts through control surface hinges and position on knife edge balancing mandrels refer to Figure 2 for positioning of balancing control surfaces Make sure all control surfaces are in their approved flight configuration painted if applicable trim tabs installed static wicks and all tips installed Place balancing mandrels on a table or other suitable flat surface Adjust trailing edge support to fit control surface being balanced while center of balancing beam is directly over hinge line Remove balancing beam and balance the beam itself by adding washers or nuts required at end opposite the trailing edge support When positioning balancing beam on control surface avoid rivets to provide a smooth surface for
119. o field repair typical repairs applicable to structural components information relative to material substitution and fastener installation and a description of procedures that must be performed with structural repair such as protective treatment of the repair and sealing C This manual will serve as a medium through which all single engine operators will be advised of actual repairs As service records indicate a requirement this manual will be revised to include additional specific repairs Airplane Identification A To identify structural differences to associated airplanes the specific airplane identity may appear in the figure and the text Items not identified for a specific airplane or group of airplanes are suitable for all airplanes Aerofiche microfiche A The Structural Repair Manual is prepared for Aerofiche presentation in addition to 8 11 inch loose leaf manual format To facilitate the use of the aerofiche a list of chapters with an aerofiche frame reference has bee tabulated and incorporated into the Introduction of the Structural Repair Manual Aerofiche is a microform reproduction of the contents of the 8 11 inch manual in a form convenient for service areas An aerofiche reader is required to view the 4 inch by 6 inch aerofiche card Each aerofiche card contains 12 horizontal rows of 24 images each An image displays information equal INTRODUCTION 1 Cessna Aircraft Company Jun 1 2005
120. o one mating part and install rivets or fasteners while sealant is still within its work life NOTE During sealing the supply of mixed sealant must be monitored to be certain it has not exceeded the normal work life To check use a small wooden paddle or tongue depressor to gather a small amount of sealant Touch this sealant to a piece of clean sheet metal If it adheres sealant can still be used if it doesn t adhere then the sealant has exceeded the allowable work life and must not be used 57 12 00 Page 801 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Apply a fillet seal to the repaired area on the inside of the bay Apply a fay surface seal to access doors fuel quantity transmitters etc if removed and install Allow sealant to cure refer to Curing Time for time requirements Clean stains on outer surface Test fuel bay for leaks as described in Testing Integral Fuel Bay aoa on wre rea 5 Sealing Fuel Leaks A First determine the source of the fuel leak Fuel can flow along a seam or structure of the wing for several inches making the leak source difficult to find A stained area is an indication of the leak source Fuel leaks can be found by testing the complete bay as described in Testing Integral Fuel Bay Another method of detecting the source of a fuel leak is to remove access doors and blow with an air nozzle from the i
121. on date LIST OF CHAPTERS Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LIST OF EFFECTIVE PAGES CHAPTER SECTION SUBJECT 51 Title 51 List of Effective Pages 51 Record of Temporary Revisions 51 Table of Contents 51 00 00 51 10 00 51 11 00 51 30 00 51 40 00 51 60 00 51 70 00 51 71 00 51 73 00 51 73 01 51 75 00 51 76 00 Cessna Aircraft Company PAGE Page 1 Pages 1 2 Pages 1 8 Pages 1 5 Pages 1 12 Pages 1 8 Page 801 Page 801 Pages 801 802 Page 801 Pages 801 808 Pages 801 803 DATE Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 Jun 1 2005 51 LIST OF EFFECTIVE PAGES Page 1 of 1 Jun 1 2005 RECORD TEMPORARY REVISIONS Temporary Revision Number Page Number Issue Date By Date Removed CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CONTENTS STANDARD PRACTICES AND STRUCTURES 51 00 00 Page 1 Geri amp ral EHE E nt ba cad cete ese a a HA hb ale MUERE EK EE IRA 51 00 00 Page 1 Descriptions sess tesseram Nube per bei het a Ne epis 51 00 00 Page 1 DAMAGE INVESTIGATION AND CLASSIFICATION 51
122. on thin gauge aluminum alloy and the work must be accomplished without distorting or damaging the material with hammer blows or riveting tools All airplane power riveting is accomplished by upsetting the rivets against a bucking bar instead of striking the shank with a hammer To prevent deforming the rivet head a rivet set must be selected to fit each type of rivet The depth of this set must not touch material being riveted Parts requiring heat treatment should be heat treated before riveting since heat 51 40 00 Page5 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL 151 4 File Center punch Drill 3 eu File a flat on the Center punch Drill through head using manufactured head the head drill that is one size smaller than the rivet shank Drill Drift punch Drift punch wv ave x is qua q Drill through dimple Remove weakend Punch out rivet countersunk or head with drift with punch countersunk head punch using drill that is one size smaller than the rivet shank Distance E should equal twice the rivet diameter Resultant crank E OW BEES Y 2 Incorrect too Correct edge A close to edge margin 558211024
123. or electrochemical action and converts it to a metallic compound such as an oxide hydroxide or sulfate All metals used in airplane construction are subject to corrosion If exposed attack may take place over an entire metal surface It may penetrate a surface at random forming deep pits or may follow grain boundaries Corrosion may be accentuated by stresses from external loads or from lack of homogeneity in the metallic structure or from improper heat treatment It is promoted by contact between dissimilar metals or with materials which absorb moisture such as wool rubber felt dirt and so forth NOTE Foradditional information on corrosion control for aircraft refer to the FAA Advisory Circular No 43 4 1 Referto Figure 1 for a simplified illustration of the conditions which must exist for electrochemical corrosion to occur There must be a metal that corrodes and acts as the anode b There must be a less corrodible metal that acts as the cathode c There must be a continuous liquid path between the two metals which acts as the electrolyte usually condensation and salt or other contamination d There must be a conductor to carry the flow of electrons from the cathode to the anode This conductor is usually in the form of a metal to metal contact rivets bolts welds etc 2 The elimination of any one of the four conditions described above will stop the corrosion reaction process as shown in Figure 1 8 One ofthe bes
124. orrugations Refer to Chapter 51 Typical Skin Repairs Figure 802 Access to the internal fin structure is best gained by removing skin attaching rivets on one side of the rear spar and ribs and springing back the skin Access to the stabilizer may be gained by removing skin attaching rivets on one side and springing back the skin If the damaged area would require a repair which could not be made between adjacent ribs or a repair would be located in an area with compound curves replacement of parts is recommended Replacement Damage A Hinge brackets and small ribs should be replaced rather than repaired In general where parts are available the easiest and most satisfactory repairs can be accomplished by replacing the damaged parts If the damaged area would require a repair which would not be made between adjacent ribs or the repair would be located in an area with compound curves complete skin panels must be replaced Ribs and spars may be repaired but replacement is generally preferable Where damage is extensive replacement of the entire assembly is recommended 55 30 00 Page 801 Cessna Aircraft Company Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL RUDDER Rudder A The rudder is constructed of a spar ribs and skin A torque tube incorporating a lower hinge bracket is attached to the lower leading edge A balance weight is bolted to the upper tip leading edge Negligible Da
125. pairs to the same area must not be made but a larger repair incorporating both repairs may be made Decisions regarding replacement of parts should be made based on the feasibility of repair verses complete replacement of the skin 14 Wing Flap Damage Criteria Model 206 Flap A Negligible damage Any smooth dents in the flap skin that are not more than 0 030 inch 0 76 mm below contour and circumscribable with not more than a 1 5 inch 88 1 mm diameter circle that have no evidence of skin tears cracks or skin penetrations and which do not include corrugations constitute negligible damage Repairable damage Dents or dings deeper and or larger than specified may be repaired Skin tears cracks or penetration must be repaired Dings that include understructure ribs must be repaired by reforming or removal and replacement of the rib Reevaluation of the skin after the repair of the understructure will determine if the skin damage is negligible repairable or replacement damage Damage Necessitating Replacement Of Parts Multiple repairs to the same area must not be made but a larger repair incorporating both repairs may be made Skins must be replaced if damage extends across more than one rib 57 10 00 4 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL WING FUEL BAY REPAIRS 1 Preparing Damaged Area In Wing Fuel Bay for Repair A Before performing any main
126. pe wrench 5 Apply a smooth even pull when applying torque pressure If chattering or a jerking motion occurs during final torque back off and re torque 6 When installing a castellated nut start alignment with cotter pin hole at minimum recommended torque plus friction drag torque and do not exceed maximum torque plus friction drag If hole and nut castellation do not align change washers or nut and try again Exceeding maximum recommended torque is not recommended unless specifically allowed or recommended for that particular installation 11 Rivets for Plastic or Composite Parts A Unlike rivets in metallic joints blind rivets are often the rivet of choice for riveting non metallic materials because they may be installed without the hammering necessary to install solid rivets If the tail end of the rivet is adjacent to the non metal side install a washer over the shank to prevent the hole filling action built into blind rivets from overloading the non metal hole The hole in the washer should match the specified installation hole for the fastener If the tail end of the rivet is installed through metal substructure the washer is not necessary 51 40 00 Page 10 Cessna Aircraft Company Jun 1 2005 55 AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL B153 S Loose fastener 5582T1025 Red Lining of Fasteners Figure 3 Sheet 1 51 40 00 Page 11 Cessna Aircraft Company Jun 1 2005 C
127. places In cases of dimpled assemblies the process of forming the metal around a hole to form a conical indentation to receive the tapered head of a flush rivet or a screw the rivet holes shall be drilled after the sheets are dimpled When possible the exposed end of each clipped plug shall be coated with epoxy primer Blind rivets shall not be used in fuel bay areas except in cases of absolute necessity and must be sealed If blind fasteners other than blind rivets are encountered it is recommended that replacements be made with identical fasteners For a list of approved solid shank and Hi Shear rivet substitutions refer to Tables 1 and 2 Rivet Diameters A Rivet diameters range from 3 32 inch to 3 8 inch Sizes of 1 8 inch 5 32 inch and 3 16 inch are most frequently used Since smaller diameter rivets lack proper structural qualities and larger diameter rivets dangerously reduce the splice or patch area extreme care should be exercised before substituting other than the specified sizes of rivet diameter Rivet Lengths A Proper length of rivets is an important part of a repair Should too long a rivet be used the formed head will be too large or the rivet may bend or be forced between the sheets being riveted Should too short a rivet be used the formed head will be too small or the riveted material will be damaged If proper length rivets are not available longer rivets may be cut off to equal the proper length not grip
128. r until desired finish is obtained 1 Repaint finished area with matching paint Refer to the applicable Maintenance Manual Chapter 51 73 00 Page 801 Jun 1 2005 15 approximately Smooth patch area with fine sandpaper Fa Fill back side with resin as necessary to obtain original thickness CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Clean damaged area thoroughly Clean and sand surrounding area First patch placed over entire damaged and Crack or damaged area cleaned area 161 2222 gt 2222 22 22 Wa 2222 22222 222222 1 22 NOTE Refer to repair of glass fiber parts 25 s S3 gt oog 09 50 gt Os lt Go Sa 5 pE E BG ogg oc BESO 802 1 2005 before attempting glass fiber repair 51 73 00 Typical Glass Fiber Panel Repair Figure 801 Sheet 1 Cessna Aircraft Company CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL REPAIR OF THERMO FORMED THERMO PLASTIC COMPONENTS 1 Thermo formed Thermo Plastic Repair A Repair of puncture or holes in thermo formed plastics can be made by trimming out the damaged area removing any paint in the area and installing an overlapping beveled or flush patch of identical material Doublers may be installed behind
129. re of the formed sheet metal or the reinforced formed sheet metal type Repair of Webs or Flanges A following procedures are for the repair of cracked bulkhead webs or flanges 1 Acceptable methods of repairing various types of cracks occurring in service are shown in Figures 801 and 802 2 Stop drill 30 0 128 inch minimum holes at extreme ends of cracks to prevent further cracking 3 Reinforcements should be added to carry stresses across damaged portion and stiffen the joints NOTE The condition causing such cracks to develop at a particular point may be stress concentration at that point in conjunction with repetition of stress such as produced by vibration of the structure The stress concentration may be due to defects such as nicks scratches tool marks and initial stresses or cracks from forming or heat treating operations An increase in sheet thickness alone is usually beneficial but does not necessarily remedy the condition leading to the cracking Patch type repairs are generally employed and are usually satisfactory in restoring the original material strength characteristics Repair of Channels A following procedures are for the repair of severely bent kinked or torn channels 1 practical severely bent kinked or torn portions of bulkheads should be removed and replacement sections installed and joined at the original splice joint 2 Ifthe procedure outlined in the preceding step is not
130. red Dings that include ribs must be repaired by reforming or removal and replacement of the rib Reevaluation of the skin after the repair of the understructure will determine if the skin damage is negligible repairable or requires replacement Damage Necessitating Replacement Of Parts Where extreme damage has occurred complete leading edge skin panels should be replaced Bonded Leading Edge Damage Criteria A Negligible damage Any smooth dents in the wing leading edge skin that are not more than 0 030 inch 0 76 mm below contour and circumscribable with not more than a 1 5 inch 38 mm diameter circle that has no evidence of skin tears cracks or skin penetrations which are not stress wrinkles and do not interfere with internal structure constitute negligible damage However because of the critical nature of the wing leading edge this cosmetic repair should be completed 57 10 00 Page 2 Cessna Aircraft Company Jun 1 2005 10 11 12 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Repairable damage Dents or dings deeper and or larger than specified above must be repaired Skin tears cracks or penetrations must be repaired Dings that include ribs must be repaired by reforming or removal and replacement of the rib Reevaluation of the skin after the repair of the understructure will determine if the skin damage is negligible repairable or requires replacement Bonded ribs may be removed by applyi
131. result of contact with a relatively smoothly contoured object NOTE dent like form of damage to skin be the result of the peening action of a smoothly contoured object contacting it If the inner surface of skin shows no contour change consider that such damage results in a local cross sectional area change 2 Crease A damaged area which is depressed or folded back upon itself in such a manner that its boundaries are sharp or well defined lines or ridges Consider it to be the equivalent of a crack 3 Abrasion An abrasion is a damaged area of any size which results in a cross sectional area change due to scuffing rubbing scraping or other surface erosion It is usually rough and irregular 4 Gouge A gouge is a damaged area of any size which results in a cross sectional area change It is usually caused by contact with a relatively sharp object which produces a continuous sharp or smooth channel like groove in the material 5 Nick A nick is a local gouge with sharp edges Consider a series of nicks in a line pattern to be the equivalent of a gouge 6 Scratch A scratch is a line of damage of any depth in the material and results in a cross sectional area change It is usually caused by contact with a very sharp object 7 Crack A crack is a partial fracture or complete break in the material with the most significant cross sectional area change In appearance it is usually an irregular line and is normally the r
132. rings are made from high strength steel that is shot peened on the lower surface to increase the fatigue life of the part The shot peened layer is between 0 010 and 0 020 inch thick If the protective layer of paint is chipped scratched or worn away the steel may corrode rust a Ifthe corrosion pit depth is greater than the thickness of the shot peen layer the gear spring fatigue life will be greatly reduced Operation from unimproved surfaces increases the likelihood of damage Corrosion removal and repair 1 If damage to the paint finish of the landing gear spring is found examine the damage area for signs of corrosion red rust WARNING High strength steel parts are very susceptible to hydrogen embrittlement Acidic solutions such as rust removers and paint strippers have been found to cause hydrogen embrittlement Hydrogen embrittlement is an undetectable time delayed process Since the process is time delayed failure may occur after the part is returned to service The only reliable way to prevent hydrogen embrittlement is not to use chemical rust removers or paint strippers on landing gear springs Carefully remove any rust by light sanding a The sanding should blend the damage into the surrounding area in an approximate 20 1 ratio EXAMPLE An 0 005 inch pit must be blended to a 0 10 inch radius or 0 20 inch diameter b Make sure the final sanding marks are along an inboard to outboard direction or along
133. rn apio Cele Ne e UD Sd 57 23 00 Page 801 AUXILIARY SPAR REPAIRS 22 ere iyu ege ee x rec edge bed elei 57 24 00 Page 801 General RE AR ARRA RAN A aca 57 24 00 Page 801 Auxiliary Spar 57 24 00 Page 801 Auxiliary Spar 57 24 00 Page 801 CONTENTS Page 1 of 2 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL LEADING EDGE REPAIRS Leading Edge Damage Classification Leading Edge Repairs Notes and Repair Limits BONDED LEADING EDGE REPAIR Bonded Leading Edge Damage Classification Bonded Leading Edge Repair FLAP LEADING EDGE REPAIR Flap Leading Edge Damage Flap FLAPS AND AILERONS Generali erem uk n Damage Criteria Flap and Aileron Repair WING LIFT Wing Strut Damage Classification Cessna Aircraft Company 57 25 00 Page 801 57 25 00 Page 801 57 25 00 Page 801 57 25 00 Page 801 57 26 00 Page 801 57 26 00 Page 801 57 26 00 Page 801 57
134. rospace Commercially available Commercially available Commercially available USE Repair composite structures Cover patches while curing Adhesive resin Adhesive resin Adhesive resin Cleaning solvent Abrading smoothing Cover patches when applying pressure clean an area at least 1 1 2 inches larger in diameter than the cut out section 2 Prepare necessary size and number of patches of glass fabric style No 181 WARNING Always follow manufacturer s mixing instructions carefully to ensure proper cure and prevent a spontaneous fire area 5 all the glass fabric patches by laying them on a polyethylene sheet and working the resin through the glass fabric with a small brush 20 Exterior Finish Cleaning Painting for painting procedures Cessna Aircraft Company 3 Mix sufficient amount of resin in accordance with manufacturers instructions 4 Ensure that hands are free from oil grease and dirt and apply an even coat of resin on sanded Place larger patch over cutout area working out all air bubbles and wrinkles If cutout is large enough to cause the patch to sag place a suitable support behind repair area Apply a second patch over the first patch working out all wrinkles and air bubbles After all patches have been applied brush the area with an even coat of resin and allow to cure Curing time is 24 hours at 774 0 Smooth patched area with 600 grit sandpape
135. s are available a Repair rivets have a shank that is 1 64 inch 0 016 inch larger diameter than a standard rivet but have the same size and shape heads b NAS1241 repair rivets replace MS20426 rivets if they have the same suffix c NAS1242 repair rivets replace MS20470 rivets if they have the same suffix 51 40 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL d NAS1738 NAS1939 and some NAS9301 through NAS9311 blind rivets also have oversize shanks Replacement shall not be made with rivets of lower strength material Hi Shear Rivets 1 When Hi Shear rivets are not available replacement of sizes 3 16 inch or greater rivets shall be made with bolts of equal or greater strength than the rivet being replaced and with self locking nuts of the same diameter It is permissible to replace Hi Shear rivets with Hi Lok bolts of the same material diameter and grip length Blind Rivets 1 Blind rivets have higher deflection rates in shear than standard solid rivets are more susceptible to fatigue failure and are not as strong as solid rivets in thin sheets For this reason it is not advisable to replace any considerable number of solid rivets in a given joint by blind rivets because this may result in overstressing the remaining solid rivets The hollow blind rivet shall not be used The blind rivet shall be of the same or greater strength than the rivet it re
136. severe the corrosion Relative activity in descending order is as follows a Magnesium and its alloys b Aluminum alloys 1100 3003 5052 6061 220 355 356 cadmium and zinc c Aluminum alloys 2014 2017 2024 and 7075 d Iron lead and their alloys except stainless steel e Stainless steels titanium chromium nickel copper and their alloys Graphite including dry film lubricants containing graphite Intergranular Corrosion 1 Selective attack along the grain boundaries in metal alloys is referred to as intergranular corrosion It results from lack of uniformity in the alloy structure It is particularly characteristic of precipitation hardened alloys of aluminum and some stainless steels Aluminum extrusions and forgings in general may contain nonuniform areas which in turn may result in galvanic attack along the grain boundaries When attack is well advanced the metal may blister or delaminate which is referred to as exfoliation Stress Corrosion 1 This results from the combined effect of static tensile stresses applied to a surface over a period of time In general cracking susceptibility increases with stress particularly at stresses approaching the yield point and with increasing temperature exposure time and concentration of corrosive ingredients in the surrounding environment Examples of parts which are susceptible to stress corrosion cracking are aluminum alloy bell cranks landing gear
137. shock struts with pipe thread type grease fittings clevis points and shrink fits Corrosion Fatigue 1 This is a type of stress corrosion resulting from the cyclic stresses on a metal in corrosive surroundings Corrosion may start at the bottom of a shallow pit in the stressed area Once attack begins the continuous flexing prevents repair of protective surface coating or oxide films and additional corrosion takes place in the area of stress 3 Typical Corrosion Areas A This section lists typical areas of the airplane which are susceptible to corrosion These areas should be carefully inspected at periodic intervals to detect corrosion as early as possible 1 Engine Exhaust Trail Areas a Gaps seams and fairings on the lower fuselage aft of the engine exhaust pipe s are typical areas where deposits may be trapped and not reached by normal cleaning methods b Around rivet heads skin laps and inspection covers on the airplane lower fuselage aft of the engine exhaust pipe s should be carefully cleaned and inspected Battery Box and Battery Vent Opening a The battery battery cover battery box and adjacent areas especially areas below the battery box where battery electrolyte may have seeped are particularly subject to corrosive action If spilled battery electrolyte is neutralized and cleaned up at the same time of spillage corrosion can be held to a minimum by using a baking soda solution to neutralize the le
138. standard bend radius for 2024 T3 or 2024 T4 use 2024 0 and then heat treat to 2024 T42 after forming In the event that the original temper was T3 it may be necessary to increase the material thickness sufficiently to provide strength equivalent to that of the original part It is often practical to cut repair pieces from service parts listed in the parts catalog Steel sheet material for reinforcement is 4130 steel heat treated to a minimum of 90 000 pounds per square inch The firewall is annealed stainless steel sheet Extrusions and Formed Sections A Refer to Figure 1 This section provides information on extrusions and formed sections It also provides details of equivalent built up sections for extrusions Alternative materials are provided for equivalent sections and formed sections Use of equivalent built up sections for extrusions are to be utilized only when the proper extrusions are not available They are intended to be cold formed from raw stock in sheet forms that have already been heat treated to the required condition But when workability is required the parts may be formed from 2024 0 aluminum and then heat treated to the T42 condition before installation When forming the section care must be taken to ensure that the bend radii and the cross section areas are not reduced below the minimum shown in the diagrams In some cases equivalent sections are not given because it is impractical to build them from sheet stock Il
139. t ways to eliminate one of the four described conditions is to apply an organic film such as paint grease plastic etc to the surface of the metal affected This will prevent the electrolyte from connecting the cathode to the anode and since current cannot flow it prevents corrosive reaction 4 At normal atmospheric temperatures metals do not corrode appreciably without moisture but the moisture in the air is usually enough to start corrosive action 5 When components and systems constructed of many different types of metals must perform under various climatic conditions corrosion becomes a complex problem The presence of salts on metal surfaces from sea coast operation greatly increases the electrical conductivity of any moisture present and accelerates corrosion 6 Other environmental conditions which contribute to corrosion are a Moisture collecting on dirt particles b Moisture collecting in crevices between lap joints around rivets bolt and screws 2 Types of Corrosion A Direct Surface Attack 1 The most common type of general surface corrosion results from direct reaction of a metal surface with oxygen in the atmosphere Unless properly protected steel will rust and aluminum and magnesium will form oxides The attack may be accelerated by salt spray or salt bearing air by industrial gasses or by engine exhaust gasses Pitting 1 While pitting can occur in any metal it is particularly characteristic o
140. tenance in fuel bay area it will be necessary to defuel and purge the fuel bay To defuel and purge the fuel bay proceed as follows WARNING During all fuel system servicing procedures fire fighting equipment must be available WARNING Always ground airplane prior to performing any maintenance of the fuel system WARNING Avoid drainage from residual fuel held in disconnected fuel lines this accumulation constitutes a fire hazard WARNING Use NS 40 RAS 4 Snap On Tools Corp Kenosha Wisconsin MIL T 83483 thread compound anti seize graphite petrolatum or engine oil as a thread lubricant or to seal leaking connections Apply sparingly to all but first two threads of male fittings being careful not to allow entry of compound into fuel system NOTE Covers or caps should be installed on lines and fittings to prevent entry of foreign material 10 and to prevent damage to threads Ground airplane to a suitable ground stake Ensure airplane battery switch is in OFF position Turn fuel selector valves to OFF position Remove fuel filler cap on bay that is to be defueled insert defueling nozzle Remove as much fuel as possible through filler opening Remove drain valves from bottom side of fuel bay and drain remaining fuel into a clean open container Use defueling nozzle to remove fuel from container If necessary repeat procedures
141. the long dimension of the spring After the sanding is complete measure the depth of the damage removal a Make sure the depth of the damage is not more than 0 010 to 0 012 inch deep and has not penetrated the shot peen layer If the shot peened layer has been penetrated the gear spring must be removed and sent to an approved facility to be re shotpeened a The shotpeen specification is to be Almen intensity of 0 012 to 0 016 using 330 steel shot After the spring is installed refinish any damaged or removed finish paint NOTE Additional information regarding corrosion control can be found in AC 43 4 Chapter 6 or AC43 13 1B Chapter 6 51 11 00 Page7 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Axle bolt hole corrosion 1 Operation of an airplane on skis increases the loads on the lower part of the gear spring because of the unsymmetrical and twisting loads a The increased loads have produced spring fractures that originate from pits in the axle attach holes 1 Catastrophic failures have occurred from fatigue cracks as small as 0 003 to 0 010 inch long that originated at pits b Although operation on skis causes more loads the criteria applies to all airplanes There is no acceptable damage depth for pits that develop in the axle bolt holes If pits or corrosion is found it must be removed by reaming subject to the following limitations a
142. the patch where additional strength is desired MPK or any commercially available solvent that will soften and dissolve the plastic may be used as the bonding agent Dissolving some of the plastic shavings in the solvent will furnish additional working time Moderate pressure is recommended for best results Curing time will vary with the agent used but repairs should not be strained until fully cured Cracks can be repaired by saturating the crack itself with the solvent then filling with an epoxy filler or a paste made of the plastic shavings and the solvent Again the crack may be reinforced with a doubler on the back side for additional strength After the repair has been made the area may be sanded smooth and painted Parts that are extensively damaged should be replaced instead of repaired 2 Temporary Repairs A Crack Repair 1 Itis permissible to stop drill crack s that originate at the edge of a fairing if the crack is less than 2 inches 50 mm in length a Stop drill the crack with a Number 30 0 128 inch diameter drill bit b crack be stop drilled only once NOTE Acrackthat passes through a fastener hole and does not extend to the edge of the part may be stop drilled at both ends of the crack c fairing that has a crack that progresses past a stop drilled hole must be repaired or replaced d A fairing that has any of the following conditions must have a repair made as soon as practical 1 Acrack t
143. tip horizontal stabilizer tip or vertical fin tip damage If the damage is due to an aft load the rear spars should be checked for indications of compression damage for the full length including the fuselage components 3 Damage Classification A Damage to the airplane can be divided into three major categories negligible damage repairable damage and major replacement damage These categories are intended to provide the mechanic with some general guidelines to use in determining the extent and criticalness of any damage Obviously there will be some overlapping between categories and common sense should be used in determining the final action to be taken with regard to any damage 1 For damage criteria of specific structure wings fuselage and so forth refer to applicable chapters within this repair manual 4 Refinishing Damaged Areas Following Repairs A Areas of structure which are damaged and then repaired in the field must be refinished to restore the original paint and corrosion protectant properties to factory standards Refer to applicable airplane Maintenance Manual Chapter 20 Exterior Finish Cleaning Painting for refinishing procedures and required materials 51 10 00 Page 2 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL CORROSION AND CORROSION CONTROL GENERAL 1 General A Corrosion is a natural phenomenon which destroys metal by chemical
144. ts or bends in a stringer may be repaired by reforming or by replacing a section of the stringer Since aluminum work hardens it is much more likely to crack when reformed and should be carefully inspected for such cracks after rework Removal and replacement of damaged stringers is preferred to reformation Damage Necessitating Replacement Of Parts If a stringer is so badly damaged that more than one section must be spliced replacement is recommended Wing Auxiliary Spar Damage Criteria A B Negligible damage Minor scratches or abrasions are the only form of damage considered negligible to wing auxiliary spars Repairable damage Dents or bends in an auxiliary spar may be repaired by reforming or by replacing a section of the auxiliary spar Since aluminum work hardens it is much more likely to crack when reformed and should be carefully inspected for such cracks after rework Removal and replacement of a damaged section to the auxiliary spar is preferred to reformation Damage necessitating Replacement Of Parts If damage to an auxiliary spar would require a repair which could not be made between adjacent ribs the auxiliary spar must be replaced Wing Rib Damage Criteria A Negligible damage None other than minor scratches or abrasions 57 10 00 Page 1 Cessna Aircraft Company Jun 1 2005 CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL Repairable damage Dents or bends in a rib may be repaired b
145. view AA2682T1018 Rivet Removal and Rivet Edge Distance Figure 1 Sheet 1 51 40 00 Pages Cessna Aircraft Company Jun 1 2005 D CESSNA AIRCRAFT COMPANY SINGLE ENGINE STRUCTURAL REPAIR MANUAL treating process after rivet installation causes warping Assemblies that require heat treatment in a salt bath must be treated prior to assembly as the salt cannot be entirely washed out of the joints The use of hollow rivets in joining highly stressed parts is not permitted To determine if blind rivets may be substituted refer to Tables 1 and 2 Selection of the proper rivet and the proper number of rivets is very important Rivets must be of the proper length for the total thickness of the parts being riveted Ordinarily from 1 1 2 to 2 times the diameter of the rivet is the correct amount for the rivet shank to protrude through the material to form the head For heavy material such as plates or fittings from 2 to 2 1 2 times the rivet diameter may be used The rivet should not be excessively loose in the hole as this condition will cause the rivet to bend over while being driven and the shank will not be sufficiently expanded to completely fill the hole A drill from 0 002 inch to 0 004 inch larger than the rivet shank should be used for sheet and plate riveting Parts should be held firmly together by clamps screws or bolts while they are being drilled or riveted The bucking bar is to be held against the end of the rivet shank
146. wing 1 At WS 39 00 a bolt A to a dimension of 2 00 inches Grind bolt B to a dimension of 1 00 inch Place these bolts 29 50 inches from each other on the upper edge of the straightedge and secure using tape b Secure protractor to bottom of the straightedge c X Hold straightedge parallel to wing station and place bolt A on mark d Set bubble in protractor to center and lock protractor to hold this reading 2 At WS 100 50 Hold straightedge parallel to wing station Place bolt A on mark set protractor head against lower edge of straightedge and verify bubble in protractor head indicates level 3 At WS 207 00 a Remove bolt A from straightedge Grind another bolt A to a dimension of 0 45 inch Place this bolt 24 00 inches from bolt B and secure to straightedge b Hold straightedge parallel to wing station and place bolt A on mark c Check to assure that protractor bubble is still centered If proper twist is present the protractor readings will be the same parallel Oo m NOTE Forward or aft bolt may be lowered from wing 0 10 inch maximum to attain level indication 5 Model 206 T206 Series Wing Twist Check Procedure A Mark wing Station Reference Points Refer to Figure 603 1 Locate WS 39 00 Make a mark with a felt tip pen approximately 0 50 aft of lateral row of rivets in wing leading edge spar flange 2 Locate WS 100 00 Make a mark with a felt t
147. wing conditions shall have a repair made as soon as practical a Acrack that is longer than 2 inches b Acrack that does not originate from the trailing edge or a trailing edge rivet c Cracks in more than six trailing edge rivet locations per skin 6 Affected control surfaces with corrugated skins and having a stop drilled crack that does not extend past the stop drilled hole may remain in service without additional repair 7 Refer to Figure 802 as applicable for repair information B Corrugated Flap Skin Repair 1 It is permissible to stop drill crack s that originate at the trailing edge of the control surface provided the crack s is are not more than 2 inches in length 2 Stop dill crack s using a Number 30 0 128 inch diameter drill 3 crack may only be stop dilled once NOTE A crack that passes through a trailing edge rivet and does not extend to the trailing edge of the skin may be stop drilled at both ends of the crack 4 Any control surface that has a crack that progresses past a stop drilled hole shall be repaired or replaced 5 A control surface that has any of the following conditions shall have a repair made as soon as practical a Acrack that is longer than 2 inches 6 Acrack that does not originate from the trailing edge or a trailing edge rivet c Cracks in more than six trailing edge rivet locations per skin 6 Affected control surfaces with corrugated skins and having a stop drilled cra
148. xtent and criticalness of any damage Negligible Damage A Any smooth dents in the fuselage skin that are free from cracks abrasions and sharp corners and which are not stress wrinkles and do not interfere with any internal structure or mechanism may be considered as negligible damage In areas of low stress intensity cracks deep scratches or deep sharp dents which after trimming or stop drilling can be enclosed by a two inch circle can be considered negligible if the damaged area is at least one diameter of the enclosing circle away from all existing rivet lines and material edges Stop drilling is considered a temporary repair and a permanent repair must be made as soon as practical Mild corrosion appearing upon clad aluminum surfaces does not necessarily indicate incipient failure of the base metal However corrosion of all types must be carefully considered and approved remedial action taken Small cans appear in the skin structure of all metal airplanes and should not necessarily be a cause for concern However It is strongly recommended that wrinkles which appear to have originated from other sources or which do not follow the general appearance of the remainder of the skin panels be thoroughly investigated Except in the landing gear bulkhead areas wrinkles occurring over stringers which disappear when the rivet pattern is removed may be considered negligible However the stringer rivet holes may not align perfectly with
149. y reforming or by replacing a section of the rib Since aluminum work hardens it is much more likely to crack when reformed and should be carefully inspected for such cracks after rework Removal and replacement of a damaged section to the rib is preferred to reformation Damage Necessitating Replacement Of Parts Leading and trailing edge ribs that are extensively damaged can be replaced However due to the necessity of unfastening an excessive amount of skin in order to replace the rib they should be repaired if practical Center ribs between the front and rear spar should always be repaired if practical Wing Spar Damage Criteria A Negligible damage Due to the stress which wing spars encounter very little damage can be considered negligible All cracks stress wrinkles deep scratches and sharp dents must be repaired Smooth dents light scratches and abrasions may be considered negligible Repairable damage While it is possible to repair the spar channel by reforming a section of the spar replacement is preferred A service kit SK172 68 is available for replacement of the inboard end of the rear spar for damage that typically occurs with impact on the outboard leading edge Damage Necessitating Replacement of Parts Damage so extensive that repair is not practical requires replacement of complete wing spar Wing Fuel Bay Spars Rib Damage Criteria A Negligible damage Any smooth dents in the wing fuel spar and ribs that h

1996...The Cessna Single Engine Structural Repair Manual

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