# AC 43.13-1B: Acceptable Methods, Techniques, and Practices—Aircraft Inspection and Repair

- Type: Advisory circular (guidance, not mandatory unless a rule or AD requires it)
- Status: Historical
- Version in force: [AC 43.13-1B CHG 1 Ed Update](https://openairdocs.org/ac/43.13-1), approved September 27, 2001. This is a historical version
- Approved: September 8, 1998
- Last reviewed: June 11, 2024
- Family: AC 43.13-1
- Cancels: [AC 43.13-1A CHG 2](https://openairdocs.org/ac/43.13-1A-CHG-2)
- Cancelled by: [AC 43.13-1B CHG 1](https://openairdocs.org/ac/43.13-1B-CHG-1)
- 14 CFR parts: 43, 65, 145, 21, 105, 23, 91, 121, 25
- 14 CFR sections: [43.13](https://www.ecfr.gov/current/title-14/section-43.13), [43.16](https://www.ecfr.gov/current/title-14/section-43.16), [43.9](https://www.ecfr.gov/current/title-14/section-43.9), [105.43](https://www.ecfr.gov/current/title-14/section-105.43), [23.853](https://www.ecfr.gov/current/title-14/section-23.853), [91.411](https://www.ecfr.gov/current/title-14/section-91.411), [91.413](https://www.ecfr.gov/current/title-14/section-91.413), [25.869](https://www.ecfr.gov/current/title-14/section-25.869), [91.171](https://www.ecfr.gov/current/title-14/section-91.171)
- Pages: 696
- PDF: https://openairdocs.org/files/sha256/d0/82/d082f2bd98eba9877cc6ac089a1a2ac74618c8575b0e824cc9d0a910cdb0216d
- FAA DRS: https://drs.faa.gov/browse/excelExternalWindow/CCFE0D58D41D2C8086256A55006C6B67.0001
- DRS record last changed: November 21, 2025
- Page: https://openairdocs.org/ac/43.13-1B

## Versions

- [AC 43.13-1B CHG 1 Ed Update](https://openairdocs.org/ac/43.13-1) (Current), approved Sep 27, 2001, 756 pages. Editorial Update to AC 43.13-1B CHG 1.
- [AC 43.13-1B CHG 1](https://openairdocs.org/ac/43.13-1B-CHG-1) (Historical), approved Sep 9, 1998, 646 pages
- [AC 43.13-1B](https://openairdocs.org/ac/43.13-1B) (Historical), approved Sep 8, 1998, 696 pages
- [AC 43.13-1A CHG 2](https://openairdocs.org/ac/43.13-1A-CHG-2) (Historical), approved Dec 22, 1976, 358 pages
- [AC 43.13-1A CHG 1](https://openairdocs.org/ac/43.13-1A-CHG-1) (Historical), approved May 12, 1975, 356 pages
- [AC 43.13-1A](https://openairdocs.org/ac/43.13-1A) (Historical), approved Jan 1, 1972, 356 pages
- [AC 43.13-1](https://openairdocs.org/ac/43.13-1?v=E562265B3D7C20FF86257983004DB8A8) (Historical), approved Jan 1, 1965, 346 pages. AC 43.13-1 with all changes incorporated.

## ADs that cite it (95, newest 50)

- [AD 2024-19-08](https://openairdocs.org/ad/2024-19-08): Nose/Tail Landing Gear · Diamond Aircraft Industries DA 40 and DA 40F. Effective Nov 29, 2024
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- [AD 2020-16-15](https://openairdocs.org/ad/2020-16-15): Fuselage · Viking Air DHC-2 Mk.I and DHC-2 Mk.III with STC SA00299NY or SA945NE. Effective Sep 21, 2020
- [AD 2015-08-04](https://openairdocs.org/ad/2015-08-04): Wings · FS 2000 L-14, AE-1, HE-1 and 77 more. Effective Jun 3, 2015
- [AD 2015-02-15](https://openairdocs.org/ad/2015-02-15): Elevator skins · Quest Aircraft Design Kodiak 100. Effective Feb 19, 2015
- [AD 99-26-19 R1](https://openairdocs.org/ad/99-26-19-R1): Wings · Piper J-2. Effective Jan 21, 2014
- [AD 99-01-05 R1](https://openairdocs.org/ad/99-01-05-R1) (historical): Wings · FS 2000 L-14, AE-1, HE-1 and 73 more. Effective Jan 14, 2014
- [AD 2013-10-04](https://openairdocs.org/ad/2013-10-04): Engine Exhaust · Piper PA-31, PA-31-325 and PA-31-350. Effective Jul 17, 2013
- [AD 2012-17-06](https://openairdocs.org/ad/2012-17-06): Flight Controls · Piper PA-24, PA-24-250 and PA-24-260. Effective Oct 22, 2012
- [AD 2012-04-10](https://openairdocs.org/ad/2012-04-10): Wings · Rogers, Burl A. 15AC and S15AC. Effective Apr 17, 2012
- [AD 2011-15-11](https://openairdocs.org/ad/2011-15-11): Wings · Nineteen and a half(19.5 U.S.) gallons wing tip fuel tank (Flint Aero STC SA5090NM). Effective Sep 6, 2011
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- [AD 2010-21-18](https://openairdocs.org/ad/2010-21-18): Wings · Wing tip extension fuel tanks (Aviation Enterprises STC SA02055AT or 3 others). Effective Nov 3, 2010
- [AD 2009-26-11](https://openairdocs.org/ad/2009-26-11): Spar caps · Thrush 600 S-2D, S-2R, S2R-G1 and 13 more. Effective Feb 24, 2010
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- [AD 2008-22-21](https://openairdocs.org/ad/2008-22-21) (historical): Main tubular spar · Allied Ag Cat Productions G-164, G-164A, G-164B and 7 more. Effective Dec 19, 2008
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- [AD 2001-23-03](https://openairdocs.org/ad/2001-23-03): Forward Door Post · Cessna 172N, 172P, 172RG and 5 more. Effective Dec 27, 2001
- [AD 2000-25-02 R1](https://openairdocs.org/ad/2000-25-02-R1): Wood Spars · American Champion 11AC, 11BC, 11CC and 27 more. Effective Jul 13, 2001
- [AD 2000-25-02](https://openairdocs.org/ad/2000-25-02) (historical): Wood Wing Spars · American Champion 11AC, 11BC, 11CC and 15 more. Effective Jan 19, 2001
- [AD 2000-07-03](https://openairdocs.org/ad/2000-07-03): Fuel Line Assembly · Robinson Helicopter R44. Effective May 11, 2000
- [AD 99-10-07](https://openairdocs.org/ad/99-10-07): Elevator Control Cable · Textron 65-90, 65-A90, 65-A90-1 and 9 more. Effective Jun 8, 1999
- [AD 98-19-14](https://openairdocs.org/ad/98-19-14): Airbrake Control System · Centrair 101, 101A, 101AP and 1 more. Effective Nov 9, 1998
- [AD 98-09-10](https://openairdocs.org/ad/98-09-10): Rudder Control Cables · Extra Flugzeugproduktions- und Vertriebs- EA-300 and EA-300S. Effective Jun 7, 1998
- [AD 96-16-08](https://openairdocs.org/ad/96-16-08): Main Deck Cargo Door · Boeing 727-100 Series and 727-200 Series with STC SA1797SO. Effective Aug 27, 1996
- [AD 96-02-04](https://openairdocs.org/ad/96-02-04): 80 Octane Avgas · Wytwornia Sprzetu Komunikacyjnego 6A4-150-B3 and 6A4-165-B3. Effective Feb 13, 1996
- [AD 94-14-10](https://openairdocs.org/ad/94-14-10): Aviation Gasoline · Wright Aeronautical Division, Curtiss-Wright Cyclone R-1820F-2A, Cyclone R-1820F-3A, Cyclone R-1820F-56 and 6 more. Effective Feb 13, 1996
- [AD 95-12-01](https://openairdocs.org/ad/95-12-01): Forward Wing Spars · Lavia Argentina PA-25, PA-25-235 and PA-25-260. Effective Jul 7, 1995
- [AD 92-15-13](https://openairdocs.org/ad/92-15-13): Wing Structure · Textron 99, 99A, 99A (FACH) and 4 more. Effective Aug 24, 1992
- [AD 90-04-06 R1](https://openairdocs.org/ad/90-04-06-R1): Propeller Governor External Oil Line · Lycoming Engines AEIO-320-D1B, AEIO-320-D2B, AEIO-320-E1A and 427 more. Effective May 28, 1991
- [AD 90-15-07](https://openairdocs.org/ad/90-15-07): Air Conditioning System · Short Brothers SD3-60 with STC SA3749SW. Effective Jul 31, 1990
- [AD 89-18-06](https://openairdocs.org/ad/89-18-06): Front Folding Seats · American Champion 7ECA, 7GCAA, 7GCBC and 3 more. Effective Sep 21, 1989
- [AD 78-26-10 R1](https://openairdocs.org/ad/78-26-10-R1): Cockpit Heater System · Great Lakes 2T-1A-1 and 2T-1A-2. Effective Dec 17, 1986
- [AD 86-21-08](https://openairdocs.org/ad/86-21-08): Battery Tray · SOCATA TB 10 and TB 20. Effective Oct 24, 1986
- [AD 85-04-01 R1](https://openairdocs.org/ad/85-04-01-R1): Hydraulic And Oxygen Line Inspection · M7 Aerospace SA226-AT, SA226-T, SA226-TC and 1 more. Effective Oct 28, 1985
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- [AD 78-23-14 R1](https://openairdocs.org/ad/78-23-14-R1): Tow Release System · BLANIK Blanik L-13 and Blanik L-13 AC. Effective Aug 17, 1984
- [AD 84-15-01](https://openairdocs.org/ad/84-15-01): Placard -Tethered Flight · Avian Balloon Falcon II, Skyhawk and Sparrow. Effective Jul 31, 1984
- [AD 84-13-01](https://openairdocs.org/ad/84-13-01): Generator Control J-Box · M7 Aerospace SA226-TC, SA227-AC (C-26A), SA227-AT and 1 more. Effective Jul 5, 1984
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- [AD 80-19-12 R1](https://openairdocs.org/ad/80-19-12-R1): Engine Mount Structure · Textron 76. Effective Jun 14, 1982
- [AD 81-23-01 R1](https://openairdocs.org/ad/81-23-01-R1): Beech Kit or Supplemental Type Certificate · Textron 100, 65, 65-88 and 20 more. Effective Jan 4, 1982

All 95: https://openairdocs.org/ac/43.13-1/ads.md

## Text

Each page's text from the PDF; open the PDF at a page with #page=N. Some pages are OCR from the scanned FAA advisory circular. OCR can misread letters and numbers. Those pages are marked (OCR).

### Page 1

**ADVISORY**

**CIRCULAR**

AC 43.13-1B

**ACCEPTABLE METHODS,**

TECHNIQUES, AND PRACTICES  AIRCRAFT

**INSPECTION AND REPAIR**

September 8, 1998

**DEPARTMENT OF TRANSPORTATION**

FEDERAL AVIATION ADMINISTRATION Flight Standard Service Regulatory Support Division

For sale by the Superintendent of Documents; U.S. Government Printing Office; Washington, DC 20402

### Page 2

Advisory Circular

**TITLE 14 OF THE CODE OF FEDERAL REGULATIONS (14 CFR) GUIDANCE MATERIAL**

### Page 3

Contents Paragraph Page

Page 3 Subject: ACCEPTABLE METHODS, TECHNIQUES, AND PRACTICESAIRCRAFT

INSPECTION AND REPAIR Date: 9/8/98 Initiated by: AFS-610 AC No: 43.13-1B Change:

1. PURPOSE. This advisory circular (AC) contains methods, techniques, and practices acceptable to the Administrator for the inspection and repair of nonpressurized areas of civil aircraft, only when there are no manufacturer repair or maintenance instructions. This data generally pertains to minor repairs. The repairs identified in this AC may also be used as a basis for FAA approval for major repairs. The repair data may also be used as approved data, and the AC chapter, page, and paragraph listed in block 8 of FAA form 337 when:

a. the user has determined that it is appropriate to the product being repaired;

b. it is directly applicable to the repair being made; and

c. it is not contrary to manufacturer’s data.

2. CANCELLATION. The AC 43.13-1A dated 1988 is canceled.

3. REFERENCE: Title 14 of the Code of Federal Regulations part 43, section 43.13(a) states that each person performing maintenance, alteration, or preventive maintenance on an aircraft, engine, propeller, or appliance shall use the methods, techniques, and practices prescribed in the current manufacturer’s maintenance manual or Instructions for Continued Airworthiness prepared by its manufacturer, or other methods, techniques, or practices acceptable to the Administrator, except as noted in section 43.16. FAA inspectors are prepared to answer questions that may arise in this regard. Persons engaged in the inspection and repair of civil aircraft should be familiar with 14 CFR part 43, Maintenance, Preventive Maintenance, Rebuilding, and Alteration, and part 65, Subparts A, D, and E of Certification:

Airmen Other Than Flight Crewmembers, and the applicable airworthiness requirements under which the aircraft was type certificated.

4. ACKNOWLEDGMENTS. The FAA would like to thank the following persons and organization for their assistance in producing AC 43.13-1B: Richard Finch, Richard Fischer, Michael Grimes, Ray Stits, William A.

Watkins, and the SAE, Aerospace Electronics and Electrical Systems Division. Acknowledgment is also extended to all in the aviation community who commented on the document.

5. COMMENTS INVITED. Comments regarding this AC should be directed to DOT/FAA;

ATTN: Airworthiness Programs Branch, AFS-610; PO Box 25082; Oklahoma City, OK 73125

Acting Deputy Director, Flight Standards Service

### Page 4

Contents Paragraph Page

Page 4 CHAPTER 1. WOOD STRUCTURE SECTION 1. MATERIALS AND PRACTICES

1-1.......................................................................................................................... GENERAL 1- 1

1-2................................................................................................................................ ......... WOODS 1- 1 FIGURE 1-1.RELATIVE SHRINKAGE OF WOOD MEMBER DUE TO DRYING 1- 1 TABLE 1-1. ......SELECTION AND PROPERTIES OF AIRCRAFT WOOD 1- 2 FIGURE 1-2. .................................................TAPERING OF FACEPLATE 1- 3

1-3........................................................................................MODIFIED WOOD PRODUCTS 1- 3

1-4.......................................................................................................................ADHESIVES 1- 3

1-5............................................................................................... BONDING PRECAUTIONS 1- 5

1-6............................................... PREPARATION OF WOOD SURFACES FOR BONDING 1- 5

1-7..............................................................................................APPLYING THE ADHESIVE 1- 6

1-8........................................................................................ASSEMBLY TIME IN BONDING 1- 6

1-9.............................................................................................. BONDING TEMPERATURE 1- 6

1-10..................................................................................................CLAMPING PRESSURE 1- 7

1-11............................................................................METHOD OF APPLYING PRESSURE 1- 7

1-12.1-17. [RESERVED.] … 1-8

**SECTION 2. HEALTH AND SAFETY**

1-18........................................................................................................................GENERAL 1- 9

1-19........................... SANDING IN AREAS OF EXISTING BOND JOINTS AND FINISHES 1- 9

1-20.................................................................HANDLING OF ADHESIVES AND FINISHES 1- 9

1-21.1-26. [RESERVED.] … 1-9

**SECTION 3. INSPECTION**

1-27........................................................................................................................GENERAL 1- 11

1-28............................................................... TYPES OF DETERIORATION AND DAMAGE 1- 11

1-29................................................................................................ INSPECTION METHODS 1- 11 FIGURE 1-3. ........ LIKELY AREAS TO INCUR STRUCTURAL DAMAGE 1- 13

1-30.1-35. [RESERVED.] … 1-14

**SECTION 4. REPAIRS**

1-36........................................................................................................................GENERAL 1- 15

1-37.....................................REPLACEMENT OF DRAIN HOLES AND SKIN STIFFENERS 1- 15

### Page 5

Contents Paragraph Page

1-38.........................................................CONTROL SURFACE FLUTTER PRECAUTIONS 1- 15

1-39...............................................................................................................SCARF JOINTS 1- 15 FIGURE 1-4.CONSIDERATION OF GR AIN DIRECTION WHEN MAKING SCARF JOINTS 1-15

1-40..................................................................................................... SPLICING OF SPARS 1- 15 FIGURE 1-5.METHOD OF SPLICING SOLID OR LAMINATED RECTANGULAR SPARS 1-17 FIGURE 1-6. .......................... METHOD OF SPLICING SOLID “I” SPARS 1- 18 FIGURE 1- 7..........................................REPAIRS TO BUILT-UP “I” SPAR 1- 19 FIGURE 1-8.METHOD OF SPLICING BOX SPAR FLANGES (PLATE METHOD) 1- 20

1-41................................................................................................... SPAR REPLACEMENT 1- 16

1-42................................................................................... SPLICING OF BOX SPAR WEBS 1- 16 FIGURE 1-9. .........................METHOD OF SPLICING BOX SPAR WEBS 1- 21

**1-43.........................REPLACING SOLID-TYPE SPARS WITH LAMINATED-TYPE SPARS 1- 16**

### Page 6

Contents Paragraph Page

Page vi

1-44..............................................SPAR LONGITUDINAL CRACKS AND LOCAL DAMAGE

. … 1-16

**FIGURE 1-10..METHOD OF REIN FORCING A LONGITUDINAL CRACK**

AND/OR LOCAL DAMAGE IN A SOLID SPAR … 1-22

1-45.................................................................................... ELONGATED HOLES IN SPARS

. … 1-18

1-46...................................................................................................................RIB REPAIRS

. … 1-18

FIGURE 1- 11..................................................... REPAIR OF WOOD RIBS 1-23 FIGURE 1-12.................T YPICAL WING COMPRESSION RIB REPAIRS 1-24

1-47............................................................................................... PLYWOOD SKIN REPAIR

. … 1-20

**1-48......................................... DETERMINATION OF SINGLE OR DOUBLE CURVATURE**

. … 1-20

1-49.................................................REPAIRS TO SINGLE CURVATURE PLYWOOD SKIN

. … 1-20

**TABLE 1-2.MINIMUM RECOMME NDED BEND RADII FOR AIRCRAFT PLYWOOD 1-25**

1-50............................................... REPAIRS TO DOUBLE CURVATURE PLYWOOD SKIN

. … 1-22

1-51......................................................................................................TYPES OF PATCHES

. … 1-25

FIGURE 1-13. ................................................................SPLAYED PATCH 1-27 FIGURE 1-14........................................................... SURFACE PATCHES 1-28 FIGURE 1-15............SCARF PATCH ES (BACK OF SKIN ACCESSIBLE) 1-29 FIGURE 1-16... SCARF PATCHES (BACK OF SKIN NOT ACCESSIBLE) 1-30 FIGURE 1-17. ....................................... OVAL PLUG PATCH ASSEMBLY 1-32 FIGURE 1- 18.....................................ROUND PLUG PATCH ASSEMBLY 1-33

1-52...............................................................................................................FABRIC PATCH

. … 1-31

1-53.1-63. [RESERVED.] … 1-33

**SECTION 5. FINISHING WOOD STRUCTURES**

1-64 … GENERAL

. … 1-35

1-65................................................................................................ ACCEPTABLE FINISHES

. … 1-35

1-66 … PRECAUTIONS

. … 1-35

1-67..........................................................................FINISHING OF INTERIOR SURFACES

. … 1-36

1-68.........................................................................FINISHING OF EXTERIOR SURFACES

. … 1-36

1-69.......................................................................FINISHING OF END GRAIN SURFACES

. … 1-36

### Page 7

Contents Paragraph Page

Page vii

1-70..............................................................................FINISHING WITH FABRIC OR TAPE

. … 1-37

1-71............................................................................................ SEALING OF BOLT HOLES

. … 1-37

1-72.1-79. [RESERVED.] … 1-37

**CHAPTER 2. FABRIC COVERING SECTION 1. PRACTICES AND PRECAUTIONS**

2-1 … GENERAL

. … 2-1

2-2.............................................................................................................PROBLEM AREAS

. … 2-1

2-3..................................................................................... AIRCRAFT FABRIC-SYNTHETIC

. … 2-1

2-4......................................................................................... AIRCRAFT FABRIC-NATURAL

. … 2-2

**TABLE 2-1. ................................................ COTTON AND LINEN FABRICS 2-3**

2-5................................................................................................ RECOVERING AIRCRAFT

. … 2-2

**TABLE 2-2. ........................ CO TTON AND LINEN, TAPES AND THREADS 2-4**

2-6...............................................PREPARATION OF THE STRUCTURE FOR COVERING

. … 2-2

2-7.................................................................................................................FABRIC SEAMS

. … 2-5

FIGURE 2-1. ........................................................................FABRIC SEAMS 2-7 FIGURE 2-2....................... T YPICAL METHODS OF ATTACHING FABRIC 2-8

2-8..................................................................................................... COVERING METHODS

. … 2-9

### Page 8

Contents (continued) Paragraph Page

Page viii

2-9........................................................................................................REINFORCING TAPE

. … 2-10

**FIGURE 2- 3............................................ EXPLODED SIDE VIEW OF RIB 2-10**

2-10............................................................................................................................L A C I N G

. … 2-11

**FIGURE 2-4.STANDARD EXTERNAL MO DIFIED SEINE KNOT USED FOR**

SINGLE AND DOUBLE RIB LACING … 2-11

FIGURE 2- 5.................................. STARTING STITCH FOR RIB LACING 2-12 FIGURE 2-6. ....................................STANDARD SINGLE-LOOP LACING 2-13 FIGURE 2-7............... ST ANDARD KNOT FOR DOUBLE-LOOP LACING 2-14 FIGURE 2-8............. ST ANDARD DOUBLE-LOOP LACING (OPTIONAL) 2-15 FIGURE 2-9A.ALTERNATE SEQUENCE TO TIE A MODIFIED SEINE KNOT

FOR RIB LACING … 2-16

**FIGURE 2-9B.ALTERNATE SEQUENCE TO TIE A MODIFIED SEINE KNOT**

FOR RIB LACING … 2-17

**FIGURE 2-9C.ALTERNATE SEQUENCE TO TIE A MODIFIED SEINE KNOT**

FOR RIB LACING … 2-18

FIGURE 2-10. ......................................................................SPLICE KNOT 2-19 FIGURE 2-11.BLINDSTITCH LACING - SQUARE KNOT SECURED WITH HALF HITCHES 2-20

2-11........................................................................................................... STITCH SPACING

. … 2-19

**FIGURE 2-12. ..................................... FABRIC ATTACHMENT SPACING 2-21**

2-12 … FASTENERS

. … 2-21

2-13.............................................................................................................FINISHING TAPE

. … 2-22

2-14............................................................INSPECTION RINGS AND DRAIN GROMMETS

. … 2-22

2-15.2-19. [RESERVED.] … 2-23

**SECTION 2. APPLICATION OF DOPE**

2-20 … GENERAL

. … 2-25

TABLE 2-3............................................SAFET Y TIPS FOR DOPE/PAINT 2-26 TABLE 2-4................................................................... TIPS FOR DOPING 2-26

2-21......................................... DOPE APPLICATION PROCEDURE (NATURAL FABRICS)

. … 2-26

2-22........................................................................................COVERING OVER PLYWOOD

. … 2-28

2-23..............................................................................COATING APPLICATION DEFECTS

. … 2-29

2-24.2-29. [RESERVED.] … 2-31

**SECTION 3. INSPECTION AND TESTING**

### Page 9

Contents (continued) Paragraph Page

Page ix

2-30 … GENERAL

. … 2-33

2-31...............................................................................................FABRIC IDENTIFICATION

. … 2-34

2-32...........................................................................................COATING IDENTIFICATION

. … 2-34

2-33..........................................................STRENGTH CRITERIA FOR AIRCRAFT FABRIC

. … 2-35

2-34............................................................................................................FABRIC TESTING

. … 2-35

2-35................................................................................... REJUVENATION OF DOPE FILM

. … 2-36

2-36.2-41. [RESERVED.] … 2-36

**SECTION 4. REPAIRS TO FABRIC COVERING**

2-42 … GENERAL

. … 2-37

2-43............................................................ REPAIR OF TEARS AND ACCESS OPENINGS

. … 2-37

**FIGURE 2-13 ........................................... REPAIR OF TEARS IN FABRIC 2-38**

2-44...................................................................................................SEWN-PATCH REPAIR

. … 2-38

FIGURE 2- 14........................................................ HAND-STITCH DETAIL 2-39

2-45..........................................................................................DOPED-ON PATCH REPAIR

. … 2-39

2-46.2-51. [RESERVED.] … 2-40

### Page 10

Contents (continued) Paragraph Page

Page x Chapter 3. Fiberglass And Plastics SECTION 1. REPAIR OF LIGHT LOAD LAMINATE STRUCTURES

3-1 … GENERAL

. … 3-1

3-2.................................................................................FIBERGLASS LAMINATE REPAIRS

. … 3-1

FIGURE 3-1. ..............................TYPICAL LAMINATE (FACING) REPAIR 3-2 FIGURE 3- 2................................ TYPICAL CORE AND FACING REPAIR 3-2 FIGURE 3-3. .....................................TYPICAL STEPPED JOINT REPAIR 3-3

3-3..........................................................................................................REPAIRING HOLES

. … 3-3

FIGURE 3-4....................... PREPARING THE FIBERGLASS SANDWICH 3-3 FIGURE 3-5.SCARFED REPAIR TO A NONSTRUCTURAL LAMINATED

FIBERGLASS COMPONENT … 3-4

**FIGURE 3-6.BALANCED LAYUP SAME AS ORIGINAL NUMBER OF PLIES,**

PLUS ONE EXTRA PLY … 3-4

FIGURE 3-7. ......................................... TYPICAL SCARF JOINT REPAIR 3-5 FIGURE 3-8.CAREFULLY CUT THROUGH EACH LAYER OF FIBERGLASS

CLOTH AND ...............................................REMOVE IT FROM THE DAMAGED AREA 3-5

3-4................................................................ SAMPLE BAGGING AND CURING PROCESS

. … 3-5

**FIGURE 3-9.................... SAMPLE BA GGING LAYUP CROSS SECTION 3-6**

3-5.3-9. [RESERVED.] … 3-8

**SECTION 2. METALLIC SANDWICH SECONDARY STRUCTURE REPAIRS**

**3-10.............................REPAIRS TO METALLIC SANDWICH SECONDARY STRUCTURE**

. … 3-9

FIGURE 3-10.....................T YPICAL TYPES OF CORE REPLACEMENT 3-10 FIGURE 3-11............. T YPICAL UNDERCUT CORE MATERIAL CAVITY 3-10 FIGURE 3-12.CROSS SECTION OF BONDED AND BOLTED OVERLAP REPAIRS 3-11 FIGURE 3-13. ........................................HONEYCOMB CORE REMOVAL 3-11 FIGURE 3-14..... REMOVING HON EYCOMB CORE FROM A TAPERED

CONTROL SURFACE … 3-11

FIGURE 3-15. .........................REMOVAL OF CORE WITH CORE KNIFE 3-11 FIGURE 3-16. ...............................................DETAILS OF CORE REPAIR 3-12

3-11 … FINISHING

. … 3-12

3-12.3-17. [RESERVED.] … 3-13

**SECTION 3. TRANSPARENT PLASTICS**

### Page 11

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Page xi

3-18 … GENERAL

. … 3-15

3-19............................................................................................STORAGE AND HANDLING

. … 3-15

3-20............................................................ FORMING PROCEDURES AND TECHNIQUES

. … 3-15

3-21.........................................................................................................................HEATING.

. … 3-16

**FIGURE 3-17. ...................................... HANGING OF ACRYLIC SHEETS 3-16**

3-22 … FORMS

. … 3-16

3-23..................................................................................................... FORMING METHODS

. … 3-17

TABLE 3-1.TYPICAL TEMPERATURES FOR FORMING ACRYLIC SHEETS 3-17 FIGURE 3-18.DRILL HAVING AN INCLUDED ANGLE OF APPROXIMATELY 150°, ...................................................... USED TO DRILL ACRYLIC PLASTICS 3-18 FIGURE 3-19..........UNIBIT DRILL FO R DRILLING ACRYLIC PLASTICS 3-18 FIGURE 3-20.............. APPLYING PR ESSURE TO ACRYLIC PLASTICS 3-19

### Page 12

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Page xii

3-24....................................................................................................REPAIR OF PLASTICS

. … 3-19

FIGURE 3- 21...................................................STOP-DRILLING CRACKS 3-20 FIGURE 3-22........................................................... SURFACE PATCHES 3-20 FIGURE 3- 23......................................................... PLUG PATCH REPAIR 3-21

3-25..............................................CLEANING AND POLISHING TRANSPARENT PLASTIC

. … 3-21

3-26............................................................................................... REPLACEMENT PANELS

. … 3-21

3-27..................................................................................... INSTALLATION PROCEDURES

. … 3-22

3-28.3-39. [RESERVED.] … 3-22

**SECTION 4. WINDSHIELDS, ENCLOSURES, AND WINDOWS**

3-40 … GENERAL

. … 3-23

**FIGURE 3-24.TEMPORARY REPAIRS TO CRACKED WINDSHIELDS OR WINDOWS 3-24**

3-41 … PROTECTION

. … 3-24

3-42........................................................................................WINDSHIELD INSTALLATION

. … 3-24

3-43.3-47. [RESERVED.] … 3-25

Chapter 4. Metal Structure, Welding, And Brazing SECTION 1. IDENTIFICATION OF METALS

4-1 … GENERAL

. … 4-1

**TABLE 4-1.................................. FE RROUS (IRON) ALLOY MATERIALS 4-1**

4-2...............................................................................IDENTIFICATION OF STEEL STOCK

. … 4-1

TABLE 4-2.......... NUMERICAL SYSTEM FOR STEEL IDENTIFICATION 4-2 TABLE 4-3.EXAMPLES OF STAINL ESS AND HEAT-RESISTANT STEELS

NOMINAL COMPOSITION (PERCENT) … 4-2

4-3...................................................................INTERCHANGEABILITY OF STEEL TUBING

. … 4-2

4-4.................................................................................... IDENTIFICATION OF ALUMINUM

. … 4-2

**TABLE 4-4.BASIC TEMPER DESIGNAT IONS AND SUBDIVISIONS FROM**

ALUMINUM ALLOYS … 4-3

4-5.4-15. [RESERVED.] … 4-3

**SECTION 2. TESTING OF METALS**

4-16.....................................................................................................HARDNESS TESTING

. … 4-5

4-17......................................................................................ROCKWELL HARDNESS TEST

. … 4-5

4-18........................................................................................... BRINELL HARDNESS TEST

. … 4-5

### Page 13

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Page xiii

4-19.......................................................................................... VICKERS HARDNESS TEST

. … 4-6

4-20.........................................................................................MICROHARDNESS TESTING

. … 4-6

4-21 … INDENTIONS

. … 4-7

**FIGURE 4-1.COMPARISON OF INDENTATION MADE BY KNOOP AND**

**VICKERS INDENTERS ..................... IN THE SAME WORK META L AND AT THE SAME LOADS 4-7**

4-22......................................................................................................MAGNETIC TESTING

. … 4-8

4-23......................................................................................................ALUMINUM TESTING

. … 4-8

**TABLE 4-5.HARDNESS VALUES FOR ALUMINUM ALLOYS. (REFERENCE MIL-H-6088G.) 4-8**

4-24.4-35. [RESERVED.] … 4-8

### Page 14

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Page xiv SECTION 3. PRECAUTIONARY MEASURES

4-36.................................................................FLUTTER AND VIBRATION PRECAUTIONS

. … 4-9

4-37....................................................................................LOAD FACTORS FOR REPAIRS

. … 4-10

4-38................................................... TRANSFER OF STRESSES WITHIN A STRUCTURE

. … 4-10

4-39.4-49. [RESERVED.] … 4-10

**SECTION 4. METAL REPAIR PROCEDURES**

4-50 … GENERAL

. … 4-11

**4-51................................... RIVETED (OR BOLTED) STEEL TRUSS-TYPE STRUCTURES**

. … 4-11

4-52................................................................................ALUMINUM ALLOY STRUCTURES

. … 4-11

**4-53.........................................SELECTION OF ALUMINUM FOR REPLACEMENT PARTS**

. … 4-12

4-54.....................................................HEAT TREATMENT OF ALUMINUM ALLOY PARTS

. … 4-12

4-55............................................................................................................ BENDING METAL

. … 4-13

**TABLE 4-6.RECOMMENDED RADII FO R 90-DEGREE BENDS IN ALUMINUM ALLOYS 4-14**

4-56 … SETBACK

. … 4-13

FIGURE 4-2. .................................. SETBACK FOR A 90-DEGREE BEND 4-13 TABLE 4-7.K-CHART FOR DETERMINING SETBACK FOR BENDS OTHER

THAN 90 DEGREES … 4-15

**FIGURE 4-3.METHODS OF DETERMINING SETBACKS FOR BENDS OTHER**

THAN 90 DEGREES … 4-16

4-57 … RIVETING

. … 4-14

FIGURE 4-4.RIVET IDENTIFICA TION AND PART NUMBER BREAKDOWN 4-16 TABLE 4-8....................................... AIRCRAFT RIVET IDENTIFICATION 4-17 FIGURE 4-5.RIVET HOLE SPACING A ND EDGE DISTANCE FOR SINGLE-LAP

SHEET SPLICES … 4-20

FIGURE 4-6......... RIVETING PRACTICE AND RIVET IMPERFECTIONS 4-21 FIGURE 4-7........SELF PLUGGING FRICTION-LOCK CHERRY RIVETS 4-22 FIGURE 4-8.MECHANICAL-LOCK (BULBED CHERRY) CHERRY RIVET 4-23 FIGURE 4-9.............................................................CHERRY MAX RIVET 4-24 FIGURE 4- 10.............................................................OLYMIC-LOK RIVET 4-25 FIGURE 4- 11.........................................................................HUCK RIVET 4-26

### Page 15

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Page xv

**4-58.................REPAIR METHODS AND PRECAUTIONS FOR ALUMINUM STRUCTURE**

. … 4-25

**FIGURE 4-12.TYPICAL REPAIR ME THOD FOR TUBULAR MEMBERS OF**

ALUMINUM ALLOY … 4-28

FIGURE 4-13.TYPICAL REPAIR FOR BUCKLED OR CRACKED METAL WING RIB CAPSTRIPS 4-29 FIGURE 4-14.TYPICAL METAL RIB REPAIRS (USUALLY FOUND ON SMALL

AND MEDIUM-SIZE AIRCRAFT) … 4-30

FIGURE 4-15. ........................ TYPICAL REPAIRS OF TRAILING EDGES 4-31 FIGURE 4-16.TYPICAL REPAIRS OF STRESSED SHEET METAL COVERINGS 4-32 FIGURE 4-17................... TYPICAL ST RINGER AND FLANGE SPLICES 4-33 FIGURE 4-18.EXAMPLE OF STRING ER SPLICE (MATERIAL-2017 ALLOY) 4-34 FIGURE 4-19........ APPLICATION OF TYPICAL FLANGE SPLICES AND

REINFORCEMENT … 4-35

TABLE 4-9.NUMBER OF RIVETS REQU IRED FOR SPLICES (SINGLE-LAP JOINT) IN BARE 2014-T6, 2024-T3, 2024-T3, A ND 7075-T6 SHEET, CLAD 2014-T6, 2024-T3, 2024-T36, AND 7075-T6 SHEET, 2024-T4, AND 7075-T6 PLATE, ............. BAR, ROD, TUBE, AND EXTRUSIONS, 2014-T6 EXTRUSIONS 4-37 TABLE 4-10.NUMBER OF RIVETS RE QUIRED FOR SPLICES (SINGLE-LAP JOINT) . IN 2017, 1017 ALCLAD, 2024-T3 ALCL AD SHEET, PLATE, BAR, ROD, .........................................................................TUBE, A ND EXTRUSIONS 4-38 TABLE 4-11.NUMBER OF RIVETS RE QUIRED FOR SPLICES (SINGLE-LAP JOINT) ....................................................... IN 5032 (ALL HARDNESSES) SHEET 4-39

### Page 16

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Page xvi

4-59................................................................................REPAIRING CRACKED MEMBERS

. … 4-40

**FIGURE 4-20.EXAMPLE OF INTERM EDIATE FRAME STRINGER SPLICE**

(MATERIAL 2017-T3 AL ALLOY) … 4-41

**FIGURE 4-21.TYPICAL METHODS OF REPAIRING CRACKED LEADING AND**

**TRAILING EDGES**

AND RIB INTER SECTIONS … 4-42

FIGURE 4-22.TYPICAL METHODS OF REPLACING CRACKED MEMBER AT FITTINGS 4-43 FIGURE 4-23.TYPICAL METHODS OF REPAIRING CRACKED FRAME AND

STIFFENER COMBINATION … 4-44

**FIGURE 4-24.TYPICAL REPAIRS TO RUDDER AND TO FUSELAGE AT TAIL POST 4-45**

4-60................................................................................STEEL AND ALUMINUM FITTINGS

. … 4-40

**FIGURE 4-25.TYPICAL METHODS OF REPAIRING ELONGATED OR WORN BOLT HOLES 4-46**

4-61 … CASTINGS

. … 4-40

4-62.................................................. SELECTIVE PLATING IN AIRCRAFT MAINTENANCE

. … 4-40

4-63.4-73. [RESERVED.] … 4-51

**SECTION 5. WELDING AND BRAZING**

4-74 … GENERAL

. … 4-53

TABLE 4-12.CURRENT AND POLARITY SELECTION FOR INERT GAS WELDING 4-54 FIGURE 4-26.COMMON DEFECTS TO AVOID WHEN FITTING AND WELDING

AIRCRAFT CERTIFICATION CLUSTER … 4-55

TABLE 4-13.................................................................. TORCH TIP SIZES 4-55 FIGURE 4-27................. BU TT-WELDING STRENGTH CALCULATIONS 4-56

4-75......................................................................................................TYPES OF WELDING

. … 4-56

FIGURE 4-28.BASIC GAS-WELDING FLAMES: EACH HAS DISTINCTIVE SHAPE, COLOR AND SOUND.

................................................... NEUT RAL FLAME IS THE MOST USED 4-57 FIGURE 4-29.SET TIG WELDER TO DC CURRENT, STRAIGHT POLARITY

FOR WELDING MILD STEEL, ..........................................................STAINLESS STEEL AND TITANIUM 4-57 FIGURE 4-30.SET TIG TO AC CURRE NT FOR WELDING ALUMINUM AND MAGNESIUM 4-58

4-76............................................................................. ELECTRIC-RESISTANCE WELDING

. … 4-58

**FIGURE 4-31.IN SPOT WELDING, HEAT IS PRODUCED BY ELECTRICAL**

**RESISTANCE BETWEEN COPPER ELECTRODES. PRESSURE IS SIMULTANEOUSLY APPLIED TO**

**ELECTRODE TIPS TO FORCE METAL TOGETHER TO COMPLETE FUSING PROCESS. SPOT-**

**WELD-NUGGET SIZE**

### Page 17

Contents (continued) Paragraph Page

Page xvii .......................................................IS DIRECT LY RELATED TO TIP SIZE 4-58

4-77 … BRAZING

. … 4-58

TABLE 4-14.CALCULATED ALLO WABLE STRENGTH OF BASE METAL 4-59 FIGURE 4-32. ..................................................SILVER BRAZING JOINTS 4-60 FIGURE 4-33. ........................................... ELECTRIC SOLDERING IRON 4-61

4-78..................................................................... AIRCRAFT PARTS NOT TO BE WELDED

. … 4-61

4-79..........................................................................................WELDING ROD SELECTION

. … 4-62

**TABLE 4-15....CHART SHOWING WELDING FILLER ROD SELECTION 4-62**

4-80.................................................................................REPAIR OF TUBULAR MEMBERS

. … 4-62

**FIGURE 4-34.FINGER PATCH REPAIRS FOR MEMBERS DENTED AT A CLUSTER 4-63**

4-81.......................................................................................REPAIR BY WELDED SLEEVE

. … 4-64

**FIGURE 4-35.MEMBERS DENTED IN A BAY (REPAIRS BY WELDED SLEEVE) 4-65**

4-82........................................................................................REPAIR BY BOLTED SLEEVE

. … 4-64

4-83..............................................................................................WELDED-PATCH REPAIR

. … 4-64

**FIGURE 4- 36. ................................................. WELDED PATCH REPAIR 4-66**

4-84........................................................SPLICING TUBING BY INNER-SLEEVE METHOD

. … 4-64

**FIGURE 4-37. ...........................SPLICING BY INNER-SLEEVE METHOD 4-67**

4-85...................................................... SPLICING TUBING BY OUTER-SLEEVE METHOD

. … 4-66

**FIGURE 4-38.SPLICING BY OUTE R-SLEEVE METHOD (REPLACEMENT BY**

WELDED OUTSIDE SLEEVE) … 4-68

**FIGURE 4-39.TUBE REPLACEMENT AT A STATION BY WELDED OUTER SLEEVES 4-69**

### Page 18

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Page xviii

**4-86............................... SPLICING USING LARGER DIAMETER REPLACEMENT TUBES**

. … 4-67

**FIGURE 4-40.SPLICING USING LARG ER DIAMETER REPLACEMENT TUBE 4-69**

4-87.............................................................. REPAIRS AT BUILT-IN FUSELAGE FITTINGS

. … 4-70

**FIGURE 4-41................... REPAIRS AT BUILT-IN FUSELAGE FITTINGS 4-71**

4-88............................................................................................ ENGINE-MOUNT REPAIRS

. … 4-70

4-89................................................................BUILT-UP TUBULAR WING OR TAIL-SPARS

. … 4-72

**FIGURE 4-42......... STREAMLINE TUBE SPLICE USING ROUND TUBE**

(APPLICABLE TO LANDING GEAR) … 4-73

4-90....................................................WING-BRACE STRUTS AND TAIL-BRACE STRUTS

. … 4-72

FIGURE 4-43........ STREAMLINE TUBE SPLICE USING SPLIT SLEEVE (APPLICABLE TO WING AND TAIL SURFACE BRACE STRUTS AND OTHER MEMBERS) 4-74

4-91................................................................................................LANDING GEAR REPAIR

. … 4-72

**FIGURE 4-44......... STREAMLINE TUBE SPLICE USING SPLIT INSERT**

(APPLICABLE TO LANDING GEAR) … 4-75

**FIGURE 4-45.STREAMLINE TUBE SPL ICE USING PLATES (APPLICABLE TO**

LANDING GEAR) … 4-76

FIGURE 4-46.REPRESENTATI VE TYPES OF REPAIRABLE AXLE ASSEMBLIES 4-77 FIGURE 4-47.LANDING GEAR ASSEMBLIES THAT CANNOT BE REPAIRED BY WELDING 4-78

4-92...........................................................................REPAIRS TO WELDED ASSEMBLIES

. … 4-73

4-93..................................................................................STAINLESS STEEL STRUCTURE

. … 4-74

4-94.4-103. [RESERVED.] … 4-78

**SECTION 6. WELDING AND BRAZING SAFETY**

4-104 … GENERAL

. … 4-79

4-105..................................................................................FIRE AND EXPLOSION SAFETY

. … 4-79

4-106................................................................................................WELDING WORK AREA

. … 4-79

4-107.......................................................................................................FIRE PROTECTION

. … 4-79

4-108...............................................................................................PROTECTIVE APPAREL

. … 4-80

4-109.............................................................................................................FIRST-AID KITS

. … 4-80

4-110.4-128. [RESERVED.] … 4-80

Chapter 5. NonDestructive Inspection (Ndi) SECTION 1. GENERAL

### Page 19

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Page xix

5-1 … GENERAL

. … 5-1

5-2.............................................................................................APPROVED PROCEDURES

. … 5-1

5-3......................................................................................................................NDT LEVELS

. … 5-1

5-4...................................................... TRAINING, QUALIFICATION, AND CERTIFICATION

. … 5-2

5-5 … FLAWS

. … 5-2

5-6.......................................................................................SELECTING THE NDI METHOD

. … 5-5

**TABLE 5-1. ADVANTAGES AND DI SADVANTAGES OF NDI METHODS 5-7**

5-7................................................................................................ TYPES OF INSPECTIONS

. … 5-6

5-8.5-14. [RESERVED.] … 5-7

**SECTION 2. VISUAL INSPECTION**

5-15 … GENERAL

. … 5-9

5-16..............................................................................SIMPLE VISUAL INSPECTION AIDS

. … 5-9

5-17 … BORESCOPES

. … 5-10

**FIGURE 5-1. ....................................... TYPICAL BORESCOPE DESIGNS 5-11**

### Page 20

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Page xx

5-18........................................................................... VISUAL INSPECTION PROCEDURES

. … 5-10

**FIGURE 5-2...........USING A FLASH LIGHT TO INSPECT FOR CRACKS 5-13**

5-19.5-24. [RESERVED.] … 5-13

**SECTION 3. EDDY CURRENT INSPECTION**

5-25......................................................................................EDDY CURRENT INSPECTION

. … 5-15

FIGURE 5-3. ....................................GENERATING AN EDDY CURRENT 5-15 FIGURE 5-4. .......................................DETECTING AN EDDY CURRENT 5-15 FIGURE 5-5. ..................................... TYPICAL INSTRUMENT DISPLAYS 5-16

5-26........................................................................ EDDY CURRENT COILS AND PROBES

. … 5-15

5-27........................................... FIELD APPLICATION OF EDDY CURRENT INSPECTION

. … 5-15

5-28................................................................................................. SURFACE INSPECTION

. … 5-16

**FIGURE 5-6. .............................................. TYPICAL SURFACE CRACKS 5-16**

5-29..........................................................................................SUBSURFACE INSPECTION

. … 5-17

**FIGURE 5-7. .......................................TYPICAL SUBSURFACE CRACKS 5-18**

5-30............................................................................................CORROSION INSPECTION

. … 5-17

**FIGURE 5-8. ................................TYPICAL STRUCTURAL CORROSION 5-18**

**5-31................................ESTABLISHING EDDY CURRENT INSPECTION PROCEDURES**

. … 5-17

5-32.5-39. [RESERVED.] … 5-18

**SECTION 4. MAGNETIC PARTICLE INSPECTION**

5-40 … GENERAL

. … 5-19

5-41........................................................................................PRINCIPLES OF OPERATION

. … 5-19

FIGURE 5-9. ........................................... MAGNETIC FIELD DISRUPTED 5-19 FIGURE 5-10.CRACK DETECTION BY MAGNETIC PARTICLE INSPECTION 5-19

5-42 … APPLICATIONS

. … 5-20

5-43...................................................................ELECTRICAL MAGNETIZING EQUIPMENT

. … 5-20

5-44....................................... MATERIALS USED IN MAGNETIC PARTICLE INSPECTION

. … 5-20

TABLE 5-2... LISTING OF COMMONLY ACCEPTED STANDARDS AND ................SPECIFICATIONS FOR MA GNETIC PARTICLE INSPECTION 5-21

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Page xxi

5-45....................................................................................... PREPARATION OF SURFACE

. … 5-22

5-46........................................................................................METHODS OF EXAMINATION

. … 5-22

5-47.................................................................. APPLICATION OF MAGNETIC PARTICLES

. … 5-22

5-48 … MAGNETIZATION

. … 5-23

FIGURE 5-11. ............................................ CIRCULAR MAGNETIZATION 5-23 FIGURE 5-12. .................................... LONGITUDINAL MAGNETIZATION 5-24

5-49.....................................................................DETERMINATION OF FIELD STRENGTH

. … 5-25

5-50........................................................................SPECIAL EXAMINATION TECHNIQUES

. … 5-25

**5-51......................................DEMAGNETIZATION AND POST-EXAMINATION CLEANING**

. … 5-27

5-52.5-59. [RESERVED.] … 5-28

**SECTION 5. PENETRANT INSPECTION**

5-60 … GENERAL

. … 5-29

**FIGURE 5- 13.......................... PENETRANT AND DEVELOPER ACTION 5-30**

**5-61........................... EQUIPMENT USED IN THE PENETRANT INSPECTION PROCESS**

. … 5-29

**TABLE 5-3.CLASSIFICATION OF PENETRANT INSPECTION MATERIALS**

COVERED BY MIL-I-25135E … 5-30

### Page 22

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Page xxii

**5-62.......................................BASIC STEPS TO PERFORM PENETRATION INSPECTION**

. … 5-29

TABLE 5-4.. FLUORESCENT AND VI SIBLE PENETRANT INSPECTION ........................GENERAL PR OCESSING PROCEDURES FLOWSHEET 5-31 TABLE 5-5.PRE-CLEANING METH ODS FOR PENETRANT INSPECTION 5-32

5-63 … CLEANERS

. … 5-33

5-64. TECHNICAL STANDARDS … 5-34

**TABLE 5-6... LISTING OF COMMONLY ACCEPTED STANDARDS AND**

SPECIFICATIONS FOR ........................................................................PENET RANT INSPECTION 5-35

5-65.5-72. [Reserved.] … 5-35

**SECTION 6. RADIOGRAPHY (X-RAY) INSPECTION**

5-73 … GENERAL

. … 5-37

FIGURE 5-14 … RADIOGRAPHY

5-38

5-74 … LIMITATIONS

. … 5-37

5-75..................................................................................FILM OR PAPER RADIOGRAPHY

. … 5-37

5-76...........................................................................................REAL-TIME RADIOGRAPHY

. … 5-37

**5-77.............ADVANTAGE OF REAL-TIME RADIOGRAPHY OVER FILM RADIOGRAPHY**

. … 5-37

5-78.................................................................................COMPUTED TOMOGRAPHY (CT)

. … 5-37

5-79..............................................................................................USES OF RADIOGRAPHY

. … 5-38

5-80...............................................................COMPARISON WITH OTHER NDI METHODS

. … 5-38

5-81 … FLAWS

. … 5-39

5-82........................................................................................................ FIELD INSPECTION

. … 5-39

5-83 … SAFETY

. … 5-39

5-84.5-88. [RESERVED.] … 5-39

**SECTION 7. ULTRASONIC INSPECTION**

5-89 … GENERAL

. … 5-41

FIGURE 5- 15 … ULTRASOUND

5-42

5-90.....................................................................................................SOUND REFLECTION

. … 5-41

5-91..................................................................ULTRASONIC INSPECTION TECHNIQUES

. … 5-41

**FIGURE 5-16.PULSE-ECHO AND T HROUGH-TRANSMISSION ULTRASONIC**

INSPECTION TECHNIQUES … 5-42

### Page 23

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Page xxiii

5-92..........................................................................................................FLAW DETECTION

. … 5-41

5-93........................................................................................................ BASIC EQUIPMENT

. … 5-41

FIGURE 5-17.TYPICAL PORTABLE ULTRASONIC INSPECTION INSTRUMENT 5-43 FIGURE 5-18.................EXAMPLE OF POSITION FIXTURE AND SHOE 5-44 FIGURE 5-19.EXAMPLE OF THE U SE IF A TRANSDUCER POSITIONING FIXTURE 5-44 FIGURE 5-20........EXAMPLE OF A TYPICAL REFERENCE STANDARD 5-44

5-94................................................................... INSPECTION OF BONDED STRUCTURES

. … 5-43

**FIGURE 5-21.EXAMPLES OF BONDED STRUCTURE CONFIGURATIONS**

AND SUGGESTED INSPECTION COVERAGE … 5-47

**TABLE 5-7........ ACCEPTABLE UL TRASONIC INSPECTION METHODS**

**ASSOCIATED WITH THE EXAMPLE BONDED STRUCTURE CONFIGURATIONS SHOWN IN FIGURE 5-**

**21. 5-48 TABLE 5-8.ULTRASONIC INSPECTION METHODS FOR BONDED STRUCTURES 5-49**

5-95................................................................................... BOND TESTING INSTRUMENTS

. … 5-46

5-96..................................................................................... THICKNESS MEASUREMENTS

. … 5-50

5-97................................................................................................................LEAK TESTING

. … 5-50

5-98.5-104. [RESERVED.] … 5-51

### Page 24

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Page xxiv SECTION 8. TAP TESTING

5-105 … GENERAL

. … 5-53

**FIGURE 5-22. ................................SAMPLE OF SPECIAL TAP HAMMER 5-53**

5-106.5-111. [RESERVED.] … 5-53

**SECTION 9. ACOUSTIC-EMISSION**

5-112 … GENERAL

. … 5-53

5-113 … APPLICATIONS

. … 5-54

5-114.5-119. [RESERVED.] … 5-54

**SECTION 10. THERMOGRAPHY**

5-120 … GENERAL

. … 5-54

5-121.5-126. [RESERVED.] … 5-54

**SECTION 11. HOLOGRAPHY**

5-127 … GENERAL

. … 5-54

5-128.5-133. [RESERVED.] … 5-54

**SECTION 12. SHEAROGRAPHY**

5-134 … GENERAL

. … 5-55

5-135.5-140. [RESERVED.] … 5-55

Chapter 6. Corrosion, Inspection & Protection SECTION 1. GENERAL

6-1 … GENERAL

. … 6-1

**FIGURE 6-1.SIMPLIFIED CORRO SION CELL SHOWING CONDITIONS**

**WHICH MUST EXIST FOR**

ELECTROCHEMICA L CORROSION … 6-2

**FIGURE 6-2. ELIMINATION OF CORROSION BY APPLICATION OF AN**

ORGANIC FILM TO METAL SURFACE … 6-2

6-2......................................................................... FACTORS INFLUENCING CORROSION

. … 6-1

6-3....................................................................................COMMON CORROSIVE AGENTS

. … 6-3

6-4 … MICRO-ORGANISMS

. … 6-3

6-5.6-10. [RESERVED.] … 6-4

**SECTION 2. TYPES OF CORROSION**

6-11 … GENERAL

. … 6-5

FIGURE 6- 3........................................................... CORROSION ATTACK 6-5

6-12...............................................................................GENERAL SURFACE CORROSION

. … 6-5

### Page 25

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Page xxv FIGURE 6-4. .....................................GENERAL SURFACE CORROSION 6-5

6-13.................................................................................................... PITTING CORROSION

. … 6-5

FIGURE 6-5(A). PITTING CORRO SION (EXTERNAL VIEW) … 6-5

**FIGURE 6-5(B). PITTING CORRO SION (MAGNIFIED CROSS SECTION) 6-5**

### Page 26

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Page xxvi

6-14.........................................................................CONCENTRATION CELL CORROSION

. … 6-6

**FIGURE 6-6. ...............................CONCENTRATION CELL CORROSION 6-6**

6-15............................................................................................... ACTIVE-PASSIVE CELLS

. … 6-6

FIGURE 6-7.........................................................ACTIVE-PASSIVE CELL 6-7

6-16................................................................................................. FILIFORM CORROSION

. … 6-7

FIGURE 6- 8........................................................ FILIFORM CORROSION 6-7

6-17................................................................................... INTERGRANULAR CORROSION

. … 6-7

**FIGURE 6-9. INTERGRANULAR CORROSION OF 7050-T6 ALUMINUM**

ADJACENT TO STEEL FASTENER … 6-8

6-18.......................................................................................... EXFOLIATION CORROSION

. … 6-7

**FIGURE 6- 10............................................... EXFOLIATION CORROSION 6-8**

6-19................................................................................................GALVANIC CORROSION

. … 6-8

**FIGURE 6-11.GALVANIC CORROSION OF MAGNESIUM ADJACENT TO**

STEEL FASTENER … 6-9

6-20................................................................................STRESS CORROSION CRACKING

. … 6-8

**FIGURE 6-12. ....................................STRESS CORROSION CRACKING 6-9**

6-21...................................................................................................FATIGUE CORROSION

. … 6-9

6-22................................................................................................ FRETTING CORROSION

. … 6-10

**FIGURE 6- 13..................................................... FRETTING CORROSION 6-10**

6-23.6-28. [RESERVED.] … 6-10

**SECTION 3. CORROSION PROTECTION MEASURES FOR BASIC MATERIALS**

6-29 … GENERAL

. … 6-11

6-30....................................................................ANODIZING AND RELATED PROCESSES

. … 6-11

6-31 … PLATING

. … 6-11

6-32....................................................................................PHOSPHATE RUST-PROOFING

. … 6-11

6-33....................................................................................CHROME-PICKLE TREATMENT

. … 6-11

6-34..........................................................................................DICHROMATE TREATMENT

. … 6-11

6-35......................................................................... STANNATE IMMERSION TREATMENT

. … 6-11

6-36...........................................................................GALVANIC ANODIZING TREATMENT

. … 6-12

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Page xxvii

6-37 … CLADDING

. … 6-12

6-38..........................................................................................................METAL SPRAYING

. … 6-12

6-39 … SHOT-PEENING

. … 6-12

6-40.....................................................................................................ORGANIC COATINGS

. … 6-12

6-41........................................................................................................... DOPE PROOFING

. … 6-12

6-42........................................................................................................... TUBE INTERIORS

. … 6-12

6-43.6-49. [RESERVED.] … 6-12

**SECTION 4. CORROSION PREVENTIVE MAINTENANCE**

6-50........................................................................................ GUIDELINES: ALL AIRCRAFT

. … 6-13

6-51......................................GUIDELINES: AIRCRAFT OPERATING OVER SALT WATER

. … 6-13

6-52.6-62. [RESERVED.] … 6-14

**SECTION 5. VISUAL CORROSION INSPECTION GUIDE FOR AIRCRAFT**

6-63 … GENERAL

. … 6-15

6-64................................................................................................ EXHAUST TRAIL AREAS

. … 6-15

**6-65................................BATTERY COMPARTMENTS AND BATTERY VENT OPENINGS**

. … 6-15

6-66.........................................................................................LAVATORIES AND GALLEYS

. … 6-15

6-67..................................................................................................................BILGE AREAS

. … 6-15

6-68..........................................................................WHEEL WELLS AND LANDING GEAR

. … 6-15

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Page xxviii

6-69................................................................................................EXTERNAL SKIN AREAS

. … 6-15

6-70.....................................................................................WATER ENTRAPMENT AREAS

. … 6-16

6-71.............................................................................................ENGINE FRONTAL AREAS

. … 6-16

6-72................................................................ELECTRONIC PACKAGE COMPARTMENTS

. … 6-16

6-73......................................................................................FLEXIBLE HOSE ASSEMBLIES

. … 6-16

6-74.......................................................................................................SANDWICH PANELS

. … 6-16

6-75.........................................................................................................CONTROL CABLES

. … 6-16

6-76.................................................................................................INTEGRAL FUEL CELLS

. … 6-16

6-77........................................................................................ ELECTRICAL CONNECTORS

. … 6-16

6-78.6-88. [RESERVED.] … 6-16

**SECTION 6. CORROSION REMOVAL PROCEDURES**

6-89 … GENERAL

. … 6-17

6-90................................................................................................ SAFETY PRECAUTIONS

. … 6-17

6-91..........................................................CORROSION CONTROL WORK PROCEDURES

. … 6-18

6-92.6-112. [RESERVED.] … 6-19

**SECTION 7. BASIC CORROSION REMOVAL TECHNIQUES**

6-113 … GENERAL

. … 6-21

6-114................................................................................. PREPARATIONS FOR REWORK

. … 6-21

6-115........................................................... FAIRING OR BLENDING REWORKED AREAS

. … 6-21

TABLE 6-1........................... ABRASIVES FOR CORROSION REMOVAL 6-22 FIGURE 6-14...... TYPICAL EXAMPLE OF ACCEPTABLE CLEANUP OF

CORROSION PITS … 6-23

FIGURE 6-15.BLENDOUT OF CORROSION AS A SINGLE DEPRESSION 6-24 FIGURE 6-16.BLENDOUT OF MULTIPLE PITS IN A CORRODED AREA 6-24

6-116......................................................................CORROSION REMOVAL BY BLASTING

. … 6-22

6-117...........................................................CLEANERS, POLISHES, AND BRIGHTENERS

. … 6-22

6-118.................................................................................................STANDARD METHODS

. … 6-23

6-119.6-131. [RESERVED.] … 6-24

**SECTION 8. ALUMINUM AND ALUMINUM ALLOYS**

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Page xxix

6-132 … GENERAL

. … 6-25

6-133...................................................SPECIAL TREATMENT OF ANODIZED SURFACES

. … 6-25

6-134.........................................................REPAIR OF ALUMINUM ALLOY SHEET METAL

. … 6-25

**6-135.CORROSION REMOVAL AROUND COUNTERSUNK FASTENERS IN ALUMINUM ALLOY**

. … 6-26

**6-136.EXAMPLES OF REMOVING CORROSION FROM ALUMINUM AND ALUMINUM ALLOYS**

. … 6-26

6-137.6-147. [RESERVED.] … 6-28

**SECTION 9. MAGNESIUM AND MAGNESIUM ALLOYS**

6-148 … GENERAL

. … 6-29

**6.149................... TREATMENT OF WROUGHT MAGNESIUM SHEETS AND FORGINGS**

. … 6-29

**6.150.REPAIR OF MAGNESIUM SHEET METAL AFTER EXTENSIVE CORROSION REMOVAL**

. … 6-29

6.151...............................................IN-PLACE TREATMENT OF MAGNESIUM CASTINGS

. … 6-30

**6.152................................. EXAMPLE OF REMOVING CORROSION FROM MAGNESIUM**

. … 6-30

6-153.6-163. [RESERVED.] … 6-33

### Page 30

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Page xxx SECTION 10. FERROUS METALS

6-164 … GENERAL

. … 6-35

6-165................................................SPECIAL TREATMENT OF HIGH STRENGTH STEEL

. … 6-35

6-166..........................................................SPECIAL TREATMENT OF STAINLESS STEEL

. … 6-35

**6-167...................... EXAMPLE OF REMOVING CORROSION FROM FERROUS METALS**

. … 6-36

6-168.6-178. [RESERVED.] … 6-36

**SECTION 11. OTHER METALS AND ALLOYS**

**6-179..................................NOBLE METAL COATINGS - CLEANUP AND RESTORATION**

. … 6-37

6-180................................................................................COPPER AND COPPER ALLOYS

. … 6-37

6-181.............................................................................TITANIUM AND TITANIUM ALLOYS

. … 6-37

6-182.6-192. [RESERVED.] … 6-38

**SECTION 12. PLATED PARTS**

6-193................................................................CHROMIUM AND NICKEL PLATED PARTS

. … 6-39

6-194........................................................................ CADMIUM AND ZINC PLATED PARTS

. … 6-39

6-195.6-205. [RESERVED.] … 6-39

**SECTION 13. CORROSION PROOFING OF LAND PLANES CONVERTED TO SEA PLANES**

6-206 … GENERAL

. … 6-41

6-207.....................................................................NECESSARY MINIMUM PRECAUTIONS

. … 6-41

6-208...............................................................................RECOMMENDED PRECAUTIONS

. … 6-41

6-209.6-219. [RESERVED.] … 6-42

**SECTION 14. HANDLING AND CARE OF AIRCRAFT RECOVERED FROM WATER**

**IMMERSION**

6-220 … GENERAL

. … 6-43

6-221...................................................... INITIAL FRESH WATER OR DETERGENT WASH

. … 6-43

6-222......................................................RECIPROCATING ENGINES AND PROPELLERS

. … 6-43

6-223 … AIRFRAME

. … 6-43

6-224.6-234. [RESERVED.] … 6-44

Chapter 7. Aircraft Hardware, Control Cables, And Turnbuckles SECTION 1. RIVETS

7-1 … GENERAL

. … 7-1

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Page xxxi

7-2.............................................................................................. MATERIAL APPLICATIONS

. … 7-1

7-3.7-13. [RESERVED.] … 7-2

**SECTION 2. SCREWS**

7-14 … GENERAL

. … 7-3

7-15................................................................................................ STRUCTURAL SCREWS

. … 7-3

7-16........................................................................................................ MACHINE SCREWS

. … 7-3

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Page xxxii

7-17... PANHEAD SCREWS (NAS600 THROUGH NAS606, NAS610 THROUGH NAS616, ....................................................................NAS623, AND NAS1402 THROUGH NAS1406)

. … 7-3

7-18...............................................................................................SELF-TAPPING SCREWS

. … 7-3

7-19............................................................................................................. WOOD SCREWS

. … 7-4

7-20.7-33. [RESERVED.] … 7-4

**SECTION 3. BOLTS**

7-34 … GENERAL

. … 7-5

7-35 … BOLTS

. … 7-5

7-36 … IDENTIFICATION

. … 7-5

**FIGURE 7- 1................................TYPICAL AIRCRAFT BOLT MARKINGS 7-6**

7-37.................................................................................................................GRIP LENGTH

. … 7-5

7-38..........................................................................LOCKING OR SAFETYING OF BOLTS

. … 7-6

7-39..........................................................................................................................BOLT FIT

. … 7-6

7-40 … TORQUES

. … 7-6

FIGURE 7-2................TO RQUE WRENCH WITH VARIOUS ADAPTERS 7-8 TABLE 7-1.........RECOMMENDED TORQUE VALUES (INCH-POUNDS) 7-9

**7-41.......................... STANDARD AIRCRAFT HEX HEAD BOLTS (AN3 THROUGH AN20)**

. … 7-7

7-42........................................................DRILLED HEAD BOLTS (AN73 THROUGH AN81)

. … 7-7

7-43...............................................................................................................ENGINE BOLTS

. … 7-7

7-44.........................................................................................CLOSE-TOLERANCE BOLTS

. … 7-9

**7-45.INTERNAL WRENCHING BOLTS (NAS144 THROUGH NAS158 AND NAS172 THROUGH**

NAS176) … 7-10

7-46.INTERNAL WRENCHING BOLTS (MS20004 THROUGH MS20024) AND SIX HOLE, ............................... DRILLED SOCKET HEAD BOLTS (AN148551 THROUGH AN149350)

. … 7-10

**7-47. TWELVE POINT, EXTERNAL WRENCHING BOLTS, (NAS624 THROUGH NAS644)**

. … 7-10

7-48...........................................................CLOSE-TOLERANCE SHEAR BOLTS (NAS434)

. … 7-10

7-49............................................................................................... NAS6200 SERIES BOLTS

. … 7-10

7-50......................................................................CLEVIS BOLTS (AN21 THROUGH AN36)

. … 7-10

7-51.............................................................................EYEBOLTS (AN42 THROUGH AN49)

. … 7-10

7-52.7-62. [RESERVED.] … 7-10

### Page 33

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Page xxxiii SECTION 4. NUTS

7-63 … GENERAL

. … 7-11

7-64.................................................................................................... SELF-LOCKING NUTS

. … 7-11

**TABLE 7-2.MINIMUM PREVAILING TORQUE VALUES FOR REUSED SELF- LOCKING NUTS 7-12**

7-65................................................................................NUT IDENTIFICATION - FINISHES

. … 7-11

7-66.....................................................................................................CASTLE NUT (AN310)

. … 7-12

7-67...........................................................................CASTELLATED SHEAR NUT (AN320)

. … 7-13

7-68................................................................................... PLAIN NUT (AN315 AND AN335)

. … 7-13

7-69........................................................................LIGHT HEX NUTS (AN340 AND AN345)

. … 7-13

7-70.......................................................................................................CHECKNUT (AN316)

. … 7-13

7-71....................................................................................................... WINGNUTS (AN350)

. … 7-13

7-72..................................................................................... SHEET SPRING NUTS (AN365)

. … 7-13

7-73.7-84. [RESERVED.] … 7-13

### Page 34

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Page xxxiv SECTION 5. WASHERS

7-85 … GENERAL

. … 7-15

7-86.........................................................................PLAIN WASHERS (AN960 AND AN970)

. … 7-15

7-87..........................................................................LOCKWASHERS (AN935 AND AN936)

. … 7-15

7-88....................................... BALL SOCKET AND SEAT WASHERS (AN950 AND AN955)

. … 7-15

7-89.....................................................................................TAPER PIN WASHERS (AN975)

. … 7-15

7-90.7-100. [RESERVED.] … 7-15

**SECTION 6. PINS**

7-101..................................................................................................................TAPER PINS

. … 7-17

7-102............................................................. FLATHEAD PINS (AN392 THROUGH AN406)

. … 7-17

7-103.................................................................................................COTTER PINS (AN380)

. … 7-17

7-104................................................................................................................ SPRING PINS

. … 7-17

7-105.................................................................................................QUICK-RELEASE PINS

. … 7-17

7-106.7-121. [RESERVED.] … 7-17

**SECTION 7. SAFETYING**

7-122 … GENERAL

. … 7-19

7-123............................................................................................................... SAFETY WIRE

. … 7-19

FIGURE 7-3...SECURING SCREWS , NUTS, BOLTS, AND SNAPRINGS 7-20 FIGURE 7-3A . ............................................... WIRE TWISTING BY HAND 7-20

7-124................................................................................ SAFETY-WIRING PROCEDURES

. … 7-20

7-125..............................................................................TWISTING WITH SPECIAL TOOLS

. … 7-21

**FIGURE 7- 4....................................USE OF A TYPICAL WIRE TWISTER 7-22**

7-126...............................................SECURING OIL CAPS, DRAIN COCKS, AND VALVES

. … 7-21

FIGURE 7-4A. ....SECURING OIL CAPS, DRAIN COCKS, AND VALVES 7-22 FIGURE 7-5. ........................................ SAFETY-WIRING PROCEDURES 7-23 FIGURE 7-5A. ..................................... SAFET Y-WIRING PROCEDURES 7-24 FIGURE 7-5B. ..................................... SAFET Y-WIRING PROCEDURES 7-25

7-127.................................................................................SECURING WITH COTTER PINS

. … 7-26

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Page xxxv FIGURE 7-6. .........................................SECURING WITH COTTER PINS 7-26 FIGURE 7-7.ALTERNATE METHOD FOR SECURING WITH COTTER PINS 7-26

7-128.7-139. [RESERVED.] … 7-26

**SECTION 8. INSPECTION AND REPAIR OF CONTROL CABLES AND TURNBUCKLES**

7-140 … GENERAL

. … 7-27

7-141....................................................................................................CABLE DEFINITIONS

. … 7-27

7-142........................................................................................................FLEXIBLE CABLES

. … 7-27

TABLE 7-3.FLEXIBLE CABLE CONSTRUCTION AND PHYSICAL PROPERTIES 7-28 FIGURE 7-8. ...................................FLEXIBLE CABLE CROSS SECTION 7-29

7-143............................................................................................NYLON-COATED CABLES

. … 7-27

7-144...............................................................................................NONFLEXIBLE CABLES

. … 7-28

TABLE 7-4...NONFLEXIBLE CABLE CONSTRUCTION AND PHYSICAL PROPERTIES 7-29 FIGURE 7- 9............................NONFLEXIBLE CABLE CROSS SECTION 7-29

7-145.............................................................................................CABLE SPECIFICATIONS

. … 7-29

7-146.................................................................................................CABLE PROOF LOADS

. … 7-29

7-147.......................................................................................REPLACEMENT OF CABLES

. … 7-30

### Page 36

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Page xxxvi

7-148.............................................. MECHANICALLY-FABRICATED CABLE ASSEMBLIES

. … 7-30

TABLE 7-5....... STRAIGHT-SHANK TERMINAL DIMENSIONS. (CROSS REFERENCE AN TO MS:

....... AN-666 TO MS2159, AN-667 TO MS 20667, AN-668 TO MS 20668, ...............................................................................AN- 669 TO MS 21260.)

7-31 FIGURE 7-10. .......................... INSERTION OF CABLE INTO TERMINAL 7-31 FIGURE 7-11............. GA UGING TERMINAL SHANK AFTER SWAGING 7-31 FIGURE 7- 12.................................................. TYPICAL CABLE SPLICES 7-32 FIGURE 7-13. ............................................. TYPICAL TERMINAL GAUGE 7-32 FIGURE 7-14. ........................................TYPICAL THIMBLE-EYE SPLICE 7-33 TABLE 7-6. ..............................................COPPER OVAL SLEEVE DATA 7-34 TABLE 7-7. ..............................................COPPER STOP SLEEVE DATA 7-34 FIGURE 7-15. ............................................. TYPICAL TERMINAL GAUGE 7-34

7-149.....................................................................................CABLE SYSTEM INSPECTION

. … 7-35

FIGURE 7-16. .................................... CABLE INSPECTION TECHNIQUE 7-35 FIGURE 7- 17...................................................CABLE WEAR PATTERNS 7-36 FIGURE 7-18................ WORN CABLE (REPLACEMENT NECESSARY) 7-36 FIGURE 7- 19....................................................INTERNAL CABLE WEAR 7-37 FIGURE 7- 20.................................................PULLEY WEAR PATTERNS 7-38

7-150..................................................................... CORROSION AND RUST PREVENTION

. … 7-38

FIGURE 7- 21 … CORROSION

7-39

7-151..............................................................................................................WIRE SPLICES

. … 7-39

FIGURE 7-22. ....................................MANUFACTURER’S WIRE SPLICE 7-39

7-152................................................................................................ CABLE MAINTENANCE

. … 7-39

7-153.................................................................................CABLE TENSION ADJUSTMENT

. … 7-39

7-154.7-164. [RESERVED.] … 7-39

**SECTION 9. TURNBUCKLES**

7-165 … GENERAL

. … 7-41

7-166.................................................................................... TURNBUCKLE INSTALLATION

. … 7-41

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Page xxxvii FIGURE 7-23. ..............................TURNBUCKLE THREAD TOLERANCE 7-41

7-167............................................................................................................ WITNESS HOLE

. … 7-41

7-168.7-178. [RESERVED.] … 7-41

**SECTION 10. SAFETY METHODS FOR TURNBUCKLES**

7-179..................................................................................................................... GENERAL.

. … 7-43

7-180..........................................................................................DOUBLE-WRAP METHOD..

. … 7-43

**FIGURE 7- 24.............................................. SAFETYING TURNBUCKLES 7-44**

7-181.............................................................................................SINGLE-WRAP METHOD.

. … 7-43

7-182................................................................ SAFETY-WIRE SECURED TURNBUCKLES

. … 7-44

TABLE 7-8........................................ TURNBUCKLE SAFETYING GUIDE 7-45 FIGURE 7- 25................................................SECURING TURNBUCKLES 7-46

7-183.................................................................................... SPECIAL LOCKING DEVICES..

. … 7-45

**FIGURE 7-26. ........................................... CLIP-TYPE LOCKING DEVICE 7-46**

**7-184.............................ASSEMBLING AND SECURING CLIP-LOCKING TURNBUCKLES**

. … 7-45

TABLE 7-9................................................ LOCKING-CLIP APPLICATION 7-47 FIGURE 7-27.ASSEMBLING AND SECURI NG CLIP-LOCKING TURNBUCKLES 7-47

7-185.7-195. [RESERVED.] … 7-47

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Page xxxviii SECTION 11. HARDWARE IDENTIFICATION TABLES

TABLE 7-10................................................................. T ABLE OF RIVETS 7-49 TABLE 7-11.............................................................. T ABLE OF SCREWS 7-50 TABLE 7-12.................................................................. T ABLE OF BOLTS 7-57 TABLE 7- 13.................................................................... TABLE OF NUTS 7-63 TABLE 7-14. ...........................................................TABLE OF WASHERS 7-69 TABLE 7- 15......................................................................TABLE OF PINS 7-71

7-196.7-206. [RESERVED.] … 7-74

Chapter 8. Engines, Fuel, Exhaust, And Propellers SECTION 1. ENGINES

8-1 … GENERAL

. … 8-1

8-2..................................................................................................... SPECIAL INSPECTION

. … 8-1

**8-3.......................................CRANKSHAFT INSPECTION AND REPAIR REQUIREMENTS**

. … 8-3

8-4.................................................. REPLACEMENT PARTS IN CERTIFICATED ENGINES

. … 8-3

8-5................................................................................... OIL SYSTEM LINES INSPECTION

. … 8-4

8-6................................................................................................. OIL FILTER INSPECTION

. … 8-4

8-7.....................................................CYLINDER HOLD-DOWN NUTS AND CAP SCREWS

. … 8-4

8-8...................................................................................REUSE OF SAFETYING DEVICES

. … 8-4

**8-9..................... SELF-LOCKING NUTS FOR AIRCRAFT ENGINES AND ACCESSORIES**

. … 8-4

8-10 … METALLIZING

. … 8-5

8-11 … PLATING

. … 8-5

**TABLE 8-1.CURRENT ENGINE A ND MAXIMUM PERMISSIBLE CYLINDER**

BARREL OVERSIZE … 8-6

8-12 … CORROSION

. … 8-6

8-13..............................................................................................................ENGINE RUN-IN

. … 8-6

**8-14................................COMPRESSION TESTING OF AIRCRAFT ENGINE CYLINDERS**

. … 8-6

**FIGURE 8-1.SCHEMATIC OF DI FFERENTIAL PRESSURE COMPRESSION TESTER 8-7**

8-15................................................................................................................ SPARK PLUGS

. … 8-8

**FIGURE 8-2......... CHART OF SPARK PLUG TEMPERATURE RANGES 8-9**

### Page 39

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Page xxxix FIGURE 8-3. ..........................................HOT AND COLD SPARK PLUGS 8-10 FIGURE 8- 4............................................................SPARK PLUG REACH 8-10 FIGURE 8-5. .....................METHOD OF CHECKING SPARK PLUG GAP 8-10

8-16...........................................................................................OPERATIONAL PROBLEMS

. … 8-11

FIGURE 8-6......................... TYPICAL CARBON-FOULED SPARK PLUG 8-12 FIGURE 8-6A. ........................... TYPICAL LEAD-FOULED SPARK PLUG 8-12 FIGURE 8-6B. ...............................TYPICAL OIL-FOULED SPARK PLUG 8-13 FIGURE 8-6C. ..TYPICAL SPARK PLUG WITH CRACKED CORE NOSE 8-13 FIGURE 8-6D . ................................ TYPICAL WORN OUT SPARK PLUG 8-14 FIGURE 8-6E. TYPICAL SPARK PL UG WITH BRIDGED ELECTRODES 8-14 FIGURE 8-7. .......................................SPARK PLUG WELL FLASHOVER 8-14

8-17..................................................SPARK PLUG PRE-RECONDITIONING INSPECTION

. … 8-14

8-18........................................................................... IGNITION HARNESSES INSPECTION

. … 8-14

**FIGURE 8-8..........................T YPICAL METHOD OF CLAMPING LEADS 8-15**

### Page 40

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Page xl

8-19................................................................................................ MAGNETO INSPECTION

. … 8-15

FIGURE 8-9. ..................................................NORMAL CONTACT POINT 8-16 FIGURE 8- 10............................. POINT WITH MINOR IRREGULARITIES 8-16 FIGURE 8- 11...............................POINT WITH WELL-DEFINED MOUND 8-16

8-20.................................................................................... MAGNETO-TO-ENGINE TIMING

. … 8-17

8-21.8-29. [RESERVED.] … 8-17

**SECTION 2. FUEL SYSTEMS**

8-30 … GENERAL

. … 8-19

8-31...........................................................................................FUEL LINES AND FITTINGS

. … 8-19

**FIGURE 8- 12........................ LOCATION OF CLAMPS AT TUBE BENDS 8-20**

8-32.............................................................................................. FUEL TANKS AND CELLS

. … 8-19

8-33..................................................FUEL TANK CAPS, VENTS, AND OVERFLOW LINES

. … 8-20

**8-34......... FUEL CROSS-FEED, FIREWALL SHUTOFF, AND TANK SELECTOR VALVES**

. … 8-21

8-35...................................................................................................................FUEL PUMPS

. … 8-21

8-36................................................................. FUEL FILTERS, STRAINERS, AND DRAINS

. … 8-21

8-37................................................................................................... INDICATOR SYSTEMS

. … 8-22

8-38..................................................................................... FUEL SYSTEM PRECAUTIONS

. … 8-22

8-39.8-44. [RESERVED.] … 8-22

**SECTION 3. EXHAUST SYSTEMS**

8-45 … GENERAL

. … 8-23

8-46................................................................... MUFFLER/HEAT EXCHANGER FAILURES

. … 8-23

**FIGURE 8-13.TYPICAL MUFFLER WA LL FATIGUE FAILURE AT EXHAUST OUTLET.**

(A. COMPLETE MUFFLER ASSEMBLY WI TH HEAT SHROUD REMOVED; B.

..................................................................... DETAIL VIEW OF FAILURE.)

**8-23 FIGURE 8-14.....TYPICAL MUFFLER WALL FAILURE. (A. COMPLETE**

MUFFLER ASSEMBLY WITH HEAT SHROUD REMOVED; B. DETAIL VIEW OF FAILURE; C. CROSS

SECTION OF ....................................................................................FAILED MUFFLER.)

**8--24 FIGURE 8-15.TYPICAL MUFFLER WALL FATIGUE FAILURE. (A. COMPLETE**

**MUFFLER ASSEMBLY**

### Page 41

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Page xli WITH HEAT SHROUD PARTIALLY RE MOVED; B. DETAILED VIEW OF FAILURE.) 8-25 FIGURE 8-16.TYPICAL FATIGUE FAILURE OF MUFFLER END PLATE AT STACK INLET 8-25

8-47....................................................................................... MANIFOLD/STACK FAILURES

. … 8-23

8-48...................................................................................INTERNAL MUFFLER FAILURES

. … 8-23

**FIGURE 8-17.SECTION OF A MU FFLER SHOWING TYPICAL INTERNAL**

BAFFLE FAILURE … 8-25

FIGURE 8-18........... LOOSE PIECES OF A FAILED INTERNAL BAFFLE 8-25 FIGURE 8-19.FAILED INTERNAL BAFFLE PARTIALLY OBSTRUCTING THE

MUFFLER OUTLET … 8-27

**FIGURE 8-20.FAILED INTERNAL BA FFLE COMPLETELY OBSTRUCTING THE**

MUFFLER OUTLET … 8-27

FIGURE 8-21A. ....................EXAMPLE OF EXHAUST OUTLET GUARD 8-27 FIGURE 8-21B.EXAMPLE OF EXHAU ST OUTLET GUARD INSTALLED 8-27

8-49 … INSPECTION

. … 8-26

**FIGURE 8-22........ EFFECT OF IM PROPERLY POSITIONED EXHAUST**

PIPE/MUFFLER CLAMP … 8-28

**FIGURE 8-23. .........................................PRIMARY INSPECTION AREAS 8-28**

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Page xlii

8-50 … REPAIRS

. … 8-26

8-51 … TURBO-SUPERCHARGER

. … 8-27

8-52................................................................................................AUGMENTOR SYSTEMS

. … 8-28

8-53.8-70. [RESERVED.] … 8-28

**SECTION 4. REPAIR OF METAL PROPELLERS**

8-71 … GENERAL

. … 8-29

8-72...............................................................................................................STEEL BLADES

. … 8-29

8-73...............................................................................ALUMINUM PROPELLER REPAIRS

. … 8-29

**FIGURE 8-24.METHOD OF REPAIRING SURFACE CRACKS, NICKS, ETC.,**

ON ALUMINUM-ALLOY PROPELLERS … 8-30

**FIGURE 8-25.CORRECT AND INCO RRECT METHOD OF REWORKING**

LEADING EDGE OF ........................................................... ALUMI NUM-ALLOY PROPELLERS 8-30 FIGURE 8-26.METHOD OF REPAIRI NG DAMAGED TIP OF ALUMINUM-

ALLOY PROPELLERS … 8-31

8-74................................................................................................................REPAIR LIMITS

. … 8-31

**FIGURE 8-27. EXAMPLE 1. DETERMINE THE REPAIR WIDTH LIMITS.**

**8-32 FIGURE 8-28.EXAMPLE 2. DETE RMINE THE REPAIR THICKNESS LIMITS 8-34**

8-75.....................................................................................STEEL HUBS AND HUB PARTS

. … 8-33

8-76.....................................................................PROPELLER HUB AND FLANGE REPAIR

. … 8-33

**FIGURE 8-29. REPAIR OF FIXED-P ITCH HUB AND PROPELLER WITH**

ELONGATED OR DAMAGED BOLT HOLES … 8-35

8-77......................................................................................................CONTROL SYSTEMS

. … 8-34

8-78........................................................................................................ DEICING SYSTEMS

. … 8-34

8-79.8-90. [Reserved.] … 8-35

**SECTION 5. INSPECTION OF PROPELLERS**

8-91 … GENERAL

. … 8-37

TABLE 8-2.SAMPLE MANUFACT URER’S PROPELLER INSPECTION CHECKLIST 8-38

**8-92............................................WOOD OR COMPOSITION PROPELLERS AND BLADES**

. … 8-37

8-93............................................................................ METAL PROPELLERS AND BLADES

. … 8-38

8-94........................................................................................................... PROPELLER HUB

. … 8-39

8-95......................................................................................... TACHOMETER INSPECTION

. … 8-40

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Page xliii

8-96.8-106. [RESERVED.] … 8-40

**SECTION 6. PROPELLER TRACKING AND VIBRATION**

8-107 … GENERAL

. … 8-41

8-108................................................................................ PROPELLER TRACKING CHECK

. … 8-41

**FIGURE 8-30. PROPELLER TRACKI NG (WOOD BLOCK OR COWLING**

FIXTURE SHOWN) … 8-42

8-109 … VIBRATION

. … 8-41

8-110.8-129. [RESERVED.] … 8-42

### Page 44

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Page xliv Chapter 9. Aircraft Systems And Components SECTION 1. INSPECTION AND MAINTENANCE OF LANDING GEAR

9-1 … GENERAL

. … 9-1

9-2................................................................................................... GENERAL INSPECTION

. … 9-1

9-3....................................................................................... CLEANING AND LUBRICATING

. … 9-1

9-4.............................................................................................. FIXED-GEAR INSPECTION

. … 9-1

**TABLE 9-1. ...........................................BUNGEE CORD COLOR CODES 9-2**

9-5........................................................INSPECTION OF RETRACTABLE LANDING GEAR

. … 9-3

9-6.................................................................................................. EMERGENCY SYSTEMS

. … 9-3

9-7..................................................................................... LANDING GEAR COMPONENTS

. … 9-4

9-8.....................................................................................TIRE AND TUBE MAINTENANCE

. … 9-5

9-9............................................................................................................FLOATS AND SKIS

. … 9-6

**FIGURE 9-1. ............................................A TYPICAL SKI INSTALLATION 9-6**

9-10..................................................... INSPECTION AND REPAIR OF FLOATS AND SKIS

. … 9-7

9-11...................................................................... TYPES OF LANDING GEAR PROBLEMS

. … 9-8

FIGURE 9- 2....................................................... TYPICAL BOLT CRACKS 9-8 FIGURE 9- 3...............................TYPICAL CRACKS NEAR BOLT HOLES 9-8 FIGURE 9-4. ........................................... TYPICAL BOLT HOLE CRACKS 9-9 FIGURE 9- 5................................................TYPICAL ROD-END CRACKS 9-9 FIGURE 9-6........ TYPICAL TORQUE TUBE BOLT HOLE ELONGATION 9-9

9-12.................................................................................................SPECIAL INSPECTIONS

. … 9-9

9-13..................................................................................................... RETRACTION TESTS

. … 9-9

9-14................................................................................... TIRE INSPECTION AND REPAIR

. … 9-9

**FIGURE 9-7.EXAMPLES OF TREAD W EAR INDICATING OVER-INFLATION**

AND UNDER-INFLATION … 9-11

9-15..................................................................................................... INFLATION OF TIRES

. … 9-10

9-16......................................................................................................WHEEL INSPECTION

. … 9-11

9-17..................................................................................................WHEEL INSTALLATION

. … 9-11

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Page xlv

9-18.......................................................................................REASSEMBLING THE WHEEL

. … 9-11

9-19 … SLIPPAGE

. … 9-12

9-20.9-24. [RESERVED.] … 9-12

**SECTION 2. HYDRAULIC SYSTEMS**

9-25 … GENERAL

. … 9-13

9-26....................................................................... PURPOSES OF HYDRAULIC SYSTEMS

. … 9-13

9-27...................................................................................... TYPES OF HYDRAULIC FLUID

. … 9-13

9-28...................................................................................HANDLING HYDRAULIC FLUIDS

. … 9-14

9-29....................................................HYDRAULIC SYSTEM MAINTENANCE PRACTICES

. … 9-15

9-30...............................................................................HYDRAULIC LINES AND FITTINGS

. … 9-18

TABLE 9-2. ............................................................................. TUBE DATA 9-19 FIGURE 9-8.HOSE ASSEMBLY INSTRUCT IONS (CAN BE USED FOR LOW

PRESSURE HYDRAULIC FLUID, .................................................................AND OI L LINE APPLICATIONS) 9-21

### Page 46

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Page xlvi TABLE 9-3. ..................................... AIRCRAFT HOSE SPECIFICATIONS 9-22 FIGURE 9-9. ........................................PROPER HOSE INSTALLATIONS 9-23 FIGURE 9- 10.........................................................MINIMUM BEND RADII 9-24 TABLE 9-4.BALL DIAMETERS FOR TESTING HOSE RESTRICTIONS OR KINKING 9-25 FIGURE 9-11...........SUGGESTED HANDLING OF PREFORMED HOSE 9-25

9-31.9-36. [RESERVED.] … 9-25

**SECTION 3. EMERGENCY EQUIPMENT**

9-37.....................................................................................................................LIFE RAFTS

. … 9-27

9-38.............................................................................................. LIFE RAFT INSPECTIONS

. … 9-27

FIGURE 9-12............................................................... INFLATION VALVE 9-29

9-39..........................................................................................SURVIVAL KIT INSPECTION

. … 9-28

9-40.................................................................................................SPECIAL INSPECTIONS

. … 9-31

9-41...................................................................................................INSPECTION RECORD

. … 9-32

9-42...............................................................................................................RAFT REPAIRS

. … 9-32

FIGURE 9- 13......................................................... REPAIR DIMENSIONS 9-33

9-43........................................................................................................ LIFE PRESERVERS

. … 9-34

9-44....................................................................................LIFE PRESERVER INSPECTION

. … 9-35

9-45....................................................................................REPAIR OF LIFE PRESERVERS

. … 9-36

9-46.................................................................................... MISCELLANEOUS EQUIPMENT

. … 9-36

9-47........................................................................................................OXYGEN SYSTEMS

. … 9-37

9-48 … INSPECTION

. … 9-37

9-49 … MAINTENANCE

. … 9-38

FIGURE 9-14. .......................................... OXYGEN CYLINDER DAMAGE 9-39 FIGURE 9-15. .............................. CYLINDER BRACKETS AND CLAMPS 9-39

9-50......................................................................FUNCTIONAL TESTING AFTER REPAIR

. … 9-41

9-51...................................................SERVICE REQUIREMENTS --OXYGEN CYLINDERS

. … 9-41

**TABLE 9-5........................................... TABLE OF FILLING PRESSURES 9-43**

9-52.9-59. [RESERVED.] … 9-43

### Page 47

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Page xlvii SECTION 4. CABIN INTERIOR

9-60 … GENERAL

. … 9-45

9-61.........................................................................................CAR-3 AIRCRAFT INTERIOR

. … 9-45

9-62......................................................................................PART 23 AIRCRAFT INTERIOR

. … 9-45

9-63..........................................................................................SOURCE OF INFORMATION

. … 9-46

9-64......................................................................................UPHOLSTERY AND/OR BELTS

. … 9-46

9-65.9-70. [RESERVED.] … 9-46

Chapter 10. Weight And Balance SECTION 1. TERMINOLOGY

10-1 … GENERAL

. … 10-1

10-2 … TERMINOLOGY

. … 10-1

FIGURE 10-1. ...........................................TYPICAL DATUM LOCATIONS 10-2 FIGURE 10-2. ..................ILLUSTRATION OF ARM (OR MOMENT ARM) 10-3 FIGURE 10- 3...........................EXAMPLE OF MOMENT COMPUTATION 10-4 FIGURE 10-4.......EMPTY WEIGHT CENTER OF GRAVITY FORMULAS 10-5 FIGURE 10-5.EMPTY WEIGHT AND EMPTY CENTER OF GRAVITY - TAIL-

WHEEL TYPE AIRCRAFT … 10-6

### Page 48

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Page xlviii FIGURE 10-6.EMPTY WEIGHT AND EMPTY WEIGHT CENTER OF GRAVITY -

NOSEWHEEL TYPE AIRCRAFT … 10-7

FIGURE 10-7. ....................OPERATING CENTER OF GRAVITY RANGE 10-8 FIGURE 10-8. .......................................WEIGHING POINT CENTERLINE 10-9

10-3.10-13. [RESERVED.] … 10-9

**SECTION 2. WEIGHING PROCEDURES**

10-14 … GENERAL

. … 10-11

10-15 … PROCEDURES

. … 10-11

FIGURE 10-9.EMPTY WEIGHT AND EMPTY WEIGHT CENTER OF GRAVITY ...............................................WHEN AIRCRAFT IS WEIGHED WITH OIL 10-12

10-16................................................................ WEIGHT AND BALANCE COMPUTATIONS

. … 10-11

10-17.................................................... WEIGHT AND BALANCE EXTREME CONDITIONS

. … 10-13

**FIGURE 10-10.EXAMPLE OF CHECK OF MOST FORWARD WEIGHT AND**

BALANCE EXTREME … 10-14

**FIGURE 10-11.EXAMPLE OF CHECK OF MOST REARWARD WEIGHT AND**

BALANCE EXTREME … 10-16

10-18............................................................LOADING CONDITIONS AND/OR PLACARDS

. … 10-15

FIGURE 10-12.LOADING CONDITIONS: DETERMINATION OF THE NUMBER

OF PASSENGERS .................................. AND BA GGAGE PERMISSIBLE WITH FULL FUEL 10-17 FIGURE 10-13.LOADING CONDITIONS: DETERMINATION OF THE FUEL

AND BAGGAGE PERMISSIBLE ...............................................................WITH M AXIMUM PASSENGERS 10-18 FIGURE 10-14.LOADING CONDITIONS: DETERMINATION OF THE FUEL

**AND THE NUMBER AND**

**LOCATION OF PASSENGERS PERMISSIBLE WITH MAXIMUM BAGGAGE 10-19**

10-19..........................................................................................................EQUIPMENT LIST

. … 10-15

10-20..................................................................................................EQUIPMENT CHANGE

. … 10-15

FIGURE 10-15.EFFECTS OF THE ADDITION OF EQUIPMENT ITEMS ON BALANCE 10-20 FIGURE 10-16.EXAMPLE OF MOMENT AND WEIGHT CHANGES RESULTING

FROM EQUIPMENT CHANGES … 10-21

10-21...............................................................SAMPLE WEIGHT AND BALANCE REPORT

. … 10-20

**FIGURE 10-17.SAMPLE WEIGHT AND BALANCE REPORT TO DETERMINE**

EMPTY WEIGHT CENTER OF GRAVITY … 10-22

**FIGURE 10-18.SAMPLE WEIGHT AND BALANCE REPORT INCLUDING AN**

EQUIPMENT CHANGE ..............................................................FOR AI RCRAFT FULLY LOADED 10-23

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Page xlix

10-22.......................................................................................INSTALLATION OF BALLAST

. … 10-20

**FIGURE 10-19...... PERMANENT BALLAST COMPUTATION FORMULA 10-24**

10-23...................................................................................................LOADING SCHEDULE

. … 10-24

10-24.10-34. [RESERVED.] … 10-24

Chapter 11. Aircraft Electrical Systems SECTION 1. INSPECTION AND CARE OF ELECTRICAL SYSTEMS

11-1 … GENERAL

. … 11-1

11-2.................................................................... INSPECTION AND OPERATION CHECKS

. … 11-1

**11-3.. FUNCTIONAL CHECK OF CHECK OF STAND-BY OR EMERGENCY EQUIPMENT**

. … 11-2

11-4.................................................................................CLEANING AND PRESERVATION

. … 11-2

11-5...................................................................... BATTERY ELECTROLYTE CORROSION

. … 11-2

11-6.......................................................................................... ADJUSTMENT AND REPAIR

. … 11-2

11-7................................................................INSULATION OF ELECTRICAL EQUIPMENT

. … 11-2

11-8.......................................................................................................................BUS BARS

. … 11-3

11-9.11-14. [RESERVED.] … 11-3

### Page 50

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Page l SECTION 2. STORAGE BATTERIES

11-15 … GENERAL

. … 11-5

11-16.................................................................................................. BATTERY CHARGING

. … 11-5

11-17....................................................................................................BATTERY FREEZING

. … 11-5

**TABLE 11-1.LEAD-ACID BATTERY ELECTROLYTE FREEZING POINTS 11-5**

11-18...................................................................................TEMPERATURE CORRECTION

. … 11-5

**TABLE 11-2............... SULF URIC ACID TEMPERATURE CORRECTION 11-6**

11-19............................................................................................BATTERY MAINTENANCE

. … 11-6

11-20...........................................................................................ELECTROLYTE SPILLAGE

. … 11-8

11-21......................................................................................................... NOXIOUS FUMES

. … 11-9

11-22........................................................................................ INSTALLATION PRACTICES

. … 11-9

**FIGURE 11-1. ..................................BATTERY VENTILATING SYSTEMS 11-9**

11-23.11-29. [RESERVED.] … 11-10

**SECTION 3. INSPECTION OF EQUIPMENT INSTALLATION**

11-30 … GENERAL

. … 11-11

11-31............................................................... INSTALLATION CLEARANCE PROVISIONS

. … 11-11

**11-32............................WIRES, WIRE BUNDLES, AND CIRCUIT PROTECTIVE DEVICES**

. … 11-11

11-33.......................................................................................................... OUTPUT RATING

. … 11-11

11-34................................................................................................ ALTERNATOR DIODES

. … 11-11

11-35...........................................................STATIC ELECTRICAL POWER CONVERTERS

. … 11-11

**11-36.ACCEPTABLE MEANS OF CONTROLLING OR MONITORING THE ELECTRICAL LOAD**

. … 11-12

11-37.................................................................DETERMINATION OF ELECTRICAL LOAD

. … 11-12

11-38..............................................................................JUNCTION BOX CONSTRUCTION

. … 11-13

11-39.11-46. [RESERVED.] … 11-13

**SECTION 4. INSPECTION OF CIRCUIT-PROTECTION DEVICES**

11-47 … GENERAL

. … 11-15

11-48................................................ DETERMINATION OF CIRCUIT BREAKER RATINGS

. … 11-15

11-49.............................................................................DC CIRCUIT PROTECTOR CHART

. … 11-15

### Page 51

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Page li TABLE 11-3. .................. DC WIRE AND CIRCUIT PROTECTOR CHART 11-15

11-50.......................................................RESETTABLE CIRCUIT PROTECTION DEVICES

. … 11-15

11-51.........................................................................................CIRCUIT BREAKER USAGE

. … 11-15

11-52........................................................................... CIRCUIT BREAKER MAINTENANCE

. … 11-16

11-53 … SWITCHES

. … 11-16

TABLE 11-4............................................ SWITCH DERATING FACTORS 11-17 TABLE 11-5................................ SEL ECTION OF CONTACT MATERIAL 11-17

11-54 … RELAYS

. … 11-19

11-55............................................................................................LOAD CONSIDERATIONS

. … 11-20

**11-56......................................OPERATING CONDITIONS FOR SWITCHES AND RELAYS**

. … 11-20

11-57.11-65. [RESERVED.] … 11-20

**SECTION 5. ELECTRICAL WIRE RATING**

11-66 … GENERAL

. … 11-21

TABLE 11-6..... TABULATION CHART (ALLOWABLE VOLTAGE DROP) 11-21 TABLE 11-7.EXAMPLES OF DETERMININ G REQUIRED WIRE SIZE USING FIGURE 11-2 11-22 TABLE 11-8.EXAMPLES OF DETERMINING MAXIMUM RUN LENGTH USING FIGURE 11-3 11-22

### Page 52

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Page lii FIGURE 11-2.................... CONDUCTOR CHART, CONTINUOUS FLOW 11-23 FIGURE 11-3................. CONDUCTOR CHART, INTERMITTENT FLOW 11-24 TABLE 11-9. CURRENT CARRYING CAPACITY AND RESISTANCE OF COPPER WIRE 11-25 TABLE 11-10.CURRENT CARRYING CAPACITY AND RESISTANCE OF

ALUMINUM WIRE … 11-25

**11-67..... METHODS FOR DETERMINING CURRENT CARRYING CAPACITY OF WIRES**

. … 11-26

FIGURE 11-4A. ............................ SINGLE COPPER WIRE IN FREE AIR 11-29 FIGURE 11-4B. ............................ SINGLE COPPER WIRE IN FREE AIR 11-30 FIGURE 11- 5............................................ BUNDLE DERATING CURVES 11-31 FIGURE 11-6. ...........................................ALTITUDE DERATING CURVE 11-32

**11-68......................................INSTRUCTIONS FOR USE OF ELECTRICAL WIRE CHART**

. … 11-27

11-69.......................................................COMPUTING CURRENT CARRYING CAPACITY

. … 11-33

11-70.11-75. [RESERVED.] … 11-34

**SECTION 6. AIRCRAFT ELECTRICAL WIRE SELECTION**

11-76 … GENERAL

. … 11-35

11-77......................................................................................AIRCRAFT WIRE MATERIALS

. … 11-35

11-78 … SUBSTITUTIONS

. … 11-36

11-79.11-84. [RESERVED.] … 11-37

**SECTION 7. TABLE OF ACCEPTABLE WIRES**

11-85...............................................................................................AIRCRAFT WIRE TABLE

. … 11-39

TABLE 11- 11......................................................................OPEN WIRING 11-40 TABLE 11- 12..........................................................PROTECTED WIRING 11-41

11-86. OPEN AIRFRAME IN TERCONNECTING WIRE … 11-39

11-87. PROTECTE D WIRE … 11-39

11-88. SEVERE WIND AND MOISTURE  PROBLEMS (SWAMP) … 11-39

11-89. SHIELDED  WIRE … 11-39

11-90.11-95. [RESERVED.] … 11-41

**SECTION 8. WIRING INSTALLATION INSPECTION REQUIREMENTS**

11-96 … GENERAL

. … 11-43

11-97..............................................................................................WIRING REPLACEMENT

. … 11-45

11-98....................................................................... TERMINALS AND TERMINAL BLOCKS

. … 11-46

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Page liii

11-99......................................................................................FUSES AND FUSE HOLDERS

. … 11-46

11-100 … CONNECTORS

. … 11-47

**11-101...........JUNCTION BOXES, PANELS, SHIELDS, AND MICROSWITCH HOUSINGS**

. … 11-48

**11-102...CONDUIT-RIGID METALLIC, FLEXIBLE METALLIC AND RIGID NONMETALLIC**

. … 11-48

11-103 … JUNCTIONS

. … 11-48

11-104.................................................................................................. CIRCUIT BREAKERS

. … 11-49

11-105.............................................................................................. SYSTEM SEPARATION

. … 11-49

11-106.......................................................... ELECTROMAGNETIC INTERFERENCE (EMI)

. … 11-49

11-107.............................................................................................INTERFERENCE TESTS

. … 11-49

**11-108.....IDENTIFICATION STENCILS AND PLACARDS ON ELECTRICAL EQUIPMENT**

. … 11-50

11-109.11-114. [RESERVED.] … 11-50

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Page liv SECTION 9. ENVIRONMENTAL PROTECTION AND INSPECTION

11-115..........................................................................MAINTENANCE AND OPERATIONS

. … 11-51

11-116.......................................................................................GROUP AND BUNDLE TIES

. … 11-51

FIGURE 11- 7................................................GROUP AND BUNDLE TIES 11-51 FIGURE 11-8.........COMB FOR STRAIGHTENING WIRES IN BUNDLES 11-51

11-117..................................................................................... MINIMUM WIRE BEND RADII

. … 11-51

11-118 … SLACK

. … 11-52

**FIGURE 11-9. ..........................................SLACK BETWEEN SUPPORTS 11-52**

11-119......................................................................................................POWER FEEDERS

. … 11-52

11-120....................................................................................................................RF CABLE

. … 11-52

11-121 … PRECAUTIONS

. … 11-52

**11-122.......... MOISTURE PROTECTION, WHEEL WELLS, AND LANDING GEAR AREAS**

. … 11-53

11-123...............................................PROTECTION AGAINST PERSONNEL AND CARGO

. … 11-53

11-124.................................................................................................HEAT PRECAUTIONS

. … 11-53

11-125..................................................... MOVABLE CONTROLS WIRING PRECAUTIONS

. … 11-53

11-126............................................................................FLAMMABLE FLUIDS AND GASES

. … 11-53

**FIGURE 11-10....... SEPARATION OF WIRES FROM PLUMBING LINES 11-54**

11-127.11-134. [RESERVED.] … 11-55

**SECTION 10. SERVICE LOOP HARNESSES (PLASTIC TIE STRIPS)**

11-135 … GENERAL

. … 11-55

11-136 … SUPPORT

. … 11-55

11-137.........................................................................................ANTI-CHAFING MATERIAL

. … 11-55

11-138...........................................................................................................STRAIN RELIEF

. … 11-55

11-139. .........................................................................................................SERVICE LOOP

. … 11-55

11-140.11-145. [RESERVED.] … 11-56

**SECTION 11. CLAMPING**

11-146 … GENERAL

. … 11-57

**FIGURE 11-11. ............................... SAFE ANGLE FOR CABLE CLAMPS 11-58**

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Page lv FIGURE 11-12.TYPICAL MOUNTING HARDWARE FOR MS-21919 CABLE CLAMPS 11-58 FIGURE 11-13................IN STALLING CABLE CLAMP TO STRUCTURE 11-59 FIGURE 11- 14................................. CLAMPING AT A BULKHEAD HOLE 11-60

11-147................................................................WIRE AND CABLE CLAMPS INSPECTION

. … 11-57

11-148.11-154. [RESERVED.] … 11-60

**SECTION 12. WIRE INSULATION AND LACING STRING TIE**

11-155 … GENERAL

. … 11-61

11-156.......................................................................................... INSULATION MATERIALS

. … 11-61

11-157........................................................................................... STRIPPING INSULATION

. … 11-61

**TABLE 11-13.................ALLO WABLE NICKED OR BROKEN STRANDS 11-62**

11-158...................................................................................................... LACING AND TIES

. … 11-61

FIGURE 11- 15..................................................... SINGLE CORD LACING 11-63 FIGURE 11- 16....................................................DOUBLE CORD LACING 11-63 FIGURE 11-17. .................................................................... MAKING TIES 11-64

11-159..................................................................................................... INSULATION TAPE

. … 11-62

11-160.11-166. [RESERVED.] … 11-64

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Page lvi SECTION 13. SPLICING.

11-167 … GENERAL

. … 11-65

**FIGURE 11- 18.......................STAGGERED SPLICES IN WIRE BUNDLE 11-65**

11-168.11-173. [RESERVED.] … 11-65

**SECTION 14. TERMINAL REPAIRS**

11-174 … GENERAL

. … 11-67

11-175..............................................................ATTACHMENT OF TERMINALS TO STUDS

. … 11-68

11-176.........................................................................................STUDS AND INSULATORS

. … 11-68

11-177................................................................ WIRE TERMINALS AND BINDING POSTS

. … 11-69

11-178............................................................ CRIMP ON TERMINAL LUGS AND SPLICES

. … 11-69

11-179............................................. LOCK WASHERS FOR TERMINALS ON EQUIPMENT

. … 11-70

11-180.11-184. [RESERVED.] … 11-70

**SECTION 15. GROUNDING AND BONDING**

11-185 … GENERAL

. … 11-71

11-186 … GROUNDING

. … 11-71

11-187 … BONDING

. … 11-73

11-188..............................................................................................BONDING INSPECTION

. … 11-73

FIGURE 11- 19.................................................... MILLIVOLT DROP TEST 11-74

11-189........................................................................BONDING JUMPER INSTALLATIONS

. … 11-75

TABLE 11-14...STUD BONDING OR GROUNDING TO FLAT SURFACE 11-76 TABLE 11-15.PLATE NUT BONDING OR GROUNDING TO FLAT SURFACE 11-77 TABLE 11-16.BOLT AND NUT BONDING OR GROUNDING TO FLAT SURFACE 11-78 FIGURE 11-20.COPPER JUMPER CONNECTOR TO TUBULAR STRUCTURE 11-79 FIGURE 11-21. ............................BONDING CONDUIT TO STRUCTURE 11-79 FIGURE 11-22.ALUMINUM JUMPER CONNECTION TO TUBULAR STRUCTURE 11-79

11-190.............................................................................................. CREEPAGE DISTANCE

. … 11-75

11-191.......................................................................................................... FUEL SYSTEMS

. … 11-80

11-192........................................................... ELECTRIC SHOCK PREVENTION BONDING

. … 11-80

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Page lvii

11-193...................................................................... LIGHTNING PROTECTION BONDING

. … 11-80

**11-194................. LIGHTNING PROTECTION FOR ANTENNAS AND AIR DATA PROBES**

. … 11-81

11-195................................................................................. STATIC-DISCHARGE DEVICES

. … 11-81

11-196 … CLEANING

. … 11-81

11-197............................................................................................ HARDWARE ASSEMBLY

. … 11-81

11-198.11-204. [RESERVED.] … 11-81

**SECTION 16. WIRE MARKING**

11-205 … GENERAL

. … 11-83

11-206.............................................................................................. WIRE IDENTIFICATION

. … 11-83

**11-207.IDENTIFICATION AND INFORMATION RELATED TO THE WIRE AND WIRING**

DIAGRAMS … 11-83

11-208....................................................... PLACEMENT OF IDENTIFICATION MARKINGS

. … 11-83

FIGURE 11-23.SPACING OF PRINTED IDENTIFICATION MARKS (DIRECT MARKING) 11-84 FIGURE 11-24.SPACING OF PRINTED IDENTIFICATION MARKS (INDIRECT MARKING) 11-84

11-209.....................................................................................TYPES OF WIRE MARKINGS

. … 11-83

11-210...............................................................................................HOT STAMP MARKING

. … 11-84

### Page 58

Contents (continued) Paragraph Page

Page lviii

11-211............................................................................................. DOT MATRIX MARKING

. … 11-85

11-212...................................................................................................... INK JET MARKING

. … 11-85

11-213........................................................................................................LASER MARKING

. … 11-85

11-214....................................................................................... IDENTIFICATION SLEEVES

. … 11-85

11-215.............................................................................................. IDENTIFICATION TAPE

. … 11-86

11-216.........................................................................................OPERATING CONDITIONS

. … 11-86

11-217................................................................... INSTALLATION OF PRINTED SLEEVES

. … 11-86

**TABLE 11-17. RECOMMENDED SIZE OF IDENTIFICATION SLEEVING 11-86**

**11-218...................................... IDENTIFICATION OF WIRE BUNDLES AND HARNESSES**

. … 11-86

**FIGURE 11-25.IDENTIFICATION OF WIRE BUNDLES AND HARNESSES 11-86**

11-219...............................................................TERMINAL MARKING SLEEVE AND TAGS

. … 11-87

FIGURE 11- 26.........................................STANDARD SLEEVES (135 °C) 11-87 FIGURE 11-27.INSTALLATION OF HEAT-SHRINKABLE INSULATION SLEEVES 11-87

11-220....................................................... SLEEVES AND CABLE MARKERS SELECTION

. … 11-87

FIGURE 11-28..............................................................CABLE MARKERS 11-88 FIGURE 11- 29.................................. TIE-DOWN STRAP INSTALLATION 11-88 TABLE 11-18...............SELECTION TABLE FOR STANDARD SLEEVES 11-88 TABLE 11-19...............SELECTION TABLE FOR THIN-WALL SLEEVES 11-88 TABLE 11-20.SELECTION TABL E FOR HIGH-TEMPERATURE SLEEVES 11-89 TABLE 11-21..................... SEL ECTION TABLE FOR CABLE MARKERS 11-89 TABLE 11-22.PLASTIC TIE-DOWN STRAPS (MS3367, TYPE I, CLASS 1) 11-89 FIGURE 11-30. .......................TIE-DOWN STRAP INSTALLATION TOOL 11-89 FIGURE 11-31. .......................................... COMPLETED INSTALLATION 11-90

**11-221.....................................TEMPORARY WIRE AND CABLE MARKING PROCEDURE**

. … 11-90

**FIGURE 11-32.............TE MPORARY WIRE IDENTIFICATION MARKER 11-90**

11-222. MARKER SLEEVE INSTALLA TION AFTER PRINTING … 11-90

**FIGURE 11- 33.................................... INSERTING WIRE INTO MARKER 11-90**

### Page 59

Contents (continued) Paragraph Page

Page lix FIGURE 11-34. ...................................... SHRINKING MARKER ON WIRE 11-90

11-223.11-229. [RESERVED.] … 11-90

**SECTION 17. CONNECTORS**

11-230 … GENERAL

. … 11-91

11-231 … SELECTION

. … 11-91

11-232.........................................................................................TYPES OF CONNECTORS

. … 11-91

FIGURE 11- 35.........................CONNECTOR INFORMATION EXAMPLE 11-92 FIGURE 11- 36...........................DIFFERENT TYPES OF CONNECTORS 11-93 FIGURE 11- 37...........................................COAX CABLE CONNECTORS 11-94

11-233.......................................................................... VOLTAGE AND CURRENT RATING

. … 11-96

11-234....................................................................SPARE CONTACTS (FUTURE WIRING)

. … 11-97

11-235 … INSTALLATION

. … 11-97

11-236.................................................FEED-THROUGH BULKHEAD WIRE PROTECTION

. … 11-98

11-237........................................................................... SPECIAL PURPOSE CONNECTOR

. … 11-98

11-238.............................................................................................POTTING COMPOUNDS

. … 11-98

11-239...........................................................................................POTTING CONNECTORS

. … 11-98

.......................................... FIGURE 11-38. SPARE WIRES FOR POTTING CONNECTOR

. … 11-99

11-240...................................................................................................... THROUGH BOLTS

. … 11-99

11-241.11-247. [RESERVED.] … 11-99

### Page 60

Contents (continued) Paragraph Page

Page lx SECTION 18. CONDUITS

11-248 … GENERAL

. … 11-101

11-249...................................................................................................... SIZE OF CONDUIT

. … 11-101

11-250....................................................................................................CONDUIT FITTINGS

. … 11-101

11-251.......................................................................................... CONDUIT INSTALLATION

. … 11-101

11-252.......................................................................................................... RIGID CONDUIT

. … 11-101

**TABLE 11-23..................................... BEND RADII FOR RIGID CONDUIT 11-101**

11-253...................................................................................................FLEXIBLE CONDUIT

. … 11-102

**TABLE 11-24.MINIMUM BENDING RADII FOR FLEXIBLE ALUMINUM OR**

BRASS CONDUIT … 11-102

11-254.11-259. [RESERVED.] … 11-102

**SECTION 19. PROTECTION OF UNUSED CONNECTORS**

11-260 … GENERAL

. … 11-103

11-261..................................................................................... QUICK REFERENCE CHART

. … 11-103

**TABLE 11-25.......... CONTACT CAVI TY SEALING-QUICK REFERENCE 11-104**

11-262.................................................................UNPRESSURIZED AREA CONNECTORS

. … 11-103

FIGURE 11-39. .............................................STUB WIRE INSTALLATION 11-104 TABLE 11-26. .............................................SEALING ROD DIMENSIONS 11-104

11-263...............................................................................................PRESSURIZED AREAS

. … 11-103

**FIGURE 11-40.SEALING UNUSED CO NTACT CAVITIES-UNPRESSURIZED**

AREAS-(CUT-AWAY VIEW) … 11-104

11-264.11-270. [RESERVED.] … 11-104

**SECTION 20. ELECTRICAL AND ELECTRONIC SYMBOLS**

11-271 … GENERAL

. … 11-105

11-272 … SYMBOLS

. … 11-105

Table 11-27. Electronic/ Electrical Symbols................................................................................... 1 1-105

11-273.11-283. [RESERVED.] … 11-118

Chapter 12. Aircraft Avionics Systems SECTION 1. AVIONICS EQUIPMENT MAINTENANCE

12-1 … GENERAL

. … 12-1

12-2. Handling of Components … 12-1

12-3.12-7. [RESERVED.] … 12-1

**SECTION 2. GROUND OPERATIONAL CHECKS FOR AVIONICS EQUIPMENT (ELECTRICAL)**

### Page 61

Contents (continued) Paragraph Page

Page lxi

12-8 … GENERAL

. … 12-3

12-9........................................................................... INSPECTION OF AVIONICS SYSTEM

. … 12-3

12-10......................................................................................COMMUNICATION SYSTEMS

. … 12-5

### Page 62

Par 1-1 Page 1-62

12-11...................................................................VHF OMNI-DIRECTIONAL RANGE (VOR)

. … 12-5

12-12............................................................DISTANCE MEASURING EQUIPMENT (DME)

. … 12-5

12-13....................................................................AUTOMATIC DIRECTION FINDER (ADF)

. … 12-5

12-14................................................................... INSTRUMENT LANDING SYSTEMS (ILS)

. … 12-6

12-15........................................................................................................MARKER BEACON

. … 12-6

12-16........................................................................LONG RANGE NAVIGATION (LORAN)

. … 12-6

12-17...................................................................... GLOBAL POSITIONING SYSTEM (Gps)

. … 12-7

12-18.................................................................................................AUTOPILOT SYSTEMS

. … 12-7

12-19 … ALTIMETERS

. … 12-7

12-20 … TRANSPONDERS

. … 12-7

12-21.........................................................EMERGENCY LOCATOR TRANSMITTERS (Elt)

. … 12-8

12-22...................................................................................................... INSPECTION OF Elt

. … 12-8

12-23..........................................................................................FLIGHT DATA RECORDER

. … 12-9

12-24..........................................................................COCKPIT VOICE RECORDERS (Cvr)

. … 12-10

12-25........................................................................................................WEATHER RADAR

. … 12-10

12-26...............................................................................................RADOME INSPECTION.

. … 12-10

12-27..................................................................................................................... DATA BUS

. … 12-11

12-28.12-36. [RESERVED.] … 12-11

**SECTION 3. GROUND OPERATIONAL CHECKS FOR AVIONICS EQUIPMENT (NON ELECTRICAL)**

12-37..........................................................................................COMPASS CHECK SWING

. … 12-13

12-38.................................................................................................... PNEUMATIC GYROS

. … 12-14

FIGURE 12-1.VENTURI SYSTEM FOR PROVIDING AIRFLOW THROUGH GYRO INSTRUMENTS 12-15 FIGURE 12-2.INSTRUMENT VACUUM SYSTEM USING A WET-TYPE VACUUM PUMP 12-16 FIGURE 12-3.INSTRUMENT VACUUM SYSTEM USING A DRY-TYPE AIR PUMP 12-17 FIGURE 12-4.INSTRUMENT PRESSURE SYSTEM USING A DRY-TYPE AIR PUMP 12-17

12-39.12-50. [RESERVED.] … 12-17

Section 4. Avionics Test Equipment

12-51 … GENERAL

. … 12-19

12-52.......................................................TEST EQUIPMENT CALIBRATION STANDARDS

. … 12-19

### Page 63

Par 1-1 Page 1-63

12-53.............................................................................. TEST EQUIPMENT CALIBRATION

. … 12-19

12-54.12-64. [RESERVED.] … 12-19

Appendix 1. Glossary. (10 pages)

Appendix 2................................................... Acronyms and ABBREVIATIONS (4 pages)

**APPENDIX 3.......................METRIC-BASED PREFIXES AND POWERS OF 10 (1 PAGE)**

### Page 64

Par 1-1 Page 1-64 CHAPTER 1. WOOD STRUCTURE

**SECTION 1. MATERIALS AND PRACTICES**

GENERAL. Wood aircraft construction dates back to the early days of certificated aircraft.

Today only a limited number of wood aircraft structures are produced. However, many of the older airframes remain in service. With proper care, airframes from the 1930’s through the 1950’s have held up remarkably well considering the state of technology and long term experience available at that time. It is the responsibility of the mechanic to carefully inspect such structures for deterioration and continuing airworthiness.

**WOODS.**

Quality of Wood. All wood and plywood used in the repair of aircraft structures should be of aircraft quality (reference Army Navy Commerce Department Bulletin ANC-19, Wood Aircraft Inspection and Fabrication).

Table 1-1 lists some permissible variations in characteristics and properties of aircraft wood.

However, selection and approval of woodstock for aircraft structural use are specialized skills and should be done by personnel who are thoroughly familiar with inspection criteria and methods.

Substitution of Original Wood. The wood species used to repair a part should be the same as that of the original whenever possible; however, some permissible substitutes are given in table 1-1. Obtain approval from the airframe manufacturer or the Federal Aviation Administration (FAA) for th e replacement of modified woods or other nonwood products with a substitute material.

Effects of Shrinkage. When the moisture content of a wooden part is lowered, the part shrinks. Since the shrinkage is not equal in all directions, the mechanic should consider the effect that the repair may have on the completed structure. Th e shrinkage is greatest in a tangential direction (across the fibers and parallel to the growth rings), somewhat less in a radial direction (across the fibers and perpendicular to the growth rings), and is negligible in a longitudinal direction (parallel to the fibers). Figure 1-1 illustrates the different grain directions and the effects of shrinkage on the shape of a part. These dimensional changes can have several detrimental effects upon a wood structure, such as loosening of

FIGURE 1-1. Relative shrinkage of wood members due to drying.

### Page 65

Par 1-2 Page 1-65 TABLE 1-1. Selection and Properties of Aircraft Wood. (See notes following table.)

Species of Wood

Strength properties as compared to spruce Maximum permissible grain deviation (slope of grain)

Remarks

1. 2. 3. 4.

Spruce(Picea) Sitka (P.

Sitchensis) Red (P. Rubra) White (P. Glauca).

100%

1:15 Excellent for all uses. Considered as standard for this table.

Douglas Fir (Pseudotsuga Taxifolia).

Exceeds spruce.

1:15 May be used as substitute for spruce in same sizes or in slightly reduced size s providing reductions are substantiated. Difficult to work with handtools.

Some tendency to split and splinter during fabrication and considerable more care in manufacture is necessary. Large solid pieces should be avoided due to ins pection difficulties.

Gluing satisfactory.

Noble Fir (Abies Nobiles).

Slightly exceeds spruce except 8% deficient in shear.

1:15 Satisfactory characteri stics with respect to workability, warping, and splitting. May be used as direct substitute for spruce in same sizes providing shear does not become critical. Hardness somewhat less than spruce. Gluing satisfactory.

Western Hemlock (Tsuga Heterpphylla).

Slightly exceeds spruce.

1:15 Less uniform in texture than spruce. May be used as direct substitute for spruce. Upland growth superior to lowland growth. Gluing satisfactory.

Pine, Northern White (Pinus Strobus).

Properties between 85 % and 96 % those of spruce.

1:15 Excellent working qualities and uniform in properties, but somewhat low in hardness and shock-resisting capacity. Cannot be used as substitute for spruce without increase in sizes to compensate for lesser strength. Gluing satisfactory.

White Cedar, Port Orford (Charaecyparis Lawsoniana).

Exceeds spruce.

1:15 May be used as substitute for spruce in same sizes or in slightly reduced size s providing reductions are substantiated. Easy to work with handtools. Gluing difficult, but satisfactory joints can be obtained if suitable precautions are taken.

Poplar, Yellow (Liriodendrow Tulipifera).

Slightly less than spruce except in compression (crushing) and shear.

1:15 Excellent working qualities. Should not be used as a direct substitute for spruce without carefully accounting for slightly r educed strength properties.

Somewhat low in shock-resisting capacity. Gluing satisfactory.

Notes for Table 1-1

1. Defects Permitted.

a. Cross grain. Spiral grain, diagonal grain, or a combination of the two is acceptable providing the grain does not diverge from the longitudinal axis of the material more than specified in co lumn 3. A check of all four faces of the board is necessary to determine the amount of divergence. The direction of fr ee-flowing ink will frequently assist in determining grain direction.

b. Wavy, curly, and interlocked grain. Acceptable, if local irregularities do not exceed limitations specified for spiral and diagonal grain.

c. Hard knots. Sound, hard knots up to 3/8 inch in maximum diam eter are acceptable providing: (1) they are not projecting portions of I-beams, along the edges of rectangular or bevel ed unrouted beams, or along the edges of flanges of box beams (except in lowly stressed portions); (2) they do not cause grain divergence at the edges of the board or in the flanges of a beam more than specified in column 3; and (3) they are in the cent er third of the beam and are not closer than 20 inches to anot her knot or other defect (pertains to 3/8 inch knots—smaller knots may be proportionately closer). Knots greater than 1/4 inch must be used with caution.

d. Pin knot clusters. Small clusters are acceptable providi ng they produce only a small effect on grain direction.

e. Pitch pockets. Acceptable in center porti on of a beam providing they ar e at least 14 inches apart when they lie in the same growth ring and do not exceed 1-1/2 inches length by 1/8 inch widt h by 1/8 inch depth, and providing they are not al ong the projecting portions of I-beams, along the edges of rectangular or beveled unrouted beams, or along the edges of the flanges of box beams.

f. Mineral streaks. Acceptable, providing careful inspection fails to reveal any decay.

### Page 66

Par 1-2 Page 1-66 TABLE 1-1. Selection and Properties of Aircraft Wood. (See notes following table.) (continued)

2. Defects Not Permitted.

a. Cross grain. Not acceptable, unless within limitations noted in 1a.

b. Wavy, curly, and interlocked grain. Not acceptable, unless within limitations noted in 1b.

c. Hard knots. Not acceptable, unless within limitations noted in 1c.

d. Pin knot clusters. Not acceptable, if they produce large effect on grain direction.

e. Spike knots. These are knots running completely thr ough the depth of a beam per pendicular to the annual rings and appear most frequently in quarter-sawed lumber. Reject wood containing this defect.

f. Pitch pockets. Not acceptable, unless within limitations noted in 1e.

g. Mineral streaks. Not acceptable, if accompanied by decay (see 1f).

h. Checks, shakes, and splits. Checks are longitudinal cracks extending, in general, across the annual rings.

Shakes are longitudinal cracks usually between two annual rings . Splits are longitudinal cracks induced by artificially induced stress. Reject wood containing these defects.

i. Compression wood. This defect is very detrimental to strengt h and is difficult to recognize readily. It is characterized by high specific grav ity, has the appearance of an excessive growth of summer wood, and in most species shows little contrast in color between spring wood and summer wood. In doubtful ca ses reject the material, or subject samples to toughness machine test to establish the quality of the wood. Reje ct all material containing compression wood.

j. Compression failures. This defect is caused from the w ood being overstressed in compression due to natural forces during the growth of the tree, felling trees on rough or irregul ar ground, or rough hand ling of logs or lumber. Compression failures are characterized by a buckli ng of the fibers that appear as streaks on the surface of the piece substantially at right angles to the grain, and vary from pronounced failures to very fine hairlines that require close inspection to detect. Reject wood containing obvious failures. In doubtfu l cases reject the wood, or make a further inspection in the form of microscopic examination or toughness test, the latter means being the more reliable.

k. Decay. Examine all stains and discoloration carefully to det ermine whether or not they are harmless, or in a stage of preliminary or advanced decay. All pieces must be free from rot, dote, red heart, purple heart, and all other forms of decay.

fittings and wire bracing and checking or splitting of wood members. A few suggestions for minimizing these shrinkage effects are:

Use bushings that are slightly short so that when the wood member shrinks the bushings do not protrude and the fittings may be tightened firmly against the member.

Gradually drop off plywood faceplates by feathering as shown in figure 1-2.

Thoroughly seal all wood surfaces, particularly end grain and bolt holes, with varnish, epoxy, or other acceptable sealer to slow or prevent moisture changes in the member. (See Section 5. Finishing Wood Structures.)

MODIFIED WOOD PRODUCTS. The most common forms of modified woods found in aircraft construction are plywood. Although not a wood product, Phenolic parts

FIGURE 1-2. Tapering of faceplate.

are sometimes incorporated into structures. These products are used whenever the manufacturer requires specialized strength or durability characteristics.

ADHESIVES. Because of the critical role played by adhesives in aircraft structure, the mechanic must employ only those types of adhesives that meet all of the performance requirements necessary for use in certificated civil aircraft. Use each product strictly in accordance with the aircraft and adhesive manufacturer’s instructions.

### Page 67

Par 1-2 Page 1-67 Adhesives acceptable to the FAA can be identified in the following ways:

Refer to the aircraft maintenance or repair manual for specific instructions on acceptable adhesive selection for use on that type aircraft.

Adhesives meeting the requirements of a Military Specification (Mil Spec), Aerospace Material Specification (AMS), or Technical Standard Order (TSO) for wooden aircraft structures are satisfactory providing they are found to be compatible with existing structural materials in the aircraft and the fabrication methods to be used in the repair.

Common types of adhesives that are or have been used in aircraft structure fall into two general groups: casein and synthetic-resins. Adhesive technology continues to evolve, and new types (meeting the requirements of paragraph 1-4a) may become available in the future.

Casein adhesive performance is generally inferior to other available products. These adhesives should be considered obsolete for all repairs.

CAUTION: Casein adhesive deteriorates over the years after exposure to moisture in the air and temperature variations. Some modern adhesives are incompatible with casein adhesive. If a joint that has previously been bonded with casein is to be rebonded with another type adhesive, all traces of the casein must be scraped off before the new adhesive is applied. If any casein adhesive is left, residual alkalinity may cause the new adhesive to fail to cure properly.

Synthetic-resin adhesives comprise a broad family which includes plastic resin glue, resorcinol, hot-pressed Phenol, and epoxy.

Plastic resin glue (urea-formalde-hyde resin glue) has been used in wood aircraft for many years. Caution should be used due to possible rapid deterioration (more rapidly than wood) of plastic resin glue in hot, moist environments and under cyclic swell-shrink stress. For these reasons, urea-formaldehyde should be considered obsolete for all repairs.

Any proposed use of this type adhesive should be discussed with the appropriate FAA office prior to using on certificated aircraft.

Resorcinol adhesive (resorcinol-formaldehyde resin) is a two-part synthetic resin adhesive consisting of resin and a hardener.

Resorcinol is widely used in wooden aircraft structure and fully meets necessary strength and durability requirements. The appropriate amount of hardener (per manufacturer’s instruction) is added to the resin, and it is stirred until it is uniformly mixed; the adhesive is now ready for immediate use. Quality of fit and proper clamping pressure are both critical to the achievement of full joint strength. The adhesive bond lines must be very thin and uniform in order to achieve full joint strength.

CAUTION: Read and observe material safety data. Be sure to follow the manufacturer’s instructions regarding mixing, open assembly and close assembly times, and usable temperature ranges.

Phenol-formaldehyde adhesive is commonly used in the manufacturing of aircraft grade plywood. This product is cured at elevated temperature and pressure; therefore, it is not practical for use in structural repair.

### Page 68

Par 1-4 Page 1-68 Epoxy adhesives are a two-part synthetic resin product, and are acceptable providing they meet the requirements of paragraph 1-4a. Many new epoxy resin systems appear to have excellent working properties. They have been found to be much less critical of joint quality and clamping pressure. They penetrate well into wood and plywood. However, joint durability in the presence of elevated temperature or moisture is inadequate in many epoxies. The epoxy adhesives generally consist of a resin and a hardener that are mixed together in the proportions specified by the manufacturer.

Depending on the type of epoxy, pot life may vary from a few minutes to an hour. Cure times vary between products.

CAUTION: Some epoxies may have unacceptable thermal or other hidden characteristics not obvious in a shop test. It is essential that only those products meeting the requirements of paragraph 1-4a be used in aircraft repair. Do not vary the resin-tohardener ratio in an attempt to alter the cure time. Strength, thermal, and chemical resistance will be adversely affected. Read and observe material safety data. Be sure to follow the adhesive manufacturer’s instructions regarding mixing, open and closed curing time, and usable temperature ranges.

BONDING PRECAUTIONS. Satisfactory bond joints in aircraft will develop the full strength of wood under all conditions of stress. To produce this result, the bonding operation must be carefully controlled to obtain a continuous thin and uniform film of solid adhesive in the joint with adequate adhesion and penetration to both surfaces of the wood.

Some of the more important conditions involve:

Properly prepared wood surfaces.

Adhesive of good quality, properly prepared, and properly selected for the task at hand.

Good bonding technique, consistent with the adhesive manufacturer’s instructions for the specific application.

PREPARATION OF WOOD SURFACES FOR BONDING. It is recommended that no more time than necessary be permitted to elapse between final surfacing and bonding. Keep prepared surfaces covered with a clean plastic sheet or other material to maintain cleanliness prior to the bonding operation. The mating surfaces should be machined smooth and true with planers, joiners, or special miter saws.

Planer marks, chipped or loosened grain, and other surface irregularities are not permitted.

Sandpaper must never be used to smooth softwood surfaces that are to be bonded.

Sawn surfaces must approach well-planed surfaces in uniformity, smoothness, and freedom from crushed fibers. It is advisable to clean both joint surfaces with a vacuum cleaner just prior to adhesive application.

Wood surfaces ready for bonding must be free from oil, wax, varnish, shellac, lacquer, enamel, dope, sealers, paint, dust, dirt, adhesive, crayon marks, and other extraneous materials.

Roughening smooth, well-planed surfaces of normal wood before bonding is not recommended. Such treatment of well-planed wood surfaces may result in local irregularities and objecti onable rounding of edges.

When surfaces cannot be freshly machined before bonding, such as plywood or inaccessible members, very slight sanding of the surface with a fine grit such as 220, greatly improves penetration by the adhesive of aged or polished

### Page 69

Par 1-4 Page 1-69 surfaces. Sanding should neve r be continued to the extent that it alters the fl atness of the surface. Very light sanding may also improve the wetting of the adhesive to very hard or resinous materials.

Wetting tests are useful as a means of detecting the presence of wax, old adhesive, and finish. A drop of water placed on a surface that is difficult to wet and thus difficult to bond will not spread or wet the wood rapidly (in seconds or minutes). The surface may be difficult to wet due to the presence of wax, exposure of the surface to heat and pressure as in the manufacture of hot press bonded plywood, the presence of synthetic resins or wood extractives, or simply chemical or physical changes in the wood surface with time. Good wettability is only an indication that a surface can be bonded satisfactorily. After performing wetting tests, allow adequate time for wood to dry before bonding.

Preliminary bonding tests and tests for bond strength are the only positive means of actually determining the bonding characteristics of the adhesive and material combinations. (See paragraph 1-29h.)

APPLYING THE ADHESIVE. To make a satisfactory bonded joint, spread the adhesive in a thin, even layer on both surfaces to be joined. It is recommended that a clean brush be used and care taken to see that all surfaces are covered. Spreading of adhesive on only one of the two surfaces is not recommended.

Be sure to read and follow the adhesive manufacturer’s application instructions.

**ASSEMBLY TIME IN BONDING.**

Resorcinol, epoxy, and other adhesives cure as a result of a chemical reaction. Time is an important consideration in the bonding process. Specific time constraints are as follows:

Pot life is the usable life of the adhesive from the time that it is mixed until it must be spread onto the wood surface. Once pot life has expired, the remaining adhesive must be discarded. Do not add thinning agents to the adhesive to extend the life of the batch.

Open assembly time is the period from the moment the adhesive is spread until the parts are clamped together.

Where surfaces are coated and exposed freely to the air, some adhesives experience a much more rapid change in consistency than whe n the parts are laid together as soon as the spreading has been completed.

Closed assembly time is the period from the moment that the structure parts are placed together until clamping pressure is applied. The consistency of the adhesive does not change as rapidly when the parts are laid together.

Pressing (or clamping) time is the period during which the parts are pressed tightly together and the adhesive cures. The pressing time must be sufficient to ensure that joint strength is adequate before handling or machining the bonded structure.

NOTE: Follow the adhesive manufacturer’s instructions for all time limits in the bonding process. If the recommended open or closed assembly periods are exceeded, the bond process should not be continued. Discard the parts if feasible. If the parts cannot be discarded, remove the partially cured adhesive and clean the bond line per adhesive manufacturer’s instructions before application of new adhesive.

BONDING TEMPERATURE. Temperature of the bond line affects the cure rate of the adhesive.

Some adhesive types, such as resorcinol, require a minimum temperature which must be maintained throughout the

### Page 70

Par 1-9 Page 1-70 curing process. Each type of adhesive requires a specific temperature during the cure cycle , and the manufacturer’s recommendations should be followed.

**CLAMPING PRESSURE.**

Use the recommended pressure to squeeze adhesive out into a thin, continuous film between the wood layers. This forces air from the joint and brings the wood surfaces into intimate contact. Pressure should be applied to the joint before the adhesive becomes too thick to flow and is accomplished by means of clamps, presses, or other mechanical devices.

Nonuniform clamping pressure commonly results in weak and strong areas in the same joint. The amount of pressure required to produce strong joints in aircraft assembly operations varies with the type of adhesive used and the type of wood to be bonded. Typical pressures when using resorcinol may vary from 125 to 150 pounds per square inch for softwoods and 150 to 200 pounds per square inch for hardwoods. Insufficient pressure or poorly machined wood surfaces usually result in thick bond lines , which indicate a weak joint, and should be carefully guarded against. Some epoxy adhesives require much less clamping pressure to produce acceptable joint strength. Be sure to read and follow the manufacturer’s instructions in all cases.

METHOD OF APPLYING PRESSURE. The methods of applying pressure to joints in aircraft bonding operations range from the use of brads, nails, small screws, and clamps; to the use of hydraulic and electrical power presses. The selection of appropriate clamping means is important to achieving sound bond joints.

Hand nailing is used rather extensively in the bonding of ribs and in the application of plywood skins to the wing, control surfaces , and fuselage frames. Small brass screws may also be used advantageously when the particular parts to be bonded are relatively small and do not allow application of pressure by means of clamps.

Both nails and screws produce adverse after effects.

There is considerable risk of splitting small parts when installing nails or screws. Metal fasteners also provide vulnerable points for moisture to enter during service.

On small joints using thin plywood for gussets or where plywood is used as an outer skin, the pressure is usually applied by nailing or stapling. Thin plywood nailing strips are often used to spread the nailing pressure over a larger area and to facilitate removal of the nails after the adhesive has cured.

The size of the n ails must vary with the size of the members. If multiple rows of nails are required, the nails should be 1 inch apart in rows spaced 1/2 inch apart. The nails in adjacent rows should be staggered.

In no case should the nails in adjacent rows be more than 3/4 inch from the nearest nail. The length of the nails should be such that they penetrate the wood below the joint at least 3/8 inch. In the case of small members, the end of the nail should not protrude through the member below the joint. Hit the nails with several light strokes, just seating the head into the surface of the gusset. Be careful not t o crush the wood with a heavy hammer blow.

In some cases the nails are removed after adhesive cure, while in others the na ils are left in place. The nails are employed for clamping pressure during adhesive cure and must not be expected to hold members together in service. In deciding whether to re

### Page 71

Par 1-18 Page 1-71 (and 1-10) move nails after assembly, the mechanic should examine adj acent structure to see whether nails remain from original manufacture.

On larger members (spar repairs for example), apply pressure by means of screw clamps, such as a cabinet-maker’s bar or “C-clamps.” Strips or blocks should be used to distribute clamping pressure and protect members from local crushing due to the

### Page 72

Par 1-18 Page 1-72 (and 1-10) limited pressure area of the clamps, especially when one member is thin (such as plywood). The strip or block should be at least twice as thick as the thinner member being bonded.

Immediately after clamping or nailing a member, the mechanic must exam ine the entire joint to assure uniform part contact and adhesive squeeze-out. Wipe away excess adhesive.

ASSEMBLY TIME IN BONDING. Where pieces of wood are coated and exposed freely to the air, some adhesives experience a much more rapid change in consistency than where the pieces are laid together as soon as the spreading has been done. The condition of free exposure is conveniently referred to as “open assembly” and the other as “closed assembly.”

When room temperature curing adhesives are coated on wood parts and left exposed to the atmosphere (open assembly), the time for co mplete assembly is appreciably reduced compared wi th closed assembly periods. Use assembly time and clamping pressures recommended by the adhesive manufacturer.

Resorcinol adhesive cures at “room temper atures,” but the proper cure temperature is crit ical. A minimum cure temperature of 70 °F must be main tained for several hours. Other types of adhesive require various times and temperatures for curing. Refer to manufacturer’s instructions.

Resorcinol, epoxy, and other adhesives increase in strength mainly as a result of a chemical reaction. The pressing time for room temperature curing adhesives should be 8 hours or mo re to ensure the cure is complete before handling or machining the bonded structure. The temperature of the wood and the surrounding air during the cure period should be at least 70°F for the entire pressing period unless the adhesive manufact urer’s directions state that a lower temperature is permissible.

CLAMPING PRESSURE. Use the recommended pressure to squeeze adhesive out into a thin continuous film between the wood layers. This forces air from the joint, and brings the wood surfaces into intimate contact with the adhesive, to hold them in this position during the curing of the adhesive. Too much pressure and not enough assembly time, is often as much of a problem as too little pressure, applied over a long time period.

Pressure should be applied to the joint before the adhesive becomes too thick to flow and is accomplished by means of clamps, presses, or other mechanical devices.

Non-uniform clamping pressure commonly r esults in both weak and strong areas in the same joint. The amount of pressure required to pr oduce strong joints in aircraft assembly oper ations varies with the type of adhesive used and the type of wood to be bonded. Typical pre ssures when using resorcinol or pl astic resin a dhesive may vary from 125 to 150 pounds per square inch for softwoods and 150 to 200 pounds per square inch for hardwoods. Insu fficient pressure or poorly machined wood su rfaces usually result in thick bond lin es, which indicate a weak joint, and should be carefully guarded against. Some epoxy adhesives require much le ss clamping pressure to produce accep table joint strength. Be sure to read and follow the adhesive manufacturer’s instructions in all cases.

METHOD OF APPLYING PRESSURE to joints in aircraft bonding operations range from the use of brads, nails, small screws, and clamps to the use of hydraulic and electrical power presses. The selection of appropriate clamping means is important to achieving sound bond joints.

Hand nailing is used rather extensively in the bonding of ribs and in the application of pl ywood skins to the wing, control surfaces, and fuselage frames. Small brass screws may also be used advantageously when the particular parts to be bonded are relatively small and do not allow application of pressure by means of clamps.

Both nails and screws produce adverse after effects. There is considerable risk of splitting small parts when installing nails or screws. Metal fasteners also provide vulnerable points for moisture entry in service.

On small joints using thin plywood for gussets or where pl ywood is used for single covering, the pressure is usually applied by using nails or small brass screws. Cloth or thin plywood nailing strips are often used to facilitate removal of the nails after the adhesive has cured. Plywood strips also serve to spread the nailing pressure over a larger area.

The size of the nails must vary with the size of the members. If multiple rows of nails are r equired, the nails should be 1 inch apart in rows spaced 1/2 inch apart. The nails in adjacent rows should be staggered. In no case should the nails

### Page 73

Par 1-18 Page 1-73 (and 1-10) be more than 3/4 inch from the next nearest nail. The length of the nails should be such that they penetrate the wood below the joint at least 3/8 inch. In the case of small members, the end of the nail should not pr otrude through the member below the joint. Hit the nails with several light strokes, just embe dding the head flush with the nailing strip.

Be careful not to crush the wood with a heavy hammer blow.

In some cases the nails are removed after adh esive cure, while in others the nails are left in place. The nails are employed for clamping pressure during adhesive cure a nd must not be counted upon to hold members together in service. In deciding whether or not to remove nails after assembly, the mechanic should exa mine adjacent structure to see whether nails remain from original manufacture.

On larger wood to wood or wood to plywood members, for example in spar repairs, apply pressure by means of screw clamps, such as cabinet maker’s bar, or “C” clamps. Strips or blocks should be used to distribute clamping pressure and protect members from local crushing due to the limited pressure area of the clamps, especially when one member is thin (such as plywood). The strip or block should be at least twice as thick as the thinner member being bonded.

Immediately after clamping or nailing a member, the mechanic must examine the entire joint to assure uniform part contact, glue line, and adhesive squeeze-out. Wipe away excess adhesive.

1-12.1-17. [RESERVED.]

### Page 74

Par 1-18 Page 1-74 (and 1-10) SECTION 2. HEALTH AND SAFETY

GENERAL. The possibility of an injury is an important consideration when working with wooden aircraft structures. The tools and machines used to shape wooden members can be very dangerous. In addition, there are potential health hazards in working with adhesives and finishes. The mechanic should follow manufacturer’s instructions wherever applicable to prevent injury. Federal law mandates that individual chemical manufacturers are to provide Material Safety Data Sheets (MSDS) with health hazard data to all consumers. First aid information and handling precautions must also be identified.

Most of the products used in wooden aircraft construction are flammable. Some, such as dope and paint, may be highly flammable.

SANDING IN AREAS OF EXISTING BOND JOINTS AND FINISHES. Some adhesives used in wooden aircraft construction contain biocides. A commonsense precaution when machining or sanding existing structure is to wear a respirator to avoid inhaling dust products. To lower potential fire hazards avoid using electric sanders around dope, paints, and adhesives.

**HANDLING OF ADHESIVES AND FINISHES.**

Most adhesives and finish products present at least some toxic potential to users. Injury may occur from skin or eye contact, inhalation, or accidental ingestion. Users should be aware of the manufacturer’s instructions and MSDS.

Appropriate skin, eye, ear, and respiratory protection should be worn whenever indicated.

Shop cleanliness is essential for health and fire safety.

Shop personnel should maintain awareness of others in the work area to assure that bystanders are not injured.

Proper shop ventilation is essential to disperse fumes emitted from adhesives such as resorcinol and epoxy.

1-21.1-26. [RESERVED.]

### Page 75

Par 1-27 Page 1-75

**SECTION 3. INSPECTION**

GENERAL. Inspection of wooden structure includes some methods, equipment, and awareness of failure modes which are unique to wooden aircraft.

**TYPES OF DETERIORATION AND DAMAGE.**

Wood Decay. Wood is an organic product which is subject to attack by fungi. Fungi are plants that grow on and in wood. The moisture content of the wood nominally will have to be 20 percent or greater to sustain fungus growth. The result of this growth is called decay. Decayed wood exhibits softness, swelling if still wet, excessive shrinkage when dry, cracking, and discoloration. Repair or replace wood if any amount or form of decay is found.

Splitting. Splits or cracks in wooden members occur along grain lines. When the moisture content of wood is lowered, its dimensions decrease. The dimensional change is greatest in a tangential direction (across the fibers and parallel to the growth rings), somewhat less in a radial direction (across the fibers and perpendicular to the growth rings), and is negligible in a longitudinal direction (parallel to the fibers). These dimensional changes can have detrimental effects upon a wood structure, particularly when two parts are bonded together with grains in different directions. This effect can often be seen where a plywood doubler is bonded to a spruce member. As the spruce member dries, it attempts to shrink, but is restrained by the plywood, which shrinks less. The re sulting stress in the spruce member exceeds its cross-grain strength, and a split occurs.

Bond Failure. Bond joint failure is generally due to improper fabrication technique or prolonged exposure to moisture in service. Although none of the older adhesives have been specifically found to fail by simple aging, the mechanic is advised to in spect all accessible joints carefully.

Finish Failure. The finish coat on wood structure (usually varnish) is the last line of defense to prevent water entry into wood and th e resulting decay. Finish failure can be the result of prolonged water exposure, wood splitting, ultraviolet li ght exposure, or surface abrasion.

Damage. Stress, impact, or mechanical damage to a wood structure is caused by excessive aerodynamic loads or impact loads occurri ng while the aircraft is on the ground. Overtightening of fittings can also cause crushing of the underlying wood member and possible bending of the metal fitting.

INSPECTION METHODS. Whenever possible, the aircraft should be kept in a dry, wellventilated hangar, with all inspection covers, access panels, etc., removed for as long as possible before final inspection. The aircraft should be given a preliminary inspection when first removing the inspection covers and access panels and inspected with a moisture meter at this time. If the moisture content is high, the aircraft should be thoroughly dried. If the aircraft is dry, this will facilitate later inspection, especially when determining the condition of bonded joints.

Likely locations for wood structure deterioration should be given special attention. Most damage is caused by external influence such as moisture, temperature extremes, or sunlight. Care should be taken to note all possible entry points for moisture, (i.e., cracks or breaks in the finish, fastener holes, inspection/access openings, c ontrol system openings, drain holes, and the interfaces of metal fittings

### Page 76

Par 1-27 Page 1-76 and the wood structure). The mechanic should also look for evidence of swelling or warpage of the aircraft’s wood structure, which would indicate underlying damage or decay. Particular attention should be paid to the wood structure immediately beneath the upper surfaces, especially under areas that are finished in dark colors, for signs of deteriorating adhesives. Cracks in wood spars are often hidden under metal fittings or metal rib flanges and leading edge skins. Any time a reinforcemen t plate exists that is not feathered out on its ends, a stress riser exists at the ends of the plate. A failure of the primary structure can be expected to occur at this point.

Tapping the wood structure with a light plastic hammer or screwdriver handle should produce a sharp solid report. If the suspect area sounds hollow and soft, further inspection is warranted by the following methods.

Probe the area in question, if accessible, with a sharp metal tool. The wood structure should be solid and firm. If the suspect area feels soft and mushy the mechanic should assume that the area is rotted.

Disassembly of the structure is warranted at this point.

Prying the area of a bond joint will reveal any mechanical separation of th e joint. If the mechanic detects any relative moveme nt between two adjacent wood members, a failure of the bond is evident. Any loose fittings should arouse the mechanic’s suspicion, and the fittings should be removed to check for elongated bolt holes. Disassembly is warranted for further inspection.

Odor is an important indicator of possible deterioration.

During the initial inspection, as the access panels are being removed from the structure, the mechanic should be aware of any areas that smell musty or moldy.

These odors are indicative of the presence of moisture and associated fungal growth and decay.

Visual inspection requires looking at the wood structure both externally and internally for visual signs of decay or physical damage. Any accumulations of dirt, bird nests, or rodent nests are likely places to hold moisture and promote decay.

The mechanic should remove any such accumulations that are found and inspect the area for signs of decay. Decay will appear as a dark discoloration or gray stains running along the grain and often a swelling of the wood member if still wet. Fittings will be imbedded in the wood instead of flush.

Highly suspected structurally damaged areas are shown in figure 1-3. A list of most likely areas to incur structural damage include the following:

Check front and rear spars for compression cracks adjacent to the plywood reinforcing plates, where the lift struts attach, and at the rib attach points on either side of the strut attach points. Triple-check these areas and the spar to fuselage attach points for cracks if the wingtip has contacted the ground, a hangar wall, etc.

Check all metal fittings which attach to wooden structure for looseness, corrosion, cracks, or warps. Areas of particular interest are strut attach fittings, spar butt fittings, aileron and flap hinges, jury strut fittings, compression struts, pulley brackets, and any landing gear fittings.

Check front and rear spars for longitudinal cracks at the ends of the plywood reinforcement plates where the lift struts attach. Triple-check this area if the wing has encountered any kind of ground strike.

### Page 77

Par 1-29 Page 1-77 Check ribs on either side of strut attach points for missing or loose rib-to-spar attach nails.

Check ribs on either side of strut attach points for cracks where the cap strips pass over and under the spars.

Check for cracked leading edge skin and/or failed nose ribs in the area directly in front of the jury strut.

Check the brackets which attach the struts to the spars for cracks.

Check the aileron, flap hinge, and hinge brackets for cracks and loose or missing rivets.

Check all exposed end grain wood, particularly the spar butts, for cracking or checking. Checking, or splitting, of wood spar butts is common on aircraft based in arid areas.

FIGURE 1-3. Likely areas to incur structural damage.

Also check for any cracks that indicate a bond line failure or structural failure of the wood member. Any evidence of movement of fittings, bushings, or fasteners should be cause for concern, and further inspection is warranted. Splits in fabric covering the plywood, especially on upper surfaces exposed to ultraviolet light and water, dictate that the mechanic remove the fabric around the split so the underlying plywood may be inspected for physical damage or decay. When removing metal fasteners from wood, check for evidence of corrosion. Any corrosion present indicates the presence of moisture and the strong probability of decay in the adjoining wood structure.

Any wooden member that has been overstressed is subject to compression failure (e.g., ground loop). Compression cracking and failure of the wood spars in certain aircraft are a continuing problem.

Compression failures are defined as failure of wood fibers on a plane perpendicular to the wood fiber’s longitudinal axis. If undetected, compression failures may result in structural failure of the wing during flight. Compression cracks have been found emanating from the upper surfaces of the wing spars and progressing downward.

The usual locations for cracks have been the front spar at both ends of the reinforcement plate for the lift strut and the front spar rib attach points, both inboard and outboard of the spar reinforcement plate; and the rear spar lift strut and rib attach points. An inspection of both the front and rear spars for compression cracks is recommended.

The two areas where it is possible to identify a compression crack are on the face and top surface of the spar. Using a borescope through existing inspection holes is one method of inspection. An alternate method is to cut inspection holes in the skin.

If inspection holes are cut, they should be made on the aft side of the front spar and the forward side of the rear spar. This will allow the fabric to be peeled away from the spar. Longitudinal cracks may also be detected during this inspection. Loose or missing rib nails may indicate further damage and should be thoroughly investigated. The mechanic may shine a light, at a low angle and parallel with the grain, in the area of the member

### Page 78

Par 1-29 Page 1-78 subjected to the compression load. An area of grain waviness would indicate a potential compression failure. In all cases the manufacturer’s inspection data should be followed.

Moisture Meters are effective tools for detection of excessive moisture content in wood members. An instrument such as this allows the mechanic to insert a probe into the wood member and read its moisture content directly off the meter. A correction chart usually accompanies the inst rument to correct for temperature and species of wood. Any reading over 20 percent indicates the proba bility of fungus growth in the member. Moisture content of the wood should be 8-16 percent, preferably in the 10-12 percent range (this range is during inspection). Where plywood skin covers the spar and the spar would be inaccessible without removing the skin, the moisture meter probe can be inserted through the plywood skin and into the spar to check the moisture content of the spar. The small holes made by the probe are easily sealed.

Destructive testing of sample bonded joints whenever a new bond joint is made, a sample joint should be made with the adhesive from the same batch used on the repair and scraps of wood left over from the repair.

After curing, the sample joint should be destructively tested to ensure proper bonding of the two wood pieces.

Any failure in the bond line indicates a cohesive failure of the adhesive. Any failure along the bond line indicates an adhesive failure, which is indicative of poor bonding. The ideal situation is when wood fibers are observed on both sides of the fracture surface. This indicates a failure in the wood, and indicates the bond joint is actually stronger than the wood.

1-30.1-35. [RESERVED.]

### Page 79

Par 1-36 Page 1-79

**SECTION 4. REPAIRS**

GENERAL. The basic standard for any aircraft repair is that the repaired structure must be as strong as the original structure and be equivalent to the original in rigidity and aerodynamic shape.

Repairs should be made in accordance with manufacturer specifications whenever such data is available.

REPLACEMENT OF DRAIN HOLES AND SKIN STIFFENERS. Whenever repairs are made that require replacing a portion that includes drain holes, skin stiffeners, or any other items, the repaired portion must be provided with similar drain holes, skin stiffeners, or items of the same dimensions in the same location. Additional drain holes may be required if reinforcement under a skin repair interferes with waterflow to existing drain holes.

Make any additional drain holes the same diameter as originals, usually 1/4 inch.

**CONTROL SURFACE FLUTTER PRECAUTIONS.**

When repairing or refinishing control surfaces, especially on high-performance airplanes, care must be exercised that the repairs do not involve the addition of weight aft of the hinge line. Such a procedure may adversely affect the balance of the surface to a degree that could induce flutter. As a general rule, it will be necessary to repair control surfaces in such a manner that the structure is identical to the original, and that the stiffness, weight distribution, and mass balance are not affected in any way. Consult the aircraft maintenance manual or seek manufacturer’s direction for specific requirements on checking control surface balance after repair and refinishing of any control surface.

SCARF JOINTS. The scarf joint is the most satisfactory method of making an end joint between two solid wood members. Cut both parts accurately. The strength of the joints depends upon good joint design and a thin, uniform bond line. Make the scarf cut in the general direction of the grain slope as shown in figure 1-4.

No grain deviation steeper than 1 in 15 should be present in an outer eighth of the depth of the spar. In adjacent eighths, deviations involving steeper slopes, such as a wave in a few growth layers, are unlikely to be harmful. Local grain slope deviations in excess of those specified may be permitted in spar flanges only in the inner one-fourth of the flange depth.

FIGURE 1-4. Consideration of grain direction when making scarf joints.

SPLICING OF SPARS. Unless otherwise specified by the manufacturer, a spar may be spliced at any point except under the wing attachment fittings, landing gear fittings, engine mount fittings, or lift and interplane strut fittings. These fittings may not overlap any part of the splice. A spar splice repair should not be made adjacent to a previous splice or adjacent to a reinforcing plate. Spacing between two splices or between a splice and a reinforcing plate should be no less than three times the length of the longer splice.

### Page 80

Par 1-36 Page 1-80 Splicing under minor fittings such as drag wire, antidrag wire, or compression strut fittings is acceptable under the following conditions:

The reinforcement plates of the splice should not interfere with the proper attachment or alignment of the fittings. Do not alter the locations of pulley support brackets, bellcrank support brackets, or control surface support brackets. Plates are to be tapered off, as depicted in figure 1-2.

The reinforcement plate may overlap drag wire, antidrag wire, or compression strut fittings, if the reinforcement plates are on the rear face of the rear spar or the front face of the front spar. In such cases, it will be necessary to install slightly longer bolts. The front face reinforcement plate shoul d not overlap drag strut fittings, except when it does not require sufficient shortening of compression struts or changes in dragtruss geometry, to prevent adjustment for proper rigging. Even though take up is sufficient, it may be necessary to change the angles on the fittings.

(Acceptable methods for splic ing the various types of spars are shown in figure 1-4 through figure 1-9.)

Reinforcement plates must be used as indicated on all scarf repairs to spars and the slopes of scarves shown are minimum slopes.

SPAR REPLACEMENT. Replacement of spars is a major repair. Spars may be replaced by new parts made by the manufacturer or the holder of a Parts Manufacturer Approval (PMA) for that part.

Owner-produced spars may be installed providing they are made from a manufacturer-approved drawing. Also, a spar may be made by reference to an existing spar providing sufficient evidence is presented to verify that the existing spar is an original part, and that all materials and dimensions can be determined. The dimensions and type of wood used are critical to the structural strength of the aircraft. Care should be taken that any replacement spars accurately match the manufacturer’s original design.

SPLICING OF BOX SPAR WEBS. Always splice and reinforce plywood webs with the same type of plywood as found on the original part. Do not use solid wood to replace plywood webs. Plywood is stronger in shear than solid wood of the same thickness due to the grain direction of the individual plies. The face-grain of plywood replacement webs and reinforcement plates must be in the same direction as the original member to ensure that the new web will have the required strength. (The method of splicing plywood webs is shown in figure 1-9.)

REPLACING SOLID-TYPE SPARS WITH LAMINATED-TYPE SPARS. Solid spars may be replaced with laminated spars or vice versa, provided the material is of the same high quality.

External reinforcements (plywood or solid) must always be replaced as on the original member.

SPAR LONGITUDINAL CRACKS AND LOCAL DAMAGE. Cracked spars (except box spars) may be repaired by bonding plates of spruce or plywood of sufficient thickness to develop the longitudinal shear on both sides of the spar. Extend the plates well beyond the termination of the cracks, as shown in figure 1-10. A method of repairing small local damage to either the top or bottom side of a spar is also shown in figure 1-10.

Longitudinal Cracking of Wood Wing Spars of Aircraft Operating in Arid Regions. Aircraft having wood spars and operating in arid regions may develop longitudinal spar cracks in the vicinity of the plywood reinforcement plates. Th ese cracks result from the tendency of the spar to shrink when drying takes place.

Plywood resists this tendency to

### Page 81

Par 1-44 Page 1-81

Reinforcement plates to be spruce or plywood and should be bonded only, with the ends feathered off at 5:1 slope as depicted in Figure 1-2.

Solid spars may be replaced with laminated spars or vice versa, provided the material is of the same high quality.

FIGURE 1-5. Method of splicing solid or laminated rectangular spars.

shrink and causes the basic spar stock to split (see paragraph 1-2c). Cracks start under the plywood plates, usually (but not necessarily) at a bolt hole or cutout, and usually spread in each direction until they extend a short distance beyond the ends of the plates where the resistance to spar shrinkage disappears. Cracks have also been found in the butt end of spars. Other factors, which have been found conducive to the formation of cracks are poor protective finishes, large cutouts, and metal fittings that utilize two lines of large diameter bolts.

Repairing Cracks Versus Installing a New Spar. The presence of cracks does not n ecessarily mean that the spar must be discarded. If the crack is not too long or too close to either edge and can be reinforced properly, it will probably be more economical and satisfactory to perform repair rather than install a new spar or section. However, a generally acceptable procedure suitable for all airplane models is not available. Because of the possibility of strength deficiencies contact the manufacturer. In absence of the manufacturer, the FAA should be contacted for approval before making repairs not in accordance with the manufacturer-approved instructions or the recommendations of this advisory circular. Longitudinal cracking or the recurrence of cracking can be minimized by ensuring that the moisture content of the solid wood portion is within the proper range before bonding . In arid desert areas, during bonding the moisture content should be in the range of 6-8 percent before bonding, but in other areas 10-12 percent is satisfactory. If solid or plywood repair stock is procured from another climatic region, it should be allowed to season,

### Page 82

Par 1-44 Page 1-82

If splice is made where routing is feathered to full width of spar, tapered plates (conforming to the contour of the routing) should be added first;

otherwise, the splice is the same as shown.

Reinforcement plates may be of plywood, the same material as the spar, or material of equal or higher quality; and should be bonded only with ends feathered out with a 5:1 slope.

FIGURE 1-6. Method of splicing solid “I” spars.

in the same storage area as th e part to be repaired, for no less than 2 weeks.

Preventing Cracks. An important step in the prevention of longitudinal cracking, particularly in spar butts, is to ensure that the wood is thoroughly sealed with a penetrating and highly moisture-resistant finish.

Application of a thin, slow-curing epoxy adhesive or sealer can be very effectiv e in slowing or preventing moisture changes in spar butts.

ELONGATED HOLES IN SPARS. In cases of elongated bolt holes in a spar, or cracks in the vicinity of bolt holes, splice in a new section of spar, or replace the spar entirely. If hole elongation or cracking is minimal and the bolt holes are for noncritical fittings, repair (rather than replacement) may be feasible. Obtain approval for any such repair from the manufacturer or a representative of the FAA. In many cases, it has been found advantageous to laminate the new section of the spar, particularly if the spar butts are being replaced.

RIB REPAIRS. Ribs may be replaced by new parts made by the manufacturer or the holder of a PMA for that part. Owner-produced ribs may be installed providing they are made from a manufacturer-approved drawing or by reference to an existing original rib. A rib may be made by reference to an existing rib providing sufficient evidence is presented to verify that the existing rib is an original part and that all materials and dimensions can be determined. The contour of the rib is important to the safe flying qualities of the aircraft, and care should be taken that any replacement ribs accurately match the manufacturer’s original design.

### Page 83

Par 1-46 Page 1-83

Ends of reinforcement plates

should be feathered out Section A-A with a 5:1 slope.

FIGURE 1-7. Repairs to built-up “I” spar.

Rib Repair Methods. Acceptable methods of repairing damaged ribs are shown in figure 1-11. Wood ribs should not be nailed to wood spars by driving nails through the rib cap strips, as this weakens the rib materially. The attachment should be by means of adhesive with cement coated, barbed, or spiraled nails driven through the vertical rib members on each face of the spar.

Compression Rib Repair. Acceptable methods of repairing damaged compression ribs are shown in figure 1-12.

Figure 1-12(A) illustrates the repair of a compression rib of the “I” section type; i.e., wide, shallow cap strips, and a center plywood web with a rectangular compression member on each side of the web. The rib is assumed to be cracked through the cap strips, web member, and compression member in the illustration. Cut the compression member as shown in figure 1-12(D). Cut and replace the aft portion of the cap strips, and reinforce as shown in figure 1-11. The plywood side plates are bonded on, as indicated in figure 1-12(A). These plates are added to reinforce the damaged web.

Figure 1-12(B) illustrates a compression rib of the type that is basically a standard rib with rectangular compression members added to one side and plywood web to the other side. The method used in this repair is essentially the same as

### Page 84

Par 1-46 Page 1-84 in figure 1-12(A) except that the plywood reinforcement plate,

A, B, C, D = ORIGINAL DIMENSIONS

FIGURE 1-8. Method of splicing box spar flanges (plate method).

shown in section B-B, is continued the full distance between spars.

Figure 1-12(C) illustrates a compression rib of the “I” type with a rectangular vertical member on each side of the web. The method of repair is essentially the same as in figure 1-12(A) except the plywood reinforcement plates on each side, shown as striped blocks in section C-C, are continued the full distance between spars.

PLYWOOD SKIN REPAIR. Make extensive repairs to damaged stressed skin plywood structures in accordance with specific recommendations from the aircraft manufacturer. It is recommended that repairs be made by replacing the entire panel, from one structural member to the next, if damage is very extensive. When damaged plywood skin is repaired, carefully inspect the adjacent internal structure for possible hidden damage. Repair any defective frame members prior to making skin repairs.

DETERMINATION OF SINGLE OR DOUBLE CURVATURE. Much of the outside surface of plywood aircraft is curved. On such areas, plywood used for repairs to the skin must be similarly curved.

Curved skins are either of single curvature or of double (compound) curvature. A simple test to determine which type of curvature exists may be made by laying a sheet of heavy paper on the surface in question. If the sheet can be made to conform to the surface without wrinkling, the surface is either flat or single curvature. If the sheet cannot be made to conform to the surface without wrinkling, the surface is of double curvature.

REPAIRS TO SINGLE CURVATURE PLYWOOD SKIN. Repairs to single curvature plywood skin may usually be formed from flat plywood, either by bending it dry or after soaking it in hot water. The degree of curvature to which a piece of plywood can be bent will depend upon the direction of the grain and the thickness. Table 1-2 is a guide

### Page 85

Par 1-49 Page 1-85

**1. AFTER INSERTED WEB HAS BEEN BONDED IN PLACE, BOND COVER STRIPS OVER ENTIRE**

**LENGTH OF WEB SPLICE JOINTS.**

**2. SECTIONAL SHAPE OF FILLER BLOCKS MUST CONFORM EXACTLY TO TAPER OF SPAR. THEY MUST**

NOT BE TOO TIGHTLY FITTED, FOR WEDGING ACTION WILL LOOSEN EXISTING BONDS OF WEBS TO FLANGES. IF TOO LOOSELY FITTED, CRUSHING OF WEB WILL OCCUR WHEN CLAMPING.

3. DIRECTION OF FACE GRAIN OF INSERTED WEB SECTION AND COVER STRIPS TO BE SAME AS ORIGINAL WEB.

FIGURE 1-9. Method of splicing box spar webs.

### Page 86

Par 1-49 Page 1-86

FIGURE 1-10. Method of reinforcing a longitudinal crack and/or local damage in a solid spar.

for determining which process of bending should be used for the curvature being considered.

Plywood, after softening, may be bent on a cold ventilated form, or it may be bent over the leading edge near the area being patched if space permits. In either method the repair part should be allowed to dry completely on the form. When bending plywood over a leading edge, drying may be hastened by laying a piece of coarse burlap over the leading edge before using it as a bending form. To speed drying, a fan may be used to circulate air around the repair part.

In bending pieces of small radii or to speed up the bending of a large number of parts of the same curvature, it may be necessary to use a heated bending form. The surface temperature of this form may be as high as 149 °C (300 ºF), if necessary, without danger of damage to the plywood. The plywood should be left on the heated form only long enough to dry to room conditions.

REPAIRS TO DOUBLE CURVATURE PLYWOOD SKIN. The molded plywood necessary for a repair to a damaged plywood skin of double curvature cannot be made from flat plywood unless the area to be repaired is very small or is of exceedingly slight double curvature; therefore, molded plywood of the proper curvature must be onhand before the repair can be made. If molded plywood of the proper curvature is available, the repair may be made using the same procedure as on single curvature skins.

### Page 87 (OCR)

**12F RECOMMENDED 10F MINIMUM**

**DAMAGED AREA**

**ORIGINAL GUSSET**

**PLYWOOD-**

**SPRUCE OR**

BASSWOOD AT A SPAR * 4A → ЗА

- 5A - →C→ • 5A ~ AT TRAILING EDGE 10A A,B,C,D,E,F,G = ORIGINAL DIMENSIONS.

**REINFORCEMENT PLATES**

**SHALL BE PLYWOOD AT A JOINT TO TRUSS MEMBERS THE DIRECTION OF THE GLUED AND NAILED.**

**FACE GRAIN MUST BE**

**THE SAME AS ORIGINAL GUSSET.**

— ЗА • 10A - ЗА - з‡

BETWEEN JOINTS FIGURE 1-11. Repair of wood ribs.

Par 1-50 Page 1-87

### Page 88 (OCR)

**SEE FIGURE 1-11 FOR - REPAIR REPAIR OF CAP STRIPS. DAMAGE**

SEE (D) (А) ЗА - ЗА SEE FIGURE 1-11 FOR DAMAGE REPAIR OF CAP STRIPS. REPAIR (B) SEE FIGURE 1-11 FOR

**REPAIR OF CAP STRIPS REPAIR DAMAGE - (C)**

•12A, BA COINIMENDED 6A 6 A (D)

A-A B-B

PLAND PACE GRAIN DERECTTON AS THIGINESS FIGURE 1-12. Typical wing compression rib repairs.

Par 1-50 Page 1-88

### Page 89

Par 1-50 Page 1-89 TABLE 1-2. Minimum recommended bend radii for aircraft plywood.

**10 PERCENT MOISTURE**

**CONTENT, BENT ON COLD**

**MANDRELS**

**THOROUGHLY SOAKED IN HOT**

**WATER AND BENT ON COLD**

**MANDRELS**

**PLYWOOD**

CHARACTERISTICS AT 90 º TO A T 0 º OR 45º

**FACE GRAIN**

T O F A C E

GRAIN AT 90 º TO A T 0º0R 45°

**FACE GRAIN**

**TO FACE**

**GRAIN**

**THICKNESS (INCHES) NUMBER**

OF PLIES MINIMUM BEND RADIUS (INCHES) .035 3 2.0 1.1 0.5 0.1 .070 3 5.2 3.2 1.5 0.4 .100 3 8.6 5.2 2.6 0.8 .125 3 12 7.1 3.8 1.2 .155 3 16 10 5.3 1.8 .185 3 20 13 7.1 2.6 .160 5 17 11 6 2 .190 5 21 14 7 3 .225 5 27 17 10 4 .250 5 31 20 12 5 .315 5 43 28 16 7 .375 5 54 36 21 10

TYPES OF PATCHES. There are four types of patches: splayed patch, surface (or overlay) patch, scarf patch, and plug patch. They are all acceptable for repairing plywood skins.

Splayed Patch. Small holes with their largest dimensions not over 15 times the skin thickness, in skins not more than 1/10 inch in thickness, may be repaired by using a circular splayed patch as illustrated in figure 1-13. The term “splayed” is used to denote that the edges of the patch are tapered, but the slope is steeper than is allowed in scarfing operations.

Lay out the patch according to figure 1-13. Tack a small piece of plywood over the hole for a center point and draw two circles with a divider, the inner circle to be the size of the hole and the outer circle marking the limits of the taper. The difference between the radii is 5T (5 times the thickness of the skin). If one leg of the dividers has been sharpened to a chisel edge, the dividers may be used to cut the inner circle.

Taper the hole evenly to the outer mark with a chisel, knife, or rasp.

Prepare a circular tapered patch to fit the prepared hole, and bond the patch into place with face-grain direction matching that of the original surface.

Use waxed paper or plastic wrap, (cut larger than the size of the patch) between the patch and the plywood pressure plate. This prevents excess adhesive from bonding the pressure plate to the skin. Center the pressure plate carefully over the patch.

### Page 90

AC 43.13-1B

Par 1-36 Page 1-90 As there is no reinforcement behind this patch, care must be used so that pressure is not great enough to crack the skin. On horizontal surfaces, weights or sandbags will be sufficient. On patches too far from any edge for the use of standard hand clamps, jaws of greater length may be improvised.

Table 1-2, columns (1) and (3), may also be used for determining the maximum thickness of single laminations for curved members.

Fill, sand, and refinish the patch.

Surface Patch. Plywood skins that are damaged between or along framing members may be repaired by surface or overlay patches as shown in figure 1-14.

Surface patches located entir ely aft of the 10 percent chord line, or which wrap around the leading edge and terminate aft of the 10 percent chord line, are permissible. Surface patches may have as much as a 50 inch perimeter and may cover as much as 1 frame (or rib) space. Trim the damage d skin to a rectangular or triangular shape and round the corners. The radius of rounded corners must be at least 5 times the skin thickness. Bevel the forward edges of patches located entirely aft of the 10 percent chord line to 4 times the skin thickness. The face-gr ain direction must be the same as the original skin . Cover completed surface patches with fabric to match surrounding area. The fabric must overlap the original fabric at least 2 inches.

Scarf Patch. A properly prepared and inserted scarf patch is the best repair for damaged plywood skins and is preferred for most skin repairs. Figure 1-15 shows the details and dimensions to be used when installing typical scarf skin patches, wh en the back of the skin is accessible. Follow figure 1- 16 when the back of the skin is not accessible. The scarf slope of 1 in 12, shown in both figures, is the steepest slope permitted for all kinds of plywood. If the radius of curvature of the skin at all points on the trimmed opening is greater than 100 times the skin thickness, a scarf patch may be installed.

Scarf cuts in plywood may be made by hand plane, spoke shave, scraper, or accurate sandpaper block.

Rasped surfaces, except at the corners of scarf patches and sawn surfaces, are not recommended as they are likely to be rough or inaccurate.

Nail strip or small screw clamping is often the only method available for bonding scarf joints in plywood skin repairs. It is essential that all scarf joints in plywood be backed with plywood or solid wood to provide adequate nail holding capacity.

The face-grain direction of the plywood patch must be the same as that of the original skin.

If the back of a damaged plywood skin is accessible (such as a fuselage skin), it should be repaired with a scarf patch, following the details shown in figure 1-15. Whenever possible, the edges of the patch should be supported as shown in section C-C of figure 1-15. When the damage follows or extends to a framing member, the scarf may be supported as shown in section B-B of figure 1-15. Damages that do not exceed 25 times the skin thickness in diameter after being trimmed to a circular shape and are not less than 15 times the skin thickness to a framing member, may be repaired as shown in figure 1-15, section D-D.

The backing block is carefully shaped from solid wood and fitted to the inside surface of the skin, and is temporarily held in place with nails.

Use waxed paper or plastic wrap to prevent bonding of the backing block to the skin.

### Page 91

Par 1-51 Page 1-27

FIGURE 1-13. Splayed patch.

A hole, the exact size of the inside circle of the scarf patch, is made in the block, and is centered over the trimmed area of damage.

The block is removed, after the adhesive on the patch has set, leaving a flush surface to the repaired skin.

Steps in making a scarf patch when the back of the skin is not accessible are as follows:

After removing damaged sections, install backing strips, as shown in figure 1-16, along all edges that are not fully backed by a rib or a spar. To prevent warping of the skin, backing strips should be made of a soft-textured plywood, such as yellow poplar or spruce rather than solid wood. All junctions between backing strips and ribs or spars should have the end of the backing strip supported by a saddle gusset of plywood.

If needed, nail and bond the new gusset plate to rib. It may be necessary to remove and replace the old gusset plate with a new saddle gusset, or it may be necessary to nail a saddle gusset over the original gusset.

Attach nailing strips to hold backing strips in place while the adhesive sets. Use a bucking bar, where necessary, to provide support for nailing. After the backing strips are fully bonded, install the patch.

Plug Patch. Either oval or round plug patches may be used on plywood skins provided the damage can be covered by thepatches whose dimensions are given in figure 1-17 and figure 1-18. The plug patch is strictly a skin repair, and should be used only for damage that does not involve the supporting structure under the skin.

The face-grain direction of the finished patch must match the surrounding skin.

Steps in making an oval plug patch are as follows:

Explore the area about the hole to be sure it lies at least the width of the oval

### Page 92 (OCR)

J!

**FRONT SPAR**

**TRIMMED OPENING**

- MINIMUM RADIUS 5T

- RIBS B 8 DAMAGE →A

**SADDLE GUSSET**

REAR SPAR C→ 30T PATCHI — 12T → • PLYWOOD SKIN 8T ЗT (¼" MINIMUM) RIB CAP (1" MINIMUM)

- PLYWOOD SADDLE GUSSET MINIMUM THICKNESS = T.

**SECTION A-A NAILED AND GLUED IN PLACE.**

12 T→ - PATCH - 12 T—→ PATCH -4T

**UNSUPPORTED RIB CAP LAP**

SECTION B-B SECTION C-C FIGURE 1-14. Surface patches.

Par 1-54 Sec 4 Page 28

### Page 93 (OCR)

**SADDLE GUSSET**

**NAILING STRIPS**

**MAXIMUM DIAMETER 25T**

• MINIMUM DISTANCE 15T

**TEMPORARY**

**BACKING 30T PATCH 12T**

**BACKING**

PLYWOOD SADDLE GUSSET MINIMUM THICKNESS • T

BONDED IN PLACE Section A - (1/4" min.)

**12 T - PATCH PATCH 12T**

**FRAMING FRAMING**

MEMBER MEMBER BACKING BACKING •

**CLAMP AND BOND BACKING CLAMP AND BOND BACKING**

TO FRAME AND SKIN TO FRAME AND SKIN Section B - B NAILING STRIPS T Section C - C 12T WAXED PAPER OR PLASTIC WRAP TEMPORARY BACKING 1

BLOCK-SHAPE TO FIT SKIN Section D - D FIGURE 1-15. Scarf patches (back of skin accessible).

Par 1-51 Page 1-29

### Page 94 (OCR)

FRONT SPAR →C RIBS SADDLE GUSSET +

**REAR SPAR**

- 30 T -12T= PLYWOOD SKIN 8T (1" MINIMUM) ЗT (¼" MINIMUM) - RIB CAP

**PLYWOOD SADDLE GUSSET**

SECTION A-A NAIL AND GLUE IN PLACE PLYWOOD (MINIMUM THICKNESS "T") SKIN -12T→ PATCH *12 T. - PATCH

T SPAR RIB CAP - ЗТ (¼" MINIMUM)

**PLYWOOD OR SPRUCE PLYWOOD SKIN**

SECTION B-B SECTION C-C FIGURE 1-16. Scarf patches (back of skin not accessible).

Par 1-51 Page 1-30

### Page 95

Par 1-51 Page 1-31 doubler from a rib or a spar. Refer to figure 1-17 for repair details.

Prepare a patch and a doubler of the same species plywood as the surrounding skin using the dimensions shown in figure 1-17.

Lay the oval plug patch over the damage and trace the patch onto the skin. Saw to the line, and trim the hole edges with a knife and sandpaper.

Mark the exact size of the patch on one surface of the oval doubler and apply adhesive to the area outside the line. Insert doubler through the hole and bring it, adhesive side up, to the underside of the skin with the pencil outline of the patch matching the edges of the hole. If the curvature of the surface to be repaired is greater than a rise of 1/8 inch in 6 inches, the doubler should be preformed by hot water or steam bending to the approximate curvature. As an alternative to preforming of the 1/4 inch stock, the doubler may be laminated from two thicknesses of 1/8 inch ply.

Apply nailing strips outlining the hole to apply bonding pressure between doubler and skin. Use a bucking bar to provide support for nailing. When two rows of nails are used, stagger nail spacing. Allow adhesive to cure.

Apply adhesive to remaining surface of the doubler and to the mating surface on the patch. Lay the patch in position over the doubler, and screw the pressure plate to the patch assembly using a small nail to line up the holes that have been previously made with patch and plate matching. No. 4 round head screws are used. Lead holes in the plywood doubler are not necessary. Waxed paper or plastic wrap between the plate and patch prevents adhesive from bonding the plate to the patch. No clamps or further pressure need be applied, as the nailing strips and screws exert ample pressure.

Round plug patches may be made by following the steps in figure 1-18. The steps are identical to those for making the oval patch except for the insertion of the doubler. In using the round patch, where access is from only one side, the round doubler cannot be inserted unless it has been split.

FABRIC PATCH. Small holes not exceeding 1 inch in diameter, after being trimmed to a smooth outline, may be repaired by doping a fabric patch on the outside of the plywood skin. The edges of the trimmed hole should first be sealed, and the fabric patch should overlap the plywood skin by at least 1 inch. Holes nearer than 1 inch to any frame member, or in the leading edge or frontal area of the fuselage, should not be repaired with fabric patches.

### Page 96 (OCR)

**OUTER EDGE OF DOUBLER**

**BUTT JOINT OF PATCH TO SKIN**

**INNER EDGE OF DOUBLER**

**NAIL HOLES -**

**D SCREW HOLES -**

**TO BE FILLED BEFORE**

**FINISHING**

**PATCH DIMENSIONS**

A B C D E F SMALL 1-1/2 2-1/8 |1-1/2 718 3 4-1/2 LARGE 2 3-3/8 2-1/2 1-5/8 7

**TWO ROWS OF SCREWS AND**

**NAILS REQUIRED ON LARGE BUTT JOINT OF PATCH TO SKIN PATCH.**

PLYWOOD DOUBLER (GRAIN PARALLEL TO SKIN) >E- (LAMINATE DOUBLER FROM TWO МАШЕТ "14" PIECES OF 1/8" PLY IN AREAS OF

- PLUG PATCH (GRAIN PARALLEL TO SKIN) SKIN CURVATURE,)

PLYWOOD SKIN FIGURE 1-17. Oval plug patch assembly.

Par 1-54 Page 1-32

### Page 97 (OCR)

**GRAIN DIRECTION OF SKIN, PATCH, AND DOUBLER.**

- SAW CUT IN DOUBLER.

**BUTT JOINT OF PATCH TO SKIN.**

**INNER EDGE OF DOUBLER.**

• OUTER EDGE OF DOUBLER.

**NAIL HOLES -**

**SCREW HOLES - TO BE FILLED BEFORE FINISHING.**

**BUTT JOINT OF PATCH TO SKIN - PLUG PATCH**

PLYWOOD SKIN - SAW CUT IN DOUBLER PLYWOOD DOUBLER (LAMINATE DOUBLER FROM TWO PIECES OF 1/8" PLY IN AREAS OF SKIN CURVATURE.)

DIMENSIONS B C SMALL CIRCULAR PLUG PATCH 2 6/8 1 3/8 LARGE CIRCULAR PLUG PATCH 3 718 2 1/8 (TWO ROWS OF SCREWS AND NAILS REQUIRED

FOR LARGE PATCH) FIGURE 1-18. Round plug patch assembly.

1-53.-1-63. [RESERVED.] Par 1-52 Page 1-33 (and 1-34)

### Page 98

Par 1-64 Page 1-34

**SECTION 5. FINISHING WOOD STRUCTURES**

GENERAL. Any repair to spars, ribs, skin surfaces, or other structural parts of the airframe involves finishing as the final step of the job. The surface finish is the final line of defense to prevent the destructive effects of moisture entry into the structure. The time and effort spent during the preparatory phase of the refinishing process will be reflected in the appearance and longevity of the finished surface. Adherence to the instructions issued by the finish manufacturer is necessary to obtain the appearance desired and protective characteristics for the product used. The primary objective of interior finishes is to afford protection of the wood against serious change in moisture content when exposed to damp air or to water that gains entrance to closed spaces by condensation or by penetration of rain, mist, or fog. Coatings, on contact areas between wood and metal protect the metal against corrosion from moisture in the wood.

The primary objectives of the exterior finish are to protect the wood against weathering, provide a suitable appearance, and present a smooth surface in flight.

ACCEPTABLE FINISHES. Any varnish conforming to Federal Specification TT-V-109, as amended, or other coating approved by the airframe manufacturer or the FAA is acceptable. Exterior surfaces must be further protected from the effects of abrasion, weather, and sunlight. A number of systems for exterior finishing have STC approval and are manufactured under a PMA. (See CChapter 2, Fabric Covering.) Low viscosity epoxy adhesive (meeting the requirements of paragraph 1-4 for that purpose) may be used as an internal surface finish when subsequent bonding is necessary.

**PRECAUTIONS.**

When making repairs, avoid excessive contamination of surfaces with adhesive squeeze-out at joints and on all surfaces. Excess adhesive should always be removed before appl ying finish. Because many paints and adhesives are incompatible, even a slight amount of adhesive underneath the finish may cause premature deterioration of the finish.

Soiling substances, such as oil and grease, should be removed as completely as possible. Naphtha may be used to remove surface deposits of oil and grease;

however, thinned residue may penetrate into any unprotected wood. In areas where minor amounts of oil or grease have penetrated the wood surface, removal may be accomplished by use of an absorbent type of cleaner such as gunsmith's whiting or a clothing spot lifter such as K2r . Marks that are made by grease pencils or lumber crayons containing wax are harmful and should be removed, but marks made by ordinary soft graphite pencils and nonoily stamp pad inks may be safely finished over. All dust, dirt, and other solid particles should be removed.

Sawdust, shavings, and chips must be removed from enclosed spaces before they are sealed off by replacement of skin. A vacuum cleaner is useful for such cleaning.

Since most adhesives will not bond satisfactorily to sealers, it is necessary to avoid applying sealer over the areas where adhesive will be applied. Mark off areas to receive adhesive, and allow an additional 1/4 inch on each side of the adhesive area to provide for

### Page 99

Par 1-69 1-35 misalignment when mating the parts. It is preferable to leave some unsealed areas ra ther than risk weakening the joint by accidental overlap of the sealer into the bonded areas. Wherever possible, apply sealer to the margins after the adhesive has cured. As an example, the lower skin of a wing bay would be installed first, leaving access from above to apply sealer. All low spots (where moisture would collect) are well sealed.

The top skin would be installed last, so that the only unsealed margins would be on upper surfaces where moisture is least likely to collect.

An alternative to the previous paragraph is to use an approved epoxy coating and compatible epoxy adhesive. Apply the coating, allowing 1/4 inch margins as in the previous paragraph. After the coating has cured, apply epoxy adhesi ve to joint surfaces, and overlap the adhesive onto the sealer. Close joint and clamp. The epoxy adhesive will bond satisfactorily to the coating and ensure a complete coverage of the wood surfaces. Use only approved and compatible adhesives and coatings for this method.

FINISHING OF INTERIOR SURFACES. Finish repaired ribs, spars, interior of plywood skin, and other internal members, including areas of contact between metal and wood, by applying one thinned coat (for penetration into wood grain) of varnish or other acceptable finish, followed by two full coats. Protect built-up box spars and similar closed structures on the interior in the same way. Where better protection is required, as on the surfaces of wheel wells and the bottoms of hulls below the floor boards, an additional coat of aluminized sealer consisting of 12 to 16 ounces of aluminum paste per gallon of sealer, may be applied.

FINISHING OF EXTERIOR SURFACES. Exterior surfaces require more protection than interior areas due to the effects of abrasion, weather, and sunlight. (See chapter 2.) Tests have shown that the interior temperature of wooden aircraft structures can reach 185 ºF when the aircraft is finished in a dark color and parked outdoors on a hot, still day. Exposure to prolonged high temperature is detrimental to wood, adhesives, and finishes. Wood loses approximately 25 percent of its strength at 125 ºF. For this reason, the mechanic should consider temperature effects when selecting finish colors or looking for areas of likely deterioration. The lowest temperatures are found when the aircraft is finished in white or very light colors, while darker colors produce higher temperatures. A general trend toward higher temperatures may be seen when exterior colors are yellow, pink, light blue, aluminum, purple, blue, light green, orange, tan, red, green, brown, and black. A lighter shade of a particular color helps to reduce temperatures.

FINISHING OF END GRAIN SURFACES. End grain portions of wooden members are much more absorbent than side grain. Because of this extreme vulnerability to moisture entry, it is necessary to take extra precautions to seal end grain.

Apply at least one thinned coat of acceptable sealer to ensure maximum penetration, and then follow with as many full strength coats as necessary to achieve a smooth, glossy coating. Depending on the type wood to be finished, two to four full coats will be required. A final coat of aluminized varnish may be applied to end grain surfaces. If the surfaces are to be finished with dope or lacquer, a dope-proof sealer, similar to Federal Specification TT-V-109, or epoxy sealer should be used.

Exposed end grain includes such surfaces as spar butts, skin edges, areas around vent holes, inspection holes, fittings, and exposed scarfed or tapered surfaces.

### Page 100

Par 1-70 Page 1-36 (and 1-38) FINISHING WITH FABRIC OR TAPE. A number of systems for exterior finishing have STC approval and are manufactured under a PMA.

Follow the product manufacturer's instructions for the system used.

If the finish surrounding the repair is a traditional dope system, seal the wood grain with a suitable solvent resistant one-part varnish, commonly described as “dope proof,” or a two-part epoxy varnish. Follow with two coats of clear dope, and allow sufficient drying time between coats.

Apply a third coat of clear dope and lay a piece of pinked-edge airplane cloth into the wet film. All air bubbles should be worked out by brushing to ensure maximum adhesion. When dry, apply one brush coat, to ensure proper penetration, and at least one spray coat of clear dope. The dried spray coat may be sanded with fine sandpaper to obtain a smoother finish. Complete the refinishing of the surface by application of a topcoat as required to match the adjacent area.

SEALING OF BOLT HOLES. Bolt holes in wooden structure provide a vulnerable entry point for moisture. Variations in moisture content around bolt holes can lead to decay or splitting. In addition, excessive moisture at bolt holes promotes corrosion of the bolts.

Sealing of the wood surfaces in bolt holes can be accomplished by application of varnish or other acceptable sealer into the open hole. The sealer must be allowed to dry or cure thoroughly prior to bolt installation.

1-72.1-79. [RESERVED.]

### Page 101

Par 2-1 Page 2-37

**CHAPTER 2. FABRIC COVERING**

**SECTION 1. PRACTICES AND PRECAUTIONS**

GENERAL. Cotton and Irish linen fabrics were the airframe coverings of choice from WWI through the 1950’s. However, increases in cost and the short lifespan of natural fabrics became the driving factors which resulted in almost 100 percent replacement of original airframe fabrics by man-made, STC-approved, polyester, and glass filament fabric.

**PROBLEM AREAS.**

Deterioration. Polyester fabric deteriorates only by exposure to ultraviolet radiation as used in an aircraft covering environment. When coatings completely protect the fabric its service life is infinite. Therefore, it is very important to thoroughly protect the structure from deterioration before covering and provide adequate inspection access to all areas of fabric-covered components to allow inspection for corrosion, wood rot, and mice infestation. Multiple drain holes in the lower ends of all fabric-covered s ections also provide needed ventilation to remove condensation.

Tension. Polyester fabric obtains maximum tension on an airframe at 350 ºF, and will not be excessive on aircraft originally covered with natural fabric and 12 coats of Nitrate or Butyrate Dope. However, dope applied over full heat-tauted fabric can develop excess tension after aging and damage light aircraft structures.

Coatings other than dope will not increase fabric tension after aging. The heat-tauting instructions given in the manual of each STC-approved covering process should be followed.

**AIRCRAFT FABRIC-SYNTHETIC.**

STC-Approved Covering Materials. There is a wide selection of STC-approved covering materials available which utilize synthetic fabric falling within the generic class “Polyester” and may vary in characteristics.

Difference in the fabric may be denier, tenacity, thread count, weight, shrink, tension, and weave style.

Polyester Filaments. Polyester Filaments are manufactured by polymerization of various select acids and alcohols, then extruding the resulting molten polymers through spinnerets to form filaments. The filaments are heat stretched to reduce to the desired denier or size. It is the heat stretching that imparts a memory in the filaments causing them to try and return to their original shorter length when reheated at a controlled temperature. Overheating will cancel the memory and melt the filaments.

Covering Procedures. Coating types, covering accessories, and covering pro cedures also may vary;

therefore, the covering procedures given in the pertinent manuals must be followed to comply with the STC.

The FAA STC-approved installation takes precedence over instructions in this advisory circular.

Installation. Initial installation of polyester fabric is similar to natural fabric. The fabric is installed with as little slack as possible, considering fittings and other protrusions. It may be sewn into an envelope, installed as a blanket, or installed by cementing to the airframe with a fabric cement. E ach STC may differ in the cement seam overlap, type of sewn seam, heat shrinking procedures, and temperature.

### Page 102

Par 2-1 Page 2-38 AIRCRAFT FABRIC-NATURAL. Physical specifications and minimum strength requirements for natural fiber fabric, cotton and linen, used to recover or repair components of an aircraft, are listed in table 2-1. Tear resistance is an important factor when considering aircraft fabric. A test method such as ASTM D 1424 is recommended. Technical Standard Order TSO-C15d, entitled Aircraft Fabric, Grade A (AMS 3806D); and TSO-C14b, Aircraft Fabric, Intermediate Grade (AMS 3804C) current edition, respectively, describe the minimum standards that all fabric must meet to qualify as aircraft covering material.

RECOVERING AIRCRAFT. Recover or repair aircraft with a fabric of equal quality and strength to that used by the original aircraft manufacturer. It is recommended that fabric conforming to TSO-C15d or TSO-C14b be used to recover aircraft originally covered with lower strength fabric conforming to AMS 3802, current edition.

NOTE: Recovering or repairing aircraft with any type fabric and/or coating other than the type used by the original aircraft manufacturer is considered a major alteration. Obtain approval from the FAA on fabric and installation data. Cotton and linen rib lacing cord, mach ine and hand-sewing thread, and finishing tapes should not be used with polyester and glass fabric covering.

Reinforcing tape minimum tensile strength is listed in table 2-2. Reinforcing tape meeting specification MIL-T-5661, Type I, current edition, is acceptable.

Reinforcing tape should have a minimum 40 lb.

resistance without failure when static tested in shear against a single rib lace, or a pull-through resistance when tested against a single-wi re clip, rivet, screw, or any other type of fabric-to-rib attachment. Reinforcing tape is used over the rib cab on top of the fabric and for inter-rib bracing.

Finishing Tape, sometimes referred to as surface tape, should have the same properties as the fabric used to cover the aircraft.

Lacing Cord shall have a minimum breaking strength of 40 lb. Lacing cord meeting the specifications listed in table 2-2 is acceptable. Rib lace cord should have a microcrystalline fungicidal wa x, paraffin-free wax, or beeswax coating, or other approved treatment to prevent wearing and fraying when pulling through the structure.

Machine Thread shall have a minimum breaking strength of 5 lb. Thread meeting the specifications listed in table 2-2 is acceptable.

Hand-Sewing Thread shall have a minimum breaking strength of 14 lb. Thread meeting the specifications listed in table 2-2, is acceptable. When covering with STC-approved fabric covering material, use the type of sewing thread approved by the STC and manufactured under the specific PMA.

Flutter Precautions. When re-covering or repairing control surfaces, especially on high performance airplanes, make sure that dynamic and static balances are not adversely affected. Weight distribution and mass balance must be considered to preclude to possibility of induced flutter.

PREPARATION OF THE STRUCTURE FOR COVERING. One of the most important items when covering aircraft is the proper preparation of the structure. Before covering, the airframe must be inspected and approved by a FAA-certified mechanic or repair station.

### Page 103

Par 2-6 Page 2-39 TABLE 2-1. Cotton and linen fabrics.

Materials

Specification Minimum Tensile Strength New (undoped) Minimum Tearing Strength New (undoped) (ASTM D 1424) Minimum Tensile Strength Deteriorated (undoped)

Thread Count Per Inch

Use and Remarks

Airplane cloth mercerized cotton (Grade “A”).

TSO-C15d, as amended, references Society Automotive Engineers AMS 3806d, as amended or MIL-C-5646

80 pounds per inch warp and fill.

5 pounds warp and fill.

56 pounds per inch.

80 min., 84 max.

warp and fill.

For use on all aircraft.

Required on aircraft with wing loading of 9 p.s.f. or greater or placard never exceed speed of 160 m.p.h.

or greater.

Airplane cloth mercerized cotton.

TSO-C14b, as amended, references Society Automotive Engineers AMS-3804c, as amended.

65 pounds per inch warp and fill.

4 pounds warp and fill.

46 pounds per inch.

80 min., 94 max.

warp and fill.

For use on aircraft with wing loading less than 9 p.s.f. and never exceed speed of less than 160 m.p.h.

Airplane cloth mercerized cotton.

Society Automotive Engineers AMS 3802, as amended.

50 pounds per inch warp and fill.

3 pounds warp and fill.

35 pounds per inch.

110 max.

warp and fill.

For use on gliders with wing loading of 8 p.s.f. or less, provided the placarded never-exceed speed is 135

m.p.h. or less.

Aircraft linen.

British 7F1.

This material meets the minimum strength requirements of TSO-C15.

Battery Box Treatment. An asphaltic, rubber-based acid-proof coating should be applied to the structure in the area of a battery box, by brush, for additional protection from battery acid. Control cables routed in the area of the battery box should be coated with paralketone.

Worn Holes. Oversized screw holes or worn size 4 self-tapping screw holes through ribs and other structures used to attach fabric may be redrilled a minimum 1-1/2 hole diameter distance from the original hole location with a # 44 (0.086) drill bit. Size 6 screws, drill bit size # 36 (0.1065), may be installed in stripped or worn holes drilled for size 4 sc rews, usually without redrilling. Worn holes for wire clips and wire barbs should be redrilled a minimu m 1-1/2 hole distance from the original locations using a drill jig to ensure correct spacing, with the appropriate size drill bit. Drill bit size # 30 (0.128) may be used to redrill oversize holes for 1/8-inch diameter blind rivets a minimum 1-1/2 hole diameter distance from the original location.

### Page 104

Par 2- Page 2-40 TABLE 2-2. Cotton and Linen, Tapes and Threads.

Materials

Specification

Yarn Size Minimum Tensile Strength Yards Per Pound

Use and Remarks

Reinforcing tape, cotton.

MIL-T-566 1 E, Type 1 MIL-Y-1140H

150 pounds per 1/2 inch width.

Used as reinforcing tape on fabric and under rib lacing cord. Strength of other widths approx. in proportion.

Lacing cord, prewaxed braided cotton.

Federal T-C-57 1F

40 pounds.

310 minimum.

Lacing fabric to structures. Unless already waxed, must be lightly waxed before using.

Lacing cord, braided cotton.

**MIL-C-5648A**

80 pounds.

170 minimum.

Lacing fabric to structures. Unless already waxed, must be lightly waxed before using.

Lacing cord thread, high tenacity cotton.

**MIL-T-5660B**

Ticket No.

10.

62 pounds.

480 minimum.

Lacing fabric to structures. Unless already waxed, must be lightly waxed before using.

Machine thread cotton Federal V-T- 276H

20/4 ply

5 pounds.

5,000 nominal.

Use for all machine sewing.

Hand-Sewing thread cotton.

Federal V-T- 276H Type III B

8/4 ply

14 pounds.

1,650 nominal.

Use for all hand-sewing.

Use fully waxed thread.

Finishing (Surface) tape cotton.

Same as fabric used.

Same as fabric used.

Use over seams, leading edges, trailing edges, outer edges and ribs, pinked, raveled or straight edges.

Fairing Precautions. Aluminum leading edge replacement fairings installed in short sections may telescope during normal spar bending loads or from thermal expansion and contraction. This action may cause a wrinkle to form in th e fabric, at the edge of the lap joint. Leading edge fairing sections may be fastened together with rivets or screws to prevent telescoping after installation. Trailing edges should be adequately secured to prevent movement and wrinkles.

Dope Protection . Solvents found in nitrate and butyrate dope will penetrate, wrinkle, lift, or dissolve most one-part wood varnishes and one-part metal primers. All wood surfaces th at come in contact with doped fabric should be treated with a protective coating such as aluminum foil, cellulose tape, or dope-proof paint to protect them against the action of the solvents in the dope. This can also be accomplished by recoating with a suitable, solvent resistant two-part epoxy varnish, which will be impervious to solvent penetration and damage after curing. Clad aluminum and stainless steel parts need not be dope-proofed.

A solvent-sensitive primer on ferrous metal and aluminum alloy components which will be in contact with fabric may be protected from solvent damage by overcoating with a two-part epoxy primer. Epoxy primer meeting MIL-P-53022B is acceptable.

Small metal or wood surfaces, such as rib caps, to which fabric will not be dope bonded as a part of the particular fabric attachment procedure may be protected from dope damage by cellophane tape or aluminum foil.

Chafe Protection. Fabric and finishing tape is often cut through with sandpaper over sharp edges during the coating and finishing procedure and later polishing. All sharp metal edges or protruding screws, nails, rivets, and bolt heads should be covered with an anti-chafe tape to prevent cutting and wearing through the fabric after installation. Use

### Page 105

Par 2-6 Page 2-41 appropriate non-bleeding cotton adhesive coated tape, finishing tape, or strips of fabric, cut from the fabric being used to cover the aircraft, doped in place.

Small holes cut through the fabric to accommodate flying wires, control cables, and fittings, must be reinforced with finishing tape or fabric patches cut from the same fabric used for the covering.

Areas needing additional chafe protection such as control cables routed firm against the fabric surface should be protected with patches cut from cotton duck, leather, or plastic. These patches may be sewn, doped, or cemented in place, as appropriate.

Any drag and anti-drag wires in the wings should be protected from chafing at cross points.

Inter-Rib Bracing. Use a woven fabric tape of the same quality and width as th at used for the rib lace reinforcing, where so incorporated in the wing design by the original aircraft manufacturer. When the original routing for the inter rib bracing is not known, the tape will be routed diagonally, a lternating between the top and bottom of each rib cap on each successive rib, if a single pair, half way between the front and rear spars.

The number of tape pairs will duplicate the original aircraft manufacturer’s installation. Tapes will be routed continuously from the wing butt to the wingtip bow, with one turn of tape around each intermediate rib cap strip. Care should be given to position the tape so as not to interfere with control cables, bellcranks or push-pull rods.

Preparation of Plywood Surfaces for Covering. Prior to covering plywood surfaces, prepare the surface by sanding, cleaning, and applying sealer and dope. When plywood surfaces are to be c overed with light weight glass fiber deck cloth instead of fabric, no sealer or dope should be applied to the plywood as it would inhibit penetration of epoxy resin.

Sand plywood surfaces as needed to remove old loose dope or varnish residue to provide a clean bonding surface. Remove any oil, grease, or other contamination with a suitable solvent such as naphtha. Small, rough areas and irregularities in the plywood surface and around any plywood repairs may be filled and smoothed with an appropriate commercial grade wood filler. Filling large warp depressions on plywood surfaces with a wood filler for cosmetic purposes is not acceptable.

After cleaning and sanding all plywood surfaces, seal the wood grain with a suitable solvent resistant two-part epoxy varnish. After the varnish has thoroughly dried, apply two brush or spray coats of clear dope, allowing sufficient drying time between coats.

FABRIC SEAMS. Seams parallel to the line of flight are preferable; however, spanwise seams are acceptable.

Sewn Seams.

Machine-sewn seams should be double stitched using any of the styles illustrated in figure 2-1 A, B, C, or D. A machine-sewn seam used to close an envelope at a wingtip, wing trailing edge, empennage and control surface trailing edge, and a fuselage longeron may be made with a single stitch when the seam will be positioned over a structure. (See figure 2-1 E.) The envelope size should accommodate fittings or other small protrusions with minimum excess for installation. Thick or protruding leading edge sewn seams should be avoided on thin airfoils with a sharp leading edge radius because they may act as a stall strip.

### Page 106

Par 2-6 Page 2-42 Hand sew, with plain overthrow or baseball stitches at a minimum of four stitches per inch, or permanent tacking, to the point where uncut fabric or a machine-sewn seam is reached. Lock hand sewing at a maximum of 10 stitch intervals with a double half hitch, and tie off the end stitch with a double half hitch. At the point where the hand-sewing or permanent tacking is necessary, cut the fabric so that it can be doubled under a minimum of 3/8 inch before sewing or permanent tacking is performed. (See figure 2-2.)

After hand sewing is complete, any temporary tacks used to secure the fabric over wood structures may be removed.

Cover a sewn spanwise seam on a wing’s leading edge with a minimum 4-inch wide pinked-edged surface tape with the tape centered on the seam.

Cover a spanwise-sewn seam at the wing trailing edge with pinked-edge surface tape that is at least 3 inches wide. For aircraft with never-exceed speeds in excess of 200 mph, cut V notches at least 1 inch in depth and 1/4 inch in width in both edges of the surface tape when used to cover spanwise seams on trailing edges of control surfaces. Space notches at intervals not exceeding 6 inches.

On tape less than 3 inches wide, the notches should be 1/3 the tape width. In the event the surface tape begins to separate because of poor adhesion or other causes, the tape will tear at a notched section, thus preventing progressive loosening of the entire length of the tape which could seriously affect the controllability of the aircraft. A loose tape acts as a trim tab only on a movable surface. It becomes a spoiler on a fixed surface and has no effect at the trailing edge other than drag.

Make spanwise-sewn seams on the wing’s upper or lower surfaces in a manner that will minimize any protrusions. Cover the seams with finishing tape at least 3 inches wide, centering the tape on the seam.

Sewn seams parallel to the line of flight (chordwise) may be located over ribs.

However, careful attention must be given to avoid damage to the seam threads by rib lace needles, screws, rivets, or wire clips that are used to attach the fabric to the rib. Cover chordwise seams with a finishing tape at least 3 inches wide with the tape centered on the seam.

Doped Seams.

For an overlapped and doped spanwise seam on a wing’s leading edge, overlap the fabric at least 4 inches and cover with finishing tape at least 4 inches wide, with the tape centered at the outside edge of the overlap seam.

For an overlapped and doped spanwise seam at the trailing edge, lap the fabric at least 3 inches and cover with pinked-edge surface tape at least 4 inches wide, with the tape centered on the outside edge of the overlap seam.

For an overlapped and doped seam on wingtips, wing butts, perimeters of wing control surfaces, perimeters of empennage surfaces, and all fuselage areas, overlap the fabric 2 inches and cover with a finishing tape that is at least 3 inches wide, centered on the outside edge of the overlap seam.

For an overlapped and doped seam on a wing’s leading edge, on aircraft with a velocity never exceed (Vne) speed up to and including 150 mph, overlap the fabric 2 inches and cover with a finishing tape that is at least 3 inches wide, with the tape centered on the outside edge of the overlap seam.

### Page 107 (OCR)

1/16" 3/8* - 1/16" 116→ 1114"→ 1/16"

**STITCHING**

•STITCHING A. PLAIN OVERLAP SEAM, STRAIGHT EDGE B. FRENCH FELL SEAM 1/16" 3/8" - 1/16" 1/4"￾

**STITCHING C. FOLDED FELL SEAM**

STITCHING STEP 1 1/16"-

•3/8** 1/16" STEP 2 STITCHING E. SINGLE STITCH SEAM TO CLOSE AN

**ENVELOPE WHEN THE STITCH WILL**

**BE LOCATED OVER A STRUCTURE D. MODIFIED FOLDED FELL SEAM SEWN IN**

TWO STEPS FIGURE 2-1. Fabric seams.

Par 2-7 Page 2-43

### Page 108 (OCR)

**USE A T-PIN THROUGH THE FABRIC**

**STRAPS TO TEMPORARILY SECURE**

**THE FABRIC COVERING AND REMOVE AS HAND SEWN SEAM PROGRESSES MINIMUM 4 STITCHES PER INCH.**

**TOP COVER**

- FORMER (A) ATTACHING FABRIC AT AILERON CUTOUT

**TOP COVER / TOP COVER**

FINISH WITH TAPE SPACING 1-1/4"'

**OR CLOSER**

TEMPORARY TACKS MAY BE USED TO SECURE FABRIC IN POSITION FOR HAND- SEWN SEAMS OVER A WOOD STRUCTURE. PERMANENT TACKS MAY BE USED

**TO SECURE FABRIC AT WING BUTT STRUCTURES IN LIEU OF A HAND-SEWN**

SEAM AND SHOULD BE OF THE APPROPRIATE LENGTH AND GAUGE, AND MADE OF CORROSION RESISTANT STEEL OR BRASS.

FIGURE 2-2. Typical methods of attaching fabric.

Par 2-7 Page 2-44

### Page 109

Par 2-7 Page 2-45 For an overlapped and doped seam on the perimeter of a wing (except a leading edge), perimeters of wing control surfaces, perimeters of empennage surfaces, and all areas of a fuselage, on aircraft with a Vne speed up to and including 150 mph, overlap the fabric 1 inch and cover with a finishing tape that is at least 3 inches wide, centered on the outside edge of the overlap seam.

COVERING METHODS. The method of fabric attachment should be identical, as far as strength and reliability are concerned, to the method used by the manufacturer of the airplane being recovered or repaired. Carefully remove the old fabric from the airframe, noting the location of inspection covers, drain grommets, and method of attachment. Cotton or linen fabric may be applied so that either the warp or fill-threads are parallel to the line of flight. Either the envelope method or blanket method of covering is acceptable.

The Envelope Method. A wing envelope may be developed by two methods. Machine sew together, side by side multiple fabric sections, cut to reach chordwise around the wing, starting a nd ending at the trailing edge with a minimum of 1 inch excess length. The sewn envelope is then positioned around the wing and secured with closely spaced T- Head pins at the wingtip and trailing edge. Excess material may then be trimmed. Carefully remove the envelope and complete by machine sewing at the wi ngtip and along the trailing edge, except where the geometry of the wing (aileron and flap cut out) would prevent the sewn envelope from being reinstalled. After reinstalling the envelope, the un-sewn sections and butt end are then closed by handsewn or overlapped and doped seams in accordance with the aircraft Vne speed. (Refer to paragraph 2-7 b.)

An alternative method, when fabric of sufficient width is available, is to sew together, side-by-side, two sections of fabric, placing the seams spanwise on the leading edge, then fit and sew the wingtip and trailing edge in the same manner as the multiple piece chordwise envelope.

An envelope may be developed for the fuselage in the same manner, with a final closing along a longeron by hand-sewn or overlapped and doped seams in accordance with the aircraft Vne speed.

The Blanket Method. A blanket is developed by sewing together, side-by-side, multiple sections of fabric with the seams chordwise or two wide sections of fabric, side-by-side, placing the seam spanwise on the leading edge, the same as an envelope. Close the three remaining sides with a hand-sewn seam or overlapped and doped seams in accordance with the aircraft Vne speed. Small components may be covered by wrapping one piece of fabric over a st raight leading or trailing edge, then closing three sides with hand-stitched or overlapped and doped seams in accordance with the aircraft Vne speed.

NOTE: All overlapped and doped seams will be made only over underlying supporting structures extending the full width of the seam.

Machine-sewn alternate. An alternate to machinesewn seams on a wing envelope or blanket is to use two sections of wide fabric spanwise. Attach the fabric with overlapped and doped seams at the leading and trailing edge, wingtip and wing butt, in accordance with the aircraft Vne speeds. (Reference paragraph 2-7 b.)

Smaller components may be covered in the same manner. The fuselage may be covered with multiple fabric sections with

### Page 110

Par 2-7 Page 2-46 overlapped and doped seams on the longerons or other wide fabric-forming structur es in accordance with the aircraft Vne speed. (Reference paragraph 2-7 b.)

Holes in Fabric. Never cut any holes in the fabric for inspection panels, spar fittings, or drain grommets; or attach the fabric to the airframe with rib lacing screws, rivets, clips, or rib stitch cord until the fabric has been semi-tauted and stabilized with several coats of dope.

**REINFORCING TAPE.**

Reinforcing tape should be securely bonded to the fabric surface with dope before cord lacing or installation of hardware. Where multiple attachments are in close proximity, such as on a wing rib, continuous reinforcing tape should be installed, extending at least 1 inch past the last attachment at each end. Random or wide spacing, such as on fuselage stringers or empennage surfaces, may be reinforced with 2-inch lengths of reinforcing tape centered on the attachment location.

Reinforcing tapes should be of the appropriate width for hardware attachment such as screws, rivets, wire clips, etc., which pierce the center of the tape.

Reinforcing tape under cord lacing should be the same width as the rib to which the fabric is laced and may be comprised of multiple widths positioned side-by-side to achieve the required width.

When the aircraft Vne speed is over 250 mph, anti-tear strips, cut from the same quality fabric used to cover the aircraft, are recommended for use under reinforcing tape on the entire top surface of the wing and on the portion of the wing’s bottom surface in the propeller slipstream. The propeller slipstream is considered to be the propeller diameter plus one outboard rib. The anti-tear strip should be installed completely around the wing, beginning and ending at the trailing edge in the propeller slipstream, and installed from the trailing edge over the leading edge and back under to the front spar on the balance of the ribs. Anti-tear strips should extend 1/2 inch past the wing rib cap edges and be thoroughly bonded to the fabric with dope before the reinforcing tape is installed.

(See figure 2-3.)

FIGURE 2-3. Exploded side view of rib.

### Page 111

Par 2-10 Page 2-47 LACING.

Fabric should be attached to aircraft components to prevent ballooning due to aerodynamic forces, in the identical manner and locations as used by the original aircraft manufacturer. Any deviation from the original method(s) of attachment, such as screws, rivets, wire clips, lacing cord, etc., are considered a major alteration and in conflict with the aircraft type design data.

Obtain FAA approval on any deviation.

NOTE: When the type of rib lace knot used by the original aircraft manufacturer is not known, the modified seine knot shown in figure 2-4 through figure 2-9c will be used.

FIGURE 2-4. Standard external modified seine knot used for single and double rib lacing.

During the installation of lacing cord through a wing or any other component, special attention should be given to avoid interference with the routing of any control cable, bellcrank, or any other movable item. To prevent chafing and cutting of the lacing cord, control cables or any other movable items should be tensioned or positioned to their normal alignment before rib lacing and checked afterwards to ensure adequate clearance.

When a lace cord will be chafed by a moving component, a blindstitch may be made around the top and bottom rib caps as illustrated in figure 2-11.

Stationary structures interfering with needle routing may be circumvented by aligning the needle forward or aft adjacent to the rib cap. Pull the needle through the wing and then return through the same hole and exit at the desired adjacent location.

NOTE: The first lace on a wing rib should be spaced from the leading edge fairing no more than 1/2 the required lace spacing for the balance of the rib.

Both surfaces of fabric covering on wings and control surfaces must be securely fast ened to the ribs by lacing cord or any other method originally approved for the aircraft. Care must be taken to insure that all sharp edges against which the lacing cord may bear are protected by tape in order to prevent abrasion of the cord. Separate lengths of lacing cord may be joined by the splice knot shown in figure 2-10 or tied off. The first loop is tied with a square knot as illustrated in figure 2-5 and figure 2-9a, and the knot secured with a half hitch on each side afte r the lacing is pulled tight around the rib. The needle is then routed through the wing and around the rib cap at the

### Page 112 (OCR)

**SECOND STITCH OF STANDARD LACING (SINGLE (SHOWN) OR DOUBLE)**

**HALF HITCH AROUND STITCH LOOPS C**

**WITH LEAD A BEFORE PROCEEDING TO**

**SECOND STITCH**

**LOCK SQUARE KNOT WITH HALF HITCH LEAD A AROUND STITCH LOOPS C WITH END B.**

**CUT OFF SURPLUS OF END B.**

**AFT -FINISHED FIRST KNOT.**

**CAPSTRIP 2- TIE SQUARE KNOT**

**ON CAPSTRIP WITH LEAD A AND END B.**

REINFORCING TAPE (same width as the rib).

• LEAD A END B

**STITCH LOOPS C**

**CAPSTRIP START OF**

FIRST KNOT S = NORMAL STITCH SPACING

**SCHEMATIC REPRESENTS LACING AROUND ENTIRE RIB.**

\ TIE SQUARE KNOT ON FIRST STEP OF S/2 CENTER OF CAPSTRIP

**STARTING STITCH**

LEADING EDGE FAIRING FIGURE 2-5. Starting stitch for rib lacing.

Par 2-10 Page 2-48

### Page 113 (OCR)

Modified Seine Knot , Single Loop lacing Aft Reinforcing Tape (should be same width as the rib) Capstrip Capstrip S Reinforcing tape (Same width as Rib Cap) Schematic represents lacing around entire rib.

Starting stitch S = Normal stitch spacing Wing Leading Edge Fairing FIGURE 2-6. Standard single-loop lacing.

Par 2-10 Page 2-49

### Page 114 (OCR)

- Pull No. 4

COMPLETED KNOT Edge of capstrip Pul No. 4 - 10 lb.

to complete knot Edge of capstrip Pull No. 3 pull snug OPERATION NO. 2

- Direction of lacing Pull No. 2 - 5 1b.

to tighten stitch Stitch loops around capstrip Pull No. 1 - 10 lb. From preceding to tighten loops stitch OPERATION NO. 1 FIGURE 2-7. Standard knot for double-loop lacing.

Par 2-10 Page 2-50

### Page 115 (OCR)

- Moditied Seine Knot Double loop lacing

• Reinforcing Tape (Same width as rib)

- Capstrip Capstnp - S Reinforcing tape (Same width as Rib Cap) Schematic represents lacing around entire nib.

Starting stitch S = NORMAL STITCH SPACING Wing Leading Edge Fairing FIGURE 2-8. Standard double-loop lacing (optional).

Par 2-10 Page 2-51

### Page 116 (OCR)

Square Knot Half Hitch Step 1. Tie a square knot by passing the short Step 2. Secure the tight square knot with a end of the cord thru the fold-back loop, half hitch at each side.

as illustrated.

Step 3. Route the needle back thru the right hand hole and exit at the next pre-punched lacing location.

Step 4. Route the needle back thru the exit hole and thru the opposite fabric surface leaving approximately a 3" loop around a finger on the top surface.

FIGURE 2-9a. Alternate sequence to tie a modified seine knot for rib lacing.

Par 2-10 Page 2-52

### Page 117 (OCR)

Step 5. As the needle is returned thru the top surface on the Step 6. The needle tip is routed under cord section "B" to opposite side of the rib cap the loop is rotated to hook and pull cord section "A" aft.

position cord section "A" to the forward side of the needle, then the needle is pulled thru.

Step 7. The needle tip is then routed over the top of cord Step 8. The needle tip is then routed over cord section "D" section "A" and under cord section "B." and passed thru the lacing while holding cord section "D" perpendicular to the surface to avoid cord entanglement.

FIGURE 2-9b. Alternate sequence to tie a modified seine knot for rib lacing.

Par 2-10 Page 2-53

### Page 118

Par 2-10 Page 2-54

Step 9. Pull cord section “D” perpendicular to the fabric surface to remove all slack in the cord back to the last rib lacing knot while working the loose knot to the right side. Do not pull cord section “E.”

Step 10. After all slack is removed by pulling cord section “D,” switch hands and place a thumbnail on the loose knot formed on the right-hand side, then secure the knot by pulling firmly perpendicular to the fabric surface on cord section “E.”

Step 11. After completing all lacing in the same sequence, the end is secured with a half hitch after the modified seine knot. The knot is pulled to the inside by routing the needle thru the wing before cutting the cord to leave the end inside.

FIGURE 2-9c. Alternate sequence to tie a modified seine knot for rib lacing.

### Page 119

Par 2-10 Page 2-55 next rib lace location with the cord and knot remaining on top of the fabric surface as illustrated in figure 2-5, figure 2-6, and figure 2-8. An alternate method is to route the needle under the fabric and out through the next lace location, then back down through the wing as illustrated in figure 2-9a through figure 2-9c. A modified seine knot is then tied as illustrated in figure 2-4 through figure 2-9c.

Rotate each lace loop to place the knot at the side of the rib cap to reduce the protrusion and aerodynamic interference before moving to the next lace location, or the cord routed under the fabric to the next lace location as illustrated in figure 2-9a through figure 2-9c.

The end cord is then cut off leaving a minimum of 1/4 inch stub. Lacing tension should be uniform.

Repeated pulling of long lengths of lacing cord may remove wax coating from the cord and cause fraying. Convenient lengths of rib lacing cord may be used to lace long or thick ribs. The end of each length is tied off with a half hitch as illustrated in figure 2-9c, or if needed, separate lengths of lacing cord may be joined by using the splice knot illustrated in figure 2-10.

Lacing is installed through other components, where applicable, in the same manner as a wing. Single, wide space lace attachments, usually used on empennage surfaces, are tied with a square knot and half hitch on each side, the same as a starting wing rib lace illustrated in figure 2-9a, steps 1 and 2. The lace may be rotated to place the knot under the fabric before cutting the cord.

Blind lacing on a fuselage, wing rib caps above and below a fuel tank, and any other component, when used by the original aircraft manufacturer, should be reinstalled in the same location and spacing as installed by the original aircraft manufacturer. The lace cord is routed around the stringer, rib cap, or other structure using an appropriate length, single or double pointed, curved needle as illustrated in figure 2-11. Blind laces are tied with a square knot, then pulled tight and secured with a half hitch at each side. The lace may be rotated to place the knot under the fabric surface before cutting the cord.

The splice knot is made by crossing the ends of the cord, and making four complete wraps with the small end of the free piece around the end of t he standing piece. The end is then doubled back through the formed loop. The other free end is wrapped and doubled back. The long ends of the cords are then pulled until the knot is tight. The short ends are cut close to the knot. This finishes the splice knot.

FIGURE 2-10. Splice knot.

**STITCH SPACING.**

Rib lace spacing on wings, formerly referred to as stitch spacing, should be no greater than the spacing used by the original aircra ft manufacturer. When the original spacing cannot be determined the maximum spacing illustrated in figure 2-12 should be used on the wings and wing control surfaces.

### Page 120 (OCR)

Route the needle adjacent to the rib cap at opposite sides.

Return the needle thru the exit hole and thru a third hole opposite the first entry hole.

Complete the square knot with Tie a square knot by a half hitch at each side. passing the short end of the cord thru the foldback loop,as illustrated.

Secure the tight square knot Route the needle back through the with a half hitch at each side. fabric and pull the knot inside before cutting the cord.

FIGURE 2-11. Blindstitch lacing - square knot secured with half hitches.

Par 2-10 Page 2-56

### Page 121

Par 2-11 Page 2-57

F IGURE 2-12. Fabric attachment spacing.

When the lace spacing on the empennage surfaces and fuselage, where so incorporat ed by the original aircraft manufacturer, cannot be determined, a maximum spacing equal to two times the spacing shown in figure 2-12 for the slipstream area on the wings may be used.

The installations of fabric attachments such as screws, rivets, wire clips, and rib lacing should be delayed until the fabric is stabilized and pulled taut with dope. This action is delaye d to avoid pulling wing ribs and other structures out of alignment or tearing the fabric at attachment points as the fabric becomes taut.

All lacing should be installed adjacent to the structure to which the fabric is being laced, to avoid tearing the fabric and/or creating slack in the cord loop when a load is applied. Where plastic washers were used by the aircraft manufacturer to provide increased pull-through resistance, under the heads of rivets or screws, the same diameter aluminum washer may be used as replacement. Aluminum washers are used because th ey are not affected by solvents found in adhesives or dopes, nor do they become brittle because of age or cold weather.

FASTENERS. Several light aircraft designs employ screws, rivets, or single-wire metal clips to secure the fabric to the wing.

Screws holding the old fabric can be removed after spinning a small sharpened tube around each screw or using a razor blade to cut a nd peel away the finishing tape. Care must be taken not to mark or scribe the underlying metal or wood structure. Blind rivets through ribs can be removed by drilling in the center to undercut the head.

Single-wire clips may be removed without damage to the rib by inserting a wide, thin screwdriver blade under the clip and carefully twisti ng. Apply a lifting force at the clip end to pull it up through the hole.

### Page 122

Par 2-14 Page 2-58 (and 2-24) NOTE: It is important that any damage found to ribs, such as oversize rivet or screw holes, and cracks or breaks in the rib cap, should be tagged immediately for easy location and repair later.

When repairs are made to fabric surfaces attached by special mechanical methods, duplicate the original type of fastener. When self-ta pping screws are used for the attachment of fabric to the rib structure, observe the following procedure:

Redrill the holes where necessary due to wear, distortion, etc., and in such cases, use a screw one size larger as a replacement.

Extend the length of the screw beyond the rib capstrip at least two threads.

Install a thin washer, preferably aluminum, under the heads of screws and dope pinkededge tape over each screw head.

**FINISHING TAPE.**

Finishing tape (surface tape) is installed after the fabric has been pulled taut with th e initial dope application.

This procedure is performed to prevent ripples from forming in fabric panels adjacent to newly applied tapes. Ripple formation is caused by the inability of the combined tape and fabric to tighten uniformly with adjacent fabric when additional dope is applied.

In addition to the tape widths required to be installed over fabric seams specified in paragraph 2-7, finishing tape should be installed as weather protection over all rib lacing, screws, rivets, wire clips, or other devices used to secure fabric. This includes wings, control surface ribs, empennage surface ribs, and fuselage stringers, where so installed by the original aircraft manufacturer. Tape width should be sufficient to bond the fabric a minimum of 3/8 inch on each side of all fabric attachments. Two inch width tape is normally used. Tapes over wing rib lacing should extend a minimu m of 1/2 inch past each end of any reinforcing tapes. Random or widely-spaced attachments may be covered by individual sections of fabric or finishing tape.

Installation of finishing tapes for additional wear resistance is recommended over the edges of all fabricforming structures. This includes fuselage stringers, longerons, leading and trailing edges, false or nose ribs, control surfaces, and empe nnage ribs not already covered and protected by a finishing tape that is required to be on a fabric seam or fabric attached to the structure. Compound surfaces, such as wingtip bow and empennage surfaces, are more conveniently taped using bias cut finishing tape, which easily conforms to the compound contour, rather than notching linear cut tape to fit the surface. Bias cut tape will be reduced to approximately two thirds the original cut width when pulled tight around a wingtip bow and should be considered when selecting the width of tape for the various locations.

Finishing tapes are applied by coating the fabric surface over which the tape will be applied with dope, applying the tape over the wet dope film, then brushing the tape firmly onto the fabric surface. This action will assure a good bond by thoroughly saturating and wetting the finishing tape.

**INSPECTION RINGS AND DRAIN GROMMETS.**

Inspection Rings. Inspection access is provided adjacent to or over every c ontrol bellcrank, drag-wire junction, cable guide, pulley, wing fitting, or any other component throughout the ai rcraft which will be inspected or serviced annually. They are installed only

### Page 123

Par 2-14 Page 2-59 (and 2-24) on the bottom side of the wi ngs except where installed on the top surface by the original manufacturer.

Cutting the holes may be delayed until needed; however, all covers should be finished in matching colors with any trim lines and stored until needed. Spraying matching colors a year later is expensive and time consuming.

The 3-9/16 inch inside diameter cellulose acetate butyrate (CAB) plastic inspection access rings have become popular and bond satisfactorily with Nitrate Dope or Fabric Cement. Any metal inspection hole reinforcements of a particular shape or special design or size, installed by the original manufacturer, should be reinstalled after cleaning.

Tapes or patches over aluminum reinforcements are optional, but recommended in the prop-wash areas on the wings and forward fuselage bottom.

Fabric patches over plastic rings are strongly recommended because plastic is not a stable material, becomes brittle at low temperatures, and fatigues and cracks from prop blast vibration. Plastic rings are often cracked during removal and installation of spring, clipheld covers. Patches with a minimum 1-inch overlap, should be installed with dope.

Drain Grommets. Atmospheric temperature changes cause the humidity in the air to condense on the inside of aircraft surfaces and pool in all low areas. Rainwater enters through openings in the sides and top, and when flying, everywhere throughout the structure. Taxiing on wet runways also splashes water up through any bottom holes. Therefore, provisions must be made to drain water from the lowest point in each fabric panel or plywood component throughout the airframe while in a stored attitude. Drain holes also provide needed ventilation.

Install drain grommets on the under side of all components, and the lowest point in each fabric panel when the aircraft is in stored attitude. Seaplane grommets, which feature a protruding lip to prevent water splashes through the drain hole, are recommended over drain holes subject to water splashing on land planes as well as seaplanes. The appropriate-size holes must be cut through the fabric before installing seaplane grommets. Plastic drain grommets may be doped directly to the fabric surface or mounted on fabric patches then doped to the covering. Installing a small fabric patch over flat grommets to ensure security is optional.

Alternate brass grommets are mounted on fabric patches, then doped to the fabric.

After all coating applications and sanding are completed, open all holes through flat drain grommets by cutting through the fabric with a small-blade knife. Do not attempt to open drain holes by punching with a sharp object because the drain hole will not remain open.

2-15.2-19. [RESERVED.]

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Par 2-20 Page 2-60

**SECTION 2. APPLICATION OF DOPE**

GENERAL. Nitrate dope and butyrate dope are manufactured by treating cellulose, derived from wood pulp or cotton linter with select acids, then dissolving in a blend of solvents and adding plasticizers for flexibility. After a brush or spray application on fabric, the film develops tension and strength as the solvents evaporate. The tension and strength will increase in proportion to the total film thickness. The fabric functions as a film former and carries no load until a crack develops in the dope film. An excessively-thick dope film will develop too much tension and may warp or damage a light airframe.

Viscosity adjustments for brush or spray applications may also vary between brands. When the viscosity adjustment ratio is not provided or is unclear, the product manufacturer should be contacted for detailed instructions. If instructions cannot be obtained or the source of the dope is unknown, the dope should be considered suspect and its use is not advised.

Dope, which has been stored for an extended period of time or under adverse conditi ons, should be suspected of becoming acidic and should be tested before being used on cotton or linen fabric. In some cases, fresh production dope has also been found to have a high acid content and will begin to deteriorate cotton or linen in a period of a few months. The acid content of nitrate or butyrate dope should not exceed 0.06 percent, calculated as acidic acid. An acidity test can be performed by most testing laboratories if high acid content is suspected.

Butyrate dope is superior to nitrate dope in weather exposure tests. However, nitrate dope provides better adhesion to natural fiber than butyrate dope. The adhesion of butyrate dope to natural fibers is adequate;

it is not necessary to use nitrate dope for the first application and butyrate dope for all other applications. The presence of naphtha in nitrate dope, manufactured in accordance with canceled M il Specs formulas, causes nitrate dope to be incompatible with butyrate dope;

therefore, nitrate dope should not be applied over butyrate dope for repairs or refinishing; however, butyrate dope may be applied over nitrate dope.

NOTE: Nitrate or butyrate dope thinners and retarders should not be substituted for each other, nor should automotive coating-type thinners be used.

Clear dope produces the most tension and strength.

Aluminum-pigmented dope w ill weigh slightly more than clear dope and develop less tension and strength for the same film thickness. Pigmented color finishes will produce the least tension and strength due to the higher ratio of plasticizers.

During the coating -buildup procedure, solvents released from each succeeding coat will penetrate and be absorbed into the previous dope film, temporarily releasing the tension and increasing the drying time between coats as the dope film becomes thicker. If elapsed time between coats exceeds several weeks at temperatures above 70 °F, it is recommended that several spray coats of an appropriate dope thinner or dope with retarder and/or re juvenator added be applied to the lightly-sanded, dried dope film to open the surface and provide cohesion fo r the next coat. This will reduce the possibility of surface cracks caused by dissimilar tension between the old and new dope film.

All dope coats through the final finish may be applied with a brush; however, brush

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Par 2-20 Page 2-61 marks will be noticeable in th e finish. With increasing environmental concerns, high pressure airless and highvolume low-pressure (HVLP) paint spray equipment is recommended over conventional siphon and pressure pot spraying equipment. High pressure airless and HVLP paint spraying equipment will greatly reduce paint over-spray and fogging. A spray gun, single coat is applied by overlapping each consecutive pass 50 percent of the fan width. A double coat is applied by repeating the coating application in the same direction, or at a 90 °angle to the first coat (cross coat) before the first coat has flashed off or dried dust free.

For safety and helpful tips for doping, see tables 2-3 and 2-4.

TABLE 2-3. Safety tips for dope/paint.

**SAFETY TIPS.**

Always ground the aircraft structure while sanding and painting.

Do not use an electric drill as a dope/paint mixer.

Wear leather-soled shoes in the dope/painting area.

Have an adequate, approved ventilation system.

Wear cotton clothes when doping or painting.

Wear an approved face mask or respirator when spraying.

Follow all the manufacturer’s instructions.

TABLE 2-4. Tips for doping.

**HELPFUL TIPS FOR DOPING.**

Limits for optimum application of dope: relative humidity 20 to 60%; temperature range 65° to 75 °F.

Drying time will vary with temperature, humidity, amount of thinner used, and whether or not retarder was added to the mixture.

Do not recoat until the surface is completely dry and all active solvents have left the dope film.

Spray all coats except the first three or four clear coats, to avoid brush marks.

Over thinning is preferred to under thinning.

Addition of retarder will produce a smoother coat, but drying time between coats will be extended.

To get a clean line for the trim colors, apply a light coat of clear dope directly on the masking tape prior to painting. This will help eliminate the trim color from running under the masking tape.

Remember to always bring the dope to room temperature before using.

Rubbing compound and wax polish may be applied after all solvents have escaped (usually 2 weeks, depending upon the weather).

DOPE APPLICATION PROCEDURE (Natural Fabrics).

Step 1. After the cotton or linen fabric is installed in accordance with the pro cedures specified in paragraphs 2-7 and 2-8, the fabric is wetted with distilled water to remove wrinkles and fold creases, which will show in a gloss finish. Water may be applied by rubbing with a clean sponge or rag, or by using a paint spray gun. Do not use tap water. It may contain minerals which will contaminate the fabric.

As water is absorbed by the fibers, the threads swell, resulting in temporary tauting of the fabric panel. The fabric should be allowed to dry before dope application, otherwise the water in the fibers will interfere with the dope penetration and adhesion.

After the fabric has dried, the first coat of dope is applied, brushing in one direction to set the nap with a clean, non-shedding, 2-to 6-inch wide, semi-soft, long bristle paint brush.

To offset the deteriorating effect of mildew or other fungus on natural fibers, especially in damp climates, it is recommended that a fungicide be added to the first coat of dope.

The preferred fungicide is zinc dimethyldithiocarbamate powder, which should be prepared per the manufacturer’s instructions. If no manufacturer’s instructions are available the zinc powder may be stirred in at a ratio of 4 ounces, to one gallon of unthinned nitrate or butyrate dope, after the powder is wetted to a paste with a 50/50 ratio of dope and thinner.

Pre-mixed fungicidal dope, manufactured in accordance with the formula specified in MIL-D-7850, will have a transparent purple tint to indicate the fungicide additive. Dope manufactured with other colors to

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Par 2-21 Page 2-62 identify the manufacturers’ products sold under proprietary trade names may or may not have a fungicidal additive.

The viscosity of the dope should be adjusted to uniformly wet the fabric, indicated by the fabric becoming translucent so that it penetrates through the fabric but does not drip or run down the opposite side. Any dope-runs or pooling on the opposite side will shrink and distort the fabric, and may be visible on the finished surface.

The ideal temperature for application of dope or other coatings is 65 to 75 °F and the humidity should be less than 65 percent. As a general rule, each 10 °F increase or decrease in ambient temperature will increase or decrease drying time by 100 percent. Dope should be allowed to warm to room temperature prior to attempting to adjust the viscosity.

Step 2. Depending upon the quality of the dope and the ratio of thinning, the fabric should start to become taut after the first brushed coat of dope has dried approximately 1 hour at 70 °F. A second, heavier coat is applied by cross brushing at 90 °to the first coat.

Viscosity should be adjusted only as necessary to brush out a heavy uniform coat. If the fabric is not taut, with all sag removed, after the second coat has dried approximately 2 hours, a third coat may be applied.

Step 3. After the fabric has become semi-taut and stabilized with the initial dope application, and the rib lacing and other fabric attachments are completed as detailed in paragraphs 2-9 through 2-12, it is ready for “dressing out” as described in paragraphs 2-13 and 2-14.

NOTE: “Dressing out” means applying all the finishing tapes, reinforcing patches, inspection access ports, and drain grommets, etc.

Step 4. After the covering is a dressed out, one or more coats of clear dope are brushed over all finishing tapes and fabric reinforcing patches. This will balance the thickness of the dope film with the previously coated areas of the fabric. It is very important that the porosity of the fabric be filled while brushing to avoid pinholes showing in the finish.

Step 5. After drying at least 2 hours at 70 °F, a third heavy coat of clear dope is applied over the entire surface, preferably with a paint spray gun if brush marks are to be avoided. After the third coat of dope has dried at least 2 hours at 70 °F, the fabric should be taut and the dope film should show a gloss, depending upon the dope quality and the ratio of thinner added. If not, a fourth coat of clear dope may be applied, in the same manner as the third coat.

NOTE: Three to four clear coats of dope film showing a uniform gloss combined with the aluminumpigmented coats and finish coats is considered satisfactory for light aircraft to 9 lb. per square foot wing loading. Five to eight clear coats, depending upon the quality of the dope and resulting film thickness, are recommended for higher wing loading aircraft to assure the covering does not stretch and lose tension.

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Par 2-21 Page 2-63 Step 6. After the clear coats are found to be satisfactory, two heavy cr oss-coats of aluminumpigmented dope are applied with a spray gun to provide protection from ultraviolet (UV) rays. Tests have shown that UV radiation will deteriorate cotton, linen, and polyester fabric; however, polyester fabric deteriorates at a rate half that of cotton or linen under identical exposure conditions . UV radiation does not deteriorate glass fabric. Aluminum-pigmented dope blocks UV radiation and pr ovides a sanding base. A gauge of ultraviolet protection in the field is to block all visible light from penetrating through the fabric.

Drying time between the two coats should be at least 1 hour at 70 °F.

An option to premixed aluminum dope is to use aluminum-pigment paste. Aluminum paste should be prepared per the manufacturer’s instructions. If no manufacturer’s instructions are available, mix 3 ounces (by weight), of 325-mesh aluminum-pigment paste, to 1 gallon of un-thinned, clear dope. The aluminum paste should first be mixed to a cream consistency with a 50/50 ratio of dope and thinner before mixing into the un-thinned clear dope. A higher ratio of aluminumpigment added to the dope may cause a loss of inner-coat, finish adhesion, and peeling especially when high tack tape is used to mask for the trim colors and registration numbers.

The viscosity of the mixed aluminumpigmented dope should be adjusted for satisfactory spray gun application.

Step 7. After two coats of aluminum-pigmented dope have dried at least 4 hours at 70 °F, the surface may be wet sanded with # 280 grit (or finer) waterproof sandpaper. The aluminum-pigmented dope should be sanded only to develop a smooth surface, not sanded completely off to the clear dope undercoats. Do not sand over screwheads, rib lacing, or any structural sharp edges that will quickly cut through fabric and require patching.

Additional coats of aluminum-pigmented dope may be applied and sanded, depending on the final finish desired. The last coat should not be sanded to assure ultraviolet protection along the edges of the finishing tapes and reinforcing patches is maintained.

Step 8. Three coats of pigmented color finish are applied with a paint spray gun, allowing adequate drying time between coats. The color finish may be wet sanded between coats, if desired, with a fine grit waterproof sandpaper. Adding blush retarder to the final dope finish will improve the gloss. After drying several weeks, a rubbing compound may be used to buff the finish and increase the gloss. A periodic application of a wax polish will help protect the finish from the weather and environmental pollution.

NOTE: Drain holes should be opened soon after all finishing is complete to insure drainage and to aid ventilation of the structure.

Dark colors convert more heat than light colors when exposed to the sun. High temperatures dry out wood structures and deteriorate organic materials in an aircraft structure. Preferably the lighter color shades are applied first and then overcoated with darker trim and registration number colors.

Only high-quality, solvent-resistant crepe paper or polypropylene masking tape should be used to avoid finish bleed under the tape edge. Newspaper printing ink may transfer to a fresh finish and should not be used for masking paper. Plastic sheeting should not be used as a dust cover on a fresh finish due to possible bonding and damage.

**COVERING OVER PLYWOOD.**

Exposed, stressed plywood surfaces, such as wings, must be protected from weather

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Par 2-22 Page 2-64 deterioration with fabric at least equal to that used by the original manufacturer. If the quality is not known, intermediate-grade fabric, meeting TSO-C14b specification, is acceptable. Fabric may be installed in sections with a 1/2 inch edge overlap without covering the overlap with finishing tape. Fabric may also be installed with the edges butted together, and the seam covered with a minimum 1-inch wide finishing tape.

The seams may be oriented in any direction, in reference to the line of flight. However, overlapped seams, not covered with a finishing tape, should be oriented rearward. Fabric should be wrapped completely around a wing’s l eading and trailing edges and other components, where possible, to provide fabric-to-fabric continuity around all edges to avoid a poorly-bonded fabric edge from peeling from the plywood surface causing serious aerodynamic consequences.

After the plywood surface is prepared, and the two pre-coats of clear dope have dried as recommended in paragraph 2-6 g, the fabric is pulled snug and bonded with clear dope around the perimeter of the fabric section. The fabric is then wetted with distilled water to remove fold creases, in the same manner described for fabric panel areas. After the water has evaporated, a heavy coat of low-viscosity clear dope is brushed firmly through the fabric to soften the underlying dope precoat, insuring a good bond. Brushing techniques should be accomplished by moving the brush from one side across to the opposite side to remove all air bubbles and thoroughly saturate the fabric. This is indicated by the plywood grain being easily visible through the translucent fabric. Except for very small imperfections or small dents in the plywood surface, voids are not permissible between the fabric and plywood surfaces.

Voids may allow the fabric to balloon from the plywood surface, creating adverse handling characteristics.

After the first dope coat has dried at least 1 hour at 70 °F, a second heavy coat of clear dope is applied by brush to fill the fabric weave and prevent pinholes. The installation of finish tape around the perimeter of the plywood surfaces, leading edges, and other wear points, is optional but recommended for wear and chafe protection. The application of aluminum-pigmented dope coating, sanding, and fini sh coats will be the same as that specified for fabric panel areas. Reinforcement grommets are not required on drain holes through plywood surfaces.

**COATING APPLICATION DEFECTS.**

Blushing. The appearance of li ght shaded dull areas on the surface as dope dries is th e result of moisture in the atmosphere condensing on a surface due to the cooling effect of the fast-evaporating components of dope thinner escaping from the coating. Blushing can occur at any temperature when the humidity is above 65 percent. There are several ways to remedy this problem. The drying time may be slowed by adding up to 1 quart of blush retarder to 1 gallon of dope or by increasing the temperature of the dope room and eliminating any cooling draft from blowing across the surface. Blushed surfaces may be reworked by spraying several, closely-timed coats of a 50/50 blend of blush retarder and dope thinner to soften and return the dope surface to the original liqui d state. Blush retarder, mixed with dope, may delay the full drying time by several days, but will eventually escape from the dope film if the room temperature is maintained an average of 70 °F.

Pinholes. Voids between the fabric threads that are not filled with the first coats of dope are called pinholes.

They may be caused by fabric contamination, such as oil or finger

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Par 2-22 Page 2-65 prints, but are usually the result of improper dope application. Pinholes are usually found in a second layer of fabric such as finishing tapes and reinforcing patches or over underlying, non-porous structures; such as leading edges, turtle decks, and plywood surfaces.

Any non-porous structure under fabric will act as a backstop and will resist comp lete dope penetration into the fabric. Microscopic cavities between the backstop and fabric collect escaping solvent vapors during the drying process and balloon up through the surface leaving pinholes, or become pinholes when the top of the balloon is sanded. Moisture, in the fabric or on the backstop surface, also inte rferes with complete dope penetration, resulting in pinholes. The remedy for pinholes, at any stage before the final finish coat, is to add blush retarder to low viscosity dope and carefully brush over the affected surfaces to penetrate into and fill the pinholes. Discontinue brushing after five or six strokes to avoid leaving brush marks.

Orange Peel. A rough spray gun-applied finish, similar to the texture of an orange peel, may be caused by one or more of the following conditions:

Viscosity of material being sprayed is too high.

Air temperature is too high.

Spraying in direct sunlight, onto a hot surface or in a drafty/windy condition, which causes a fast solvent evaporation.

Spray gun, tip, cap, and/or needle are not properly matched for the type material being sprayed.

Volume of air available from the compressor not sufficient for spray gun.

Wrong thinner used and drying too fast.

Spray gun not properly adjusted. The spray gun should be adjusted to a uniform spray pattern with the material atomized to deposit fine, wet particles that merge and form a smooth film.

Blisters. One or more of the following conditions my cause blisters:

Freshly coated surface placed in hot sunlight or high temperature area to accelerate drying time, causing the vapor from rapidly evaporating solvents to be trapped.

Excessive high air pressure used to spray heavy coats which “blasts” air bubbles into the coating.

Water or oil in air supply.

Runs, Sags, and Curtains. These defects may be caused by one or more of the following conditions:

Viscosity of material being sprayed is too low.

Coats applied too heavily.

Insufficient drying time between coats.

Spray gun held too close to work surface.

Improperly adjusted spray gun.

Spray Gun Laps and Streaks. These defects may be caused by one or more of the following conditions:

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Par 2-23 Page 2-66 (and 2-32) Spray gun not properly adjusted to spray a wet, smooth surface.

Overspray on a partially-dried surface.

Spray pattern not sufficiently overlapped on each pass.

Viscosity of material being sprayed is too high.

Metallic finishes sprayed too heavily allowing metallic pigments to move or flow after deposit, causing a marbled appearance.

2-24.2-29. [RESERVED.]

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Par 2-30 Page 2-67

**SECTION 3. INSPECTION AND TESTING**

GENERAL. All components of the covering should be inspected for general condition.

Loose finishing tape and reinforcing patches;

chafing under fairings; brittle, cracking, peeling, or deteriorated coatings; fabric tears and rock damage; broken or missing rib lacing;

and rodent nests are unacceptable. The entire fabric covering should be uniformly taut with no loose or wrinkled areas, or excess tension which can warp and damage the airframe.

Excess Tension. There are no methods or specifications for measuri ng acceptable fabric tension other than observation. Excess tension may pull and warp critical components, such as: longerons; wing ribs;

and trailing edges out of position, weakening the structure.

Excess tension with cotton, linen, and glass fiber fabric covering is usually caused by excessive dope film on a new covering, or continuous shrinking of an originally satisfactory dope film as the plasticizers migrate from the dope with age. Heat from sun exposure accelerates plasticizer migration.

Excess tension with polyester fabric, coated with dope, is usually caused by the combined tension of the heat tautened polyester fabric and continuous shrinking of the dope film as the plasticizers migrate from the dope with age.

Loose Fabric. Fabric that flutters or ripples in the propeller slipstream, balloons, or is depressed excessively in flight from the static position, is unacceptable.

Loose or wrinkled cotton, linen, and glass fabric covering may be caused by inadequate dope film; poor quality dope; fabric installed with excess slack; or by a bent, broken, or warped structure.

Loose or wrinkled polyester fabric covering, finished with coatings other than dope, may be caused by inadequate or excessive heat application; excess slack when the fabric was installed; or bent or warped structure.

Polyester fabric which does not meet aircraft quality specifications will very likely become loose after a short period of time.

Glass fabric covering should be tested with a large suction cup for rib lacing cord failure and reinforcing tape failure caused by chafing on all wing ribs and other structural attachments throughout the airframe. Particular attention should be given to the area within the propeller slipstream. If failure is indicated by the covering lifting from the static position, the rib lacing cord and reinforcing tape must be reinstalled with double the number of original laces.

NOTE: Temporary wrinkles will develop in any fabric coated and finished with dope, when moisture from rain, heavy fog, or dew is absorbed into a poor-quality dope film, causing the film to expand.

Temporary wrinkles may also develop with any type of thick coatings, on any type of fabric, when an aircraft is moved from a cold storage area to a warm hangar or parked in the warming sunshine, causing rapid thermal expansion of the coating.

Coating Cracks. Fabric exposed through cracks in the coating may be initially tested for deterioration by pressing firmly with a thumb to check the fabric’s strength. Natural fibers deteriorate by exposure to ultraviolet radiation, mildew, fungus from moisture, high acid-content rain, dew, fog, pollution, and age.

Polyester filaments will deteriorate by exposure to UV radiation.

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Par 2-32 Page 2-68 Glass fabric will not deteriorate from UV exposure, but will be deteriorated by acid rain, dew fallout, and chaffing if loose in the prop blast area.

Cotton, linen, and glass fabric coverings are dependent solely on the strength and tautening characteristics of the dope film to carry the airloads. Dope coatings on heattautened polyester fabric will also absorb all the airloads because the elongation of polyester filaments are considerably higher than the dope film. Polyester fabric that is coated with materials other than dope, is dependent solely on the heat tautening and low-elongation characteristics of the polyester filaments to develop tension and transmit the airloads to the airframe without excess distortion from a static position.

Cracks in coatings will allow any type of exposed fabric to deteriorate. Cracks should be closed by sealing or removing the coatings in the immediate area and replace with new coatings, or recover the component.

**FABRIC IDENTIFICATION.**

Cotton Fabric meeting TSO-C15 or TSO-C14 can be identified by an off-white color and thread count of 80 to 94 for TSO-C14b and 80 to 84 for TSO-C15d in both directions.

Aircraft linen conforming to British specification 7F1 may be identified by a slightly darker shade than cotton fabric and irregular thread spacing. The average thread count will be about the same as Grade A fabric (TSO-C15d). The non-uniformity of the linen thread size is also noticeable, with one thread half the size of the adjacent thread. When viewed under a magnifying glass, the ends of the cotton and linen fiber nap may be seen on the backside. The na p is also seen when the coating is removed from the fr ont or outside surface. A light-purple color showing on the back side of cotton or linen fabric indicates a fungicide was present in the dope to resist deterioration by fungus and mildew.

Polyester fabric conforming to TSO-C14b or TSO-C15d is whiter in color than cotton or linen. The fabric styles adapted for use as aircraft covering have a variety of thread counts, up to ninety-four (94), depending on the manufacturing source, weight, and breaking strength. Polyester is a monofilament and will not have any nap or filament ends showing.

Glass fabric is manufactured white in color, and one source is precoated with a blue-tinted dope as a primer and to reduce weave distortion during handling. Thread count will be approximately 36 threads per inch. Glass fabrics are monofilament and will not have any nap or filament ends showing unless they are inadvertently broken.

When a small fabric sample can be removed from the aircraft and all the coatings removed, a burn test will readily distinguish between natural fabric, polyester, and glass fabric. Cotton and linen will burn to a dry ash, polyester filaments will melt to a liquid and continue burning to a charred ash, and glass filaments, which do not support combustion, will become incandescent over a flame.

**COATING IDENTIFICATION.**

Nitrate or butyrate dope must be used to develop tension on cotton, linen, and glass fabrics. When a small sample can be removed, burn tests will distinguish nitrate dope-coated fabric from butyrate dope-coated fabric by its immediat e ignition and accelerated combustion. Butyrate dope will burn at less than one-half the rate of nitrate dope. Coating types other than nitrate or butyrate dope may have been used as a finish over dope on cotton, linen, and glass fiber fabric coverings.

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Par 2-32 Page 2-69 If the fabric type is determined to be polyester, coating identification should start by reviewing the aircraft records and inspecting the inside of the wings and the fuselage for the required fabric source identification stamps for covering materials authorized under the STC. The manual, furnished by the holder of the STC-approved fabric, should be reviewed to determine whether the coati ngs are those specified by the STC.

Coating types, other than those authorized by the original STC, may have been used with prior FAA approval, and this would be noted in the aircraft records. The presence of dope on polyester can be detected by a sample burn test.

**STRENGTH CRITERIA FOR AIRCRAFT FABRIC.**

Minimum performance standards for new intermediate-grade fabric is specified in TSO-C14b, which references AMS 3804C.

Minimum performance standards for new Grade A fabric are specified in TSO-C15d, which references AMS 3806D.

The condition of the fabric covering must be determined on every 100-hour and annual inspection, because the strength of the fabr ic is a definite factor in the airworthiness of an airplane. Fabric is considered to be airworthy until it deteriorates to a breaking strength less than 70 percent of the strength of new fabric required for the aircraft. For example, if grade-A cotton is used on an airplane that requires only intermediate fabric, it can deteriorate to 46 pounds per inch width (70 percent of the strength of intermediate fabric) before it must be replaced.

Fabric installed on aircraft with a wing loading less than 9 lb. per square foot (psf), and a Vne less than 160 mph, will be considered unairworthy when the breaking strength has deteriorated below 46 lb. per inch width, regardless of the fabric grade. Fabric installed on aircraft with a wing loading of 9 lb. per square foot and over, or a Vne of 160 mph and over, will be considered unairworthy when the breaking strength has deteriorated below 56 lb. per inch width.

Fabric installed on a glider or sailplane with a wing loading of 8 lb. per square foot and less, and a Vne of 135 mph or less, will be considered unairworthy when the fabric breaking strength has deteriorated below 35 lb. per inch width, regardless of the fabric grade.

FABRIC TESTING. Mechanical devices used to test fabric by pressing against or piercing the finished fabric are not FAA approved and are used at the discretion of the mechanic to base an opinion on the general fabric condition.

Punch test accuracy will depend on the individual device calibration, total coating thickness, brittleness, and types of coatings and fabric. Mechanical devices are not applicable to glass fiber fabric that will easily shear and indicate a very low reading regardless of the true breaking strength. If the fabric tests in the lower breaking strength range with the mechanical punch tester or if the overall fabric cover conditions are poor, then more accurate field tests may be made.

Cut a 1-1/4-inch wide by 4-inch long sample from a top exposed surface, remove all coatings and ravel the edges to a 1-inch width.

Clamp each end between suitable clamps with one clamp anchored to a support structure while a load is applied (see table 2-1) by adding sand in a suitable container suspended a few inches above the floor. If the breaking strength is still in question, a sample should be sent to a qualified testing laboratory and breaking strength tests made in accordance with American Society of Testing Materials (ASTM) publication D5035.

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Par 2-42 Page 2-70 NOTE: ASTM publication D1682 has been discon tinued but is still referred to in some Aerospace Material Specification (AMS). Th e grab test method previously listed in ASTM D1682, sections 1 through 16, has been superseded by ASTM publication D5034.

The strip testing method (most commonly used in aircraft) previously listed in ASTM D1682, sections 17 through 21, has been superseded by ASTM publication D5035.

REJUVENATION OF THE DOPE FILM. If fabric loses its strength, there is nothing to do but remove it and recover the aircraft. But if the fabric is good and the dope

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Par 2-42 Page 2-71 is cracked, it may be treated with rejuvenator, a mixture of very potent solvents and plasticizers, to restore its resilience. The surface of the fabric is cleaned and the rejuvenator sprayed on in a wet coat, and the solvents soften the old finish so the plasticizers can become part of the film. When the rejuvenator dries, the surface should be sprayed with two coats of aluminum-pigmented dope, then sanded and a third coat of aluminum-pigmented dope applied, followed with the colored-dope finish.

When repairing, rejuvenating, and refinishing covering materials approved under an STC, instructions in the manual furnished by the material supplier should be followed.

2-36.2-41. [RESERVED.]

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Par 2-42 Page 2-72 SECTION 4. REPAIRS TO FABRIC COVERING

GENERAL. All materials used to make repairs to fabric covering must be of a quality at least equal to the original materials. Workmanship and repair methods must be made in a manner that will return the fabric covering to its original airworthy condition.

Any combination of seams hand-sewn and overlapped and doped may be used to ma ke repairs and install new fabric sections. (See paragraph 2-7.)

All pigmented dope coats, including aluminumpigmented coats, should be removed to the clear dope preliminary coats before installing a new fabric section or finishing tape on the old fabric. The appropriate dope thinner (see paragraph 2-20) may be applied with a brush to soften the old dope. The softened coats can then be removed down to the clear dope coats by scraping with a dull-bladed knife while supporting the fabric from the back side. Removing the old dope by sanding is optional.

Avoid allowing dope to run down the back side of the fabric or drip through the wing onto the back side of the opposite surface, which will cause cosmetic damage and will show in a high gloss finish.

Repairing a new fabric section over two adjacent wing ribs is considered a major repair. A log book entry and an FAA Form 337 must be processed accordingly.

All fabric patch edges not covered with a finishing tape should have a pinked e dge or a 1/4-inch raveled edge.

Where the edge of a new fabric section will be located within 1 inch of a structural member to which the fabric is attached by rib lacing or other methods, the new fabric section should be extended 3 inches past the structural member.

NOTE: Before installing new finishing tape, duplicate the original rib lacing or other attachments without removing the original rib lacing or attachment. Removing the original finishing tape is optional.

When installing large sections of fabric on a wing or other components, all machine-sewn and/or overlapped and doped seams should be made in accordance with the guidelines specified in paragraphs 2-7 and 2-8.

When repairing a covering material other than cotton or linen, which was approved with the manufacturer’s type certificate (TC), or approved under the authority of an STC, follow the repair instructions furnished by the aircraft manufacturer or supplier of the STC-approved covering materials.

**REPAIR OF TEARS AND ACCESS OPENINGS.**

When all the original fabric is intact, an opening may be repaired by sewing the two sides together with a curved needle as illustrated in figures 2-13 and 2-14. The fabric edges should be pulled together uniformly with no wrinkles. Before sewing, remove the old dope coats down to the clear dope coats a distance of 2 inches on each side of the opening. The hand-sewn thread quality should be at least equal to that specified in table 2-2 and treated with wax (paraffin-free or beeswax) to prevent fraying, or use the proper thread on the STC application. After sewing, apply a coat of clear dope over the cleaned area and install a 3-inch wide finishing tape, centered over the stitches.

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Par 2-42 Page 2-73

FIGURE 2-13. Repair of tears in fabric.

Finishing tapes should be well saturated with dope and smoothed out with no voids or wrinkles during installation. After drying for 1 hour at 70 °F, additional coats of clear dope followed with pigmented dope are applied as detailed in paragraph 2-21.

If the opening is more than 8 inches long but less than 16 inches long in any direction, the finishing tape width should be increased to 4 inches.

The finishing tape width should be increased to 6 inches if the opening is more than 16 inches long in any direction, is located on a wing top surface, or the aircraft Vne speed is greater than 150 mph.

SEWN-PATCH REPAIR. Openings that cannot be repaired by closing with stitches may be repaired by sewing in a new fabric section.

The edges of the fabric around the opening should be trimmed straight on four sides to facilitate the installation of straight sections of finishing tape over the stitches.

After cutting out the damaged section and removing the coatings as detailed in paragraph 2-42, the new fabric section should be sized to allow folding both edges of the fabric back 1/2-inch to increase the stitch tear resistance. Temporarily attach the four corners in position with thread. Start with a double thread with a square knot at the end (see figure 2-14) and continue stitching in the manner described in figures 2-13 and 2-14. When the stitching is complete, wet the new fabric section as described in paragraph 2-21 to remove any creases. After drying, apply one coat of clear dope on the new fabric, as described in paragraph 2-21.

When the first coat of dope has dried 1 hour at 70 °F, apply a 3-inch wide finishing tape, centered over the stitches. The finishing ta pe should be well saturated with

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Par 2-44 Page 2-74 Step 1. Tie the thread ends together with a square knot.

Step 2. Start the stitch by routing through the loop, positioning the square knot on the inside.

Step 3. Complete the hand-stitch and secure the stitching with a half hitch at a maximum each 10 stitches.

FIGURE 2-14. Hand-stitch detail.

dope and smoothed out with no voids or wrinkles.

Additional coats of clear dope and pigmented dope are then applied to obtain the desired tautness and finish, as described in paragraph 2-21.

If the opening is more than 8 inches but less than 16 inches long in any direction, the finishing tape should be 4 inches wide.

The finishing tape width should be 6 inches wide if the opening is over 16 inches long in any direction, is located on a wing top surface, or the aircraft Vne speed is greater than 150 mph.

DOPED-ON PATCH REPAIR. An opening not over 8 inches in length in any direction, on an aircraft with a Vne speed less than 150 mph, may be repaired with a 2-inch overlapped and doped patch. The opening should be trimmed to eliminate any irregular edges and old pigmented dope coats removed as described in paragraph 2-42.

When installing a fabric patch over a small opening, the loose edge of the fabric around the opening may be secured by extending a series of small threads, from the edge across the opening, to the opposite side. After the patch is completed, the threads may be clipped and removed through an access port or left in place.

The fabric patch is installed by applying a coat of clear dope around the opening, then positioning the patch over the opening. Brush out any void or wrinkles while saturating only the fabric overlap area. After the first coat of clear dope around the edge has dried 1 hour at 70 °F, wet the fabric patch to remove any creases as described in paragraph 2-21. After drying, apply additional coats of clear dope and pigmented dope over the entire patch as described in paragraph 2-21.

If the opening is less than 8 inches in length in any direction and the aircraft Vne speed is greater than 150 mph, a 2-inch wide finishing tape should be installed on all sides, centered on the edge of the 2-inch overlap patch.

If the opening is more than 8 inches but less than 16 inches in length, in any direction on an aircraft with a Vne speed less than

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Par 3-1 Page 3-75 150 mph, it may be repaired with a doped patch, which is overlapped 1/4 of the opening maximum dimension. The maximum overlap should not exceed 4 inches.

If the opening is more than 8 inches but less than 16 inches in length in any direction, the repair is located on a wing top surface, and the aircraft Vne speed is greater than 150 mph, the patch overlap should be 4 inches and a 2-inch wide finishing tape installed on all sides, centered on the edge of the patch.

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Par 3-1 Page 3-76 If the opening is more than 16 inches in length in any direction and the Vne speed is less than 150 mph, the patch overlap should be 4 inches and the finishing tape should be 4 inches in width, centered on the edge of the patch.

If the opening is more than 16 inches in length in any direction and the Vne speed is greater than 150 mph, the patch overlap should be 4 inches and the finishing tape should be 6 inches in width, centered on the edge of the patch.

2-46.2-51. [RESERVED.]

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Par 3-1 Page 3-77 CHAPTER 3. FIBERGLASS AND PLASTICS

**SECTION 1. REPAIR OF LIGHT LOAD LAMINATE STRUCTURES**

GENERAL. There is a wide variation in the composition and structural application of laminates, and it is essential that these factors be given major consideration when any restoration activities are undertaken. To a similar extent, there also exist many types of laminate structure repairs that may or may not be suitable for a given condition. For this reason, it is important that the aircraft or component manufacturer’s repair data be reviewed when determining what specific type of repair is permissible and appropriate for the damage at hand.

NOTE: Review Material Safety Data Sheets for material to be used. When handling materials, prepreg fabrics, or parts with prepared surfaces, observe shelf life. Latex gloves and approved masks must be worn.

The materials used in the repair of laminate structures must preserve the strength, weight, aerodynamic characteristics, or electrical properties of the original part or assembly. Preserva tion is best accomplished by replacing damaged material with material of identical chemical composition or a substitute approved by the manufacturer.

To eliminate dangerous stress concentrations, avoid abrupt changes in cross-sectional areas. When possible, for scarf joints and facings, make small patches round or oval-shaped, and round the corners of large repairs.

Smooth and properly contour aerodynamic surfaces.

Test specimens should be prepared during the actual repair. These can then be s ubjected to a destructive test to establish the quality of the adhesive bond in the repaired part. To make this determination valid, the specimens must be assembled with the same adhesive batch mixture and subjected to curing pressure, temperature, and time identical to those in the actual repair.

FIBERGLASS LAMINATE REPAIRS. The following repairs are applicable to fiberglass laminate used for non-structural fairing, covers, cowlings, honeycomb panel facings, etc. Prior to undertaking the repair, remove any paint by using normal dry sanding methods. Bead blasting may be used but caution must be exercised not to abrade the surfaces excessively.

NOTE: Chemical paint strippers must not be used.

NOTE: These repairs are not to be used on radomes or advanced composite components, such as graphite (carbon fiber) or Kevlar.

CAUTION: Sanding fiberglass laminates gives off a fine dust that may cause skin and/or respiratory irritation unless suitable skin and respiration protection is used.

Sanding also creates static charges that attract dirt or other contaminants.

Check for voids and delamination by tap testing. (See chapter 5.) When the surface of a fiberglass laminated structure is scratched, pitted, or eroded; first wash with detergent and water to remove all of the dirt, wax, or oxide film. Then scrub th e surface with an acceptable cleaner. After the surface is thoroughly cleaned, sand it with 280-grit sandpaper, and again use an acceptable cleanser to remove any sanding residue and moisture.

This is

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Par 3-3 Page 3-78 essential, as any moisture remaining on the surface will inhibit the cure of the resin. Dry the fiberglass laminate thoroughly prior to bonding repair. Mix enough resin, using the manufacturer’s instructions, to completely cover the damaged area, and apply one or two coats.

Cover the resin with a peel ply to exclude all air from the resin while it is curing. After the resin has cured, remove the film and file or sand the surface to conform to the original shape of the part. Ensure that all edges of the laminate part are sealed to prevent water absorption. Then refinish it to match the rest of the structure.

Superficial scars, scratches, surface abrasion, or rain erosion can generally be re paired by applying one or more coats of a suitable low temperature resin, catalyzed to cure at room temperature, to the abraded surface. The number of coats required will depend upon the type of resin and the severity of the damage.

Damage not exceeding the first layer or ply of fiberglass laminate can be repaired by filling with a putty consisting of a compatible roomtemperature-setting resin and clean short glass fibers. Before the resin sets, apply a sheet of peel ply over the repair area and work out any bubbles and excess resin. After the resin has cured, sand off any excess and prepare the area for refinishing.

Damage deep enough to seriously affect the strength of the laminate (usually more than the first ply or layer of fabric) may be repaired as illustrated in figure 3-1. Coat the sanded area with room-temperature-setting resin and apply contoured pieces of glass fabric soaked in resin. Apply a peel ply sheet over the repair and work out any bubbles and excess resin.

After the resin has cured, scrape off the excess resin and sand the surface of the repair to the original contour.

FIGURE 3-1. Typical laminate (facing) repair.

Damage that extends completely through one facing and into the core requires the replacement of the damaged core and facing.

A method for accomplishing this type of repair is shown in figure 3-2. An alternate method for repairing the facing is shown in figure 3-3.

The damaged portion is carefully trimmed out to a circular or oval shape and the core material is removed completely to the opposite facing. Exercise caution so as not to damage the opposite facing or to start delamination between the facings and the core around the damage.

FIGURE 3-2. Typical core and facing repair.

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Par 3-2 Page 3-79

FIGURE 3-3. Typical stepped joint repair.

Use replacement core stock of the same material and density as the original (o r an acceptable substitute) and cut it to fit snugly in the trimmed hole. Observe the direction of the original core . When all of the pieces of replacement facing laminations are cut and soaked in resin, coat all surfaces of the hole and the scarfed area with resin. Then coat all surfaces of the core replacement with resin and inse rt it into the hole. After all of the pieces of resin-impregnated glass-fabric facing are in place and lined up with the original fiber-orientation, cover the entire area with a piece of peel ply and carefully work down the layers of fabric to remove any air bubbles and excess resin. Apply light pressure by means of sand bags or a vacuum bag.

When the resin has cured, sand the repair to match the original contour and refinish the surface.

**REPAIRING HOLES.**

Scarf Method. If the damaged area is less than 3 inches in diameter, the damage may be removed by either sanding with a power sander or hand sanding with 180-grit sandpaper.

Scarf back the edges of the hole about 50 times the thickness of the face ply.

Thoroughly clean out all of the sanding residue with a cloth wet with an acceptable cleanser.

Prepare the patches by (see figure 3-4) laying the proper weight fiberglass cloth impregnated with resin on a piece of peel ply. A weight of resin equal to the weight of the patch provides a 50-percent ratio.

FIGURE 3-4. Preparing the fiberglass sandwich.

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Par 3-2 Page 3-80 Make a sandwich by laying a second layer of peel ply over the patch before cutting it to the required size and shape. Sandwiching will prevent the patch from raveling when cut.

Brush a good coat of resin over the scarfed area. Remove one piece of peel ply from the first patch and lay the patch in place. Work all of the air out of the resin and remove the top peel ply. Cut the next larger patch so it will overlap the first patch by at least one-half inch. Remove one piece of peel ply from this patch and center the patch over the first one. Work all of the air out of the resin.

Continue laying in patches, each overlapping the one below it by at least one-half inch, until you have the required number of layers (see figures 3-5 and 3-6) plus an extra ply to restore original strength to the repaired area.

FIGURE 3-5. Scarfed repair to a nonstructural laminated fiberglass component.

FIGURE 3-6. Symmetrical layup same as original number of plies, plus one extra ply.

Cover the entire repair with peel ply and carefully work out all of the air bubbles from the resin. Apply pressure over the repair with tape or sandbags and allow it to cure. After the repair has cured, remove the excess resin by filing or sanding the surface to the contour of the original part. Smooth the surface with fine sandpaper and refinish it to match the original part.

An alternate layup method that works equally well is to place the larger patch over the scarfed area first, and then each subsequent smaller patch over this. Both types of repair are finished in exactly the same way.

The scarfed joint method (see figure 3-7) is normally used on small punctures up to 3 or 4 inches in maximum dimension and in facings that are made of thin fabric that is difficult to peel.

Step-Joint Method. The scarf method of repairing a laminated fiberglass face sheet of a honeycomb structure is the easiest method to use. In this type of repair, the damage is outlined with a compass. If a square or rectangular repair is more appropriate then the damage is outlined using a straight-edge and a compass to round out the corners.

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Par 3-3 Page 3-81

FIGURE 3-7. Typical scarf joint repair.

The configuration of the repair should be that which will remove the least amount of sound material. Extend the cleaned -out area for a distance equal to the number of plies to be removed, less 1 inch. For example, if you must remove three plies, extend the repair for 2 inches beyond the cleaned-out area. Each layer should be 1 inch beyond the layer below.

Use a sharp knife or other type of cutter to cut through the top layer, being careful not to damage the underneath layer. Use several passes with the knife rather than one deep cut. (See figure 3-8.)

Begin with one corner of the patch and carefully pry it loose and peel it up until all of the layer is removed. Next, mark the exposed layer 1/2 inch inside the opening and carefully cut and remove it. Continue until you have removed all of the damaged or delaminated layers.

FIGURE 3-8. Carefully cut through each layer of fiberglass cloth and remove it from the damaged area.

Lightly sand, then scrub the entire area with an acceptable cleanser. Prepare the patches exactly as you did for the scarf method, cutting each layer to exactly the size of the material removed. Brush in a coat of resin, lay in the patch of the smallest size, and carefully work out all of the air bubbles from the resin. Now, lay in the next larger size patch to lock the first layer of fiberglass cloth into place. Repeat the process until the damage area is filled.

Butt the top layer of cloth to the opening in the face ply and cover the entire repair with peel ply. Carefully work all of the air bubbles out of the resin and put pressure on the repair with either sandbags, or another appropriate method, such as vacuum bagging. (See figure 3-9.) After the top repair has hardened, repeat the process on the bottom.

**SAMPLE BAGGING AND CURING PROCESS.**

Figure 3-9 shows a typical bagging arrangement for a localized repair in which patch plies of prepreg are cured with a layer of adhesive, and a heating blanket is used to supply heat.

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Par 3-3 Page 3-82

FIGURE 3-9. Sample bagging layup cross section.

The materials used for most bonded repairs require elevated temperatures and pressure, during their cure, to develop full strength. The following paragraphs describe the operations required to enclose the repair in a vacuum bag. When the part to be cured can be placed in an autoclave, additional pressure and heat can be applied. For cured-in-place parts, vacuum pressure and portable heat blankets usually suffice.

When selecting materials, especially the parting agent film, the temperature at which the repair is to be cured must be known.

Polyvinyl alcohol (PVA) film is ideal when the bonding temperature does not exceed 250 °F.

PVA film has very high tear resistance and may be heat-sealed effectively. When the bond temperature is not above 180 °F, polyvinyl chloride film can be used. For temperatures up to 450 °F, a polyvinyl fluoride film is used. These three types of films are available in a variety of weights and widths.

Most nylon bagging films are used for temperatures up to 400 °F.

When all repair details are in place and ready to be cured, they are enclosed in a bag of plastic film or thin rubber. Air is removed from the bag by a vacuum source so that atmospheric pressure exerts a pressure on the repair as it is cured.

To provide a path to draw off the air initially inside the bag, layers of fiberglass cloth or similar noncontaminating materials, known as breather plies, are placed inside the bag.

When prepreg is being cured as part of the repair, it is sometimes necessary to bleed off excess resin. To do this, layers of fiberglass cloth or similar materials known as bleeder plies are placed over the prepreg. Some repairs have been made with a net resin prepreg that does not require bleeding, and therefore does not require bleeder plies.

Porous separator plies or film are used between the prepreg and the bleeder and nonporous separator plies or film are used between the bleeder and breather plies to control the flow of resin.

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Par 3-4 Page 3-83 Small parts may be envelope bagged (i.e., enclosing the entire part in the bag). Larger parts with localized repairs can be bagged by sealing the surface completely around the repair areas with sealing tape and applying the bagging material to the sealing tape.

CAUTION: The whole panel must be vacuum bagged to prevent delamination in sandwich skins when using an oven or autoclave. Contoured parts must be restrained with tooling to prevent warpage.

When the heat for curing the repair is provided by a heat blanket, the blanket can be either inside or outside the vacuum bag. However, the blanket should be covered to minimize heat loss, and the blanket should be separated from direct contact with most bagging materials by layers of fiberglass cloth.

This will prevent localized overheating that could damage the bag. It is sometimes helpful to place a thin aluminum sheet under the heating blanket to minimize localized heating. A thin rubber blanket can help smooth the surface of the material being cured. A pressure plate should be used when two or more heat blankets are applied to the same repair.

NOTE: Understanding that various resins behave differently during cure, the choice of bagging arrangements will often vary with the material being cured.

The procedure for the bagging arrangement is as follows:

Place a peel ply over the patch material to provide a surface finish for subsequent bonding or painting if not previously accomplished. Place a layer of porous separator cloth over the patch, extending beyond the prepreg and the adhesive. Smooth to avoid wrinkles.

With the patch material in place, place the end of the thermocouple wire next to the edge of the prepreg. Tape the wire to the structure inside the bag with heat-resistant tape. The tape should not be in contact with the prepreg or the adhesive.

Place bleeder plies as shown, extending 2 to 3 inches beyond the patch. The number of bleeder plies needed will vary with the type of resin and the resin content required.

Place a layer of nonporous parting film over the bleeder plies, cut 1 inch smaller than the bleeder plies. This layer is intended to stop resin flow from bleeder plies into breather plies while still providing an airflow path when vacuum is applied.

If a pressure plate is used, place it over the previous separator ply. The plate is frequently perforated with small holes to permit airflow to the breather plies. Bleeder plies may be necessary when using a pressure plate.

NOTE: Pressure on the repair will be reduced if the pressure plate does not conform to the repair.

Place the heat blanket over the assembly, making sure it extends 3 to 4 inches beyond the material to be cured.

One or more thermocouples should be in contact with the heat blanket to monitor its temperature. Additional thermocouples should be placed near the curing repair to monitor the temperature of the curing resin.

When using a heat blanket as the heat source, four to six layers of fiberglass surface breather or the equivalent should be

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Par 3-10 Page 3-84 used over the heat blanket. This will insulate and prevent damage to the nylon bagging film. Ensure that the breather plies are in contact with the bleeder plies so that an air passage exists.

Place a bead of sealing tape against the parent ma terial around the edge of the breather plies.

Seal the thermocouple wires to prevent vacuum leakage.

NOTE: Two layers of sealing tape may be required in order to provide a good seal.

Cover with a suitable vacuum bag, smoothed to minimize wrinkles. Press the bag firmly onto the sealing tape to obtain an air-tight seal. Place pl eats in vacuum bag to allow the bag material to stretch.

Install two vacuum probes or sniffers through openi ngs cut in the bag. On e will be used for the vacuum gauge and the

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Par 3-10 Page 3-85 other will be connected to the vacuum source. The vacuum probe must sit on the breather plies, but must not touch the patch or adhesive.

NOTE: Place the vacuum gauge on the opposite side of the vacuum port, where applicable. Do not place vacuum probes near repair area.

Connect the vacuum source and smooth the bag by hand pressure as the air is removed. Check for leaks and reseal as necessary. A minimum vacuum of 22 inches of mercury is required.

Place insulating material over the vacuum bag to prevent heat loss.

Apply power to the heat blanket and control its temperature as specified for the material being cured.

Observe cure time requirements established by the product manufacturer.

3-5.3-9. [RESERVED.]

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Par 3-10 Page 3-86 SECTION 2. METALLIC SANDWICH SECONDARY STRUCTURE

**REPAIRS**

REPAIRS TO METALLIC SANDWICH SECONDARY STRUCTURE. Magnesium, titanium, or stainless steel facings require special procedures that are not included in the following methods of repair. Aluminum alloys such as 7075-T6, 2024-T3, and 2014-T6 are commonly used for the repair of facings for sandwich structural parts having aluminum facings. For maximum corrosion resistance, use only clad aluminum for repairs to clad aluminum alloy facings.

Dents, scratches, or fractures, not exceeding 1/4 inch in largest dimension in aluminum facings, may be repaired with a suitable filler such as viscous epoxy resin. Dents that are delaminated shall not be filled but repaired. Thoroughly clean the repair area with fine sandpaper and acetone before a pplying the filler. After the resin has partially cured, remove any excess resin with a sharp plastic scraper. When the resin has completely cured, sand to the original contour. If the damage included a fracture, reclean the area around the filled hole and apply a surface patch.

Fractures or punctures in one facing and partial damage to the core of an aluminum-faced laminate may be repaired by several different methods. The technique used will depend upon the size of the damage, the strength required, and the aerodynamic loads of the area involved. If the repair requires aerodynamic smoothness, the facing surrounding the repair core cavity may have to be step cut to one-half its thickness.

This can be done by using a router with an end mill bit and a template.

Damage that extends completely through the core and both facings may be repaired using the same general techniques as those used for repairing fiberglass laminates when both facings are accessible.

After locating the extent of the total damaged area by tapping or other nondestructive test methods, remove the damaged facing and that portion of the core material that is also affected. The depth to which the core must be removed will depend upon the type of core material and the method of repair. The replacement core material must be the same material and core cell size as the original. Fabricate core material to shape, keeping the same core ribbon or grain direction. When a substitution is permissible, wood or glass-fabric honeycomb cores are sometimes used in the repair of aluminum honeycomb cores, as they are generally easier to shape. Typical ty pes of core replacements are shown in figure 3-10. Resin fills can be used to replace the core and facing where smaller core damage exists.

Phenolic microballoons, low-density insulating materials, and/or other ingredients are added to lower the density and give greater flexibility.

For the repair of larger holes in which it is inconvenient to use a face patch because of aerodynamic smoothness requirements in that area, both the core and facing are sometimes replaced with glassfiber fabric discs and resin. Undercut the core, as shown in figure 3-11, in order to obtain a better bonding of the fill with the facing. Fill the core cavity with accurately shaped resin-satura ted glass cloth discs, and press each ply down to remove any air bubbles. Special care should be taken that the final plies fit well against the underside of the top facing. When the core cavity is filled, close the cutout in the facing with resinimpregnated glass fiber fabric discs that have been precut to size.

Overlap repairs, typically called scab patches, have a long history of use in repairing aircraft structures.

These repairs simply cover the damaged area with patch material. Overlap repairs can be bonded and/or mechanically

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A. REPLACEMENT WITH CORE PLUG OF SAME fastened. Figure 3-12 shows a cross-section of a typical OR OTHER TYPES OF CORE MATERIAL bonded and mechanically fastened repair. The damaged CORE PLUG area may or may not need to be filled.

**ORIGINAL HONEYCOMB ROUTED AREA**

**CORE EPOXY-RESIN**

ADHESIVE Bonded overlap repairs work well on most structures. The overlap repair consists of a solid patch material such as metal, precured B. RESIN FILL laminates, prepreg or wet layup material co-cured in place.

**EPOXY-RESIN FILL**

**ROUTED AREA**

ORIGINAL PARTIAL OR COMPLETE HONEYCOMB DEPTH OF CORE Bolt or blind rivet. Mechanically fastened, CORE bolted, or blind rivet repairs are primarily used for thick structures. Primary concerns include bolt/rivet spacing, fastener diameter, number C. RESIN FILL, UNDERCUT CORE of fasteners, and sealant type.

**ROUTED AREA**

EPOXY-RESIN FILL PARTIAL OR COMPLETE /DEPTH OF CORE Core potting is the process of filling the core cutout ORIGINAL HONEYCOMB CORE UNDERCUT with a curable paste filler material. If the damage is CORE sustained in an area with an already-potted core, the replacement core should also be potted. In other cases, if the honeycomb or foam core is damaged, it may be potted rather than plugged if the damaged area is small D. GLASS-CLOTH-RESIN LAMINATE, UNDERCUT CORE (1 inch or less). Remove the face sheet with a power

GLASS-CLOTH RESIN LAMINATE router, using a router template to prevent injury to ORIGINAL CORE HONEYÇOMB ROUTED AREA UNDERCUT CORE remove one of the face skins only, a face skin and part undamaged face skin. The router may be adjusted to of the core, a face skin and all of the core, or both the face skins and the core. (See figure 3-13.)

**E. REPLACEMENT WITH CORE PLUG OF SAME**

**TYPE OF SANDWICH CONSTRUCTION**

**ORIGINAL COMPLETE SANDWICH ROUTED AREA**

**HONEYCOMB CORE PLUG EPOXY-RESIN**

CORE /ADHESIVE FIGURE 3-10. Typical types of core replacement.

- CORE - ORIGINAL CORE TO SKIN ADHESIVE

- JAGGED EDGES

**SECTION SHOWING UNDERCUT**

- FINISHED UNDERCUT AREA FIGURE 3-11. Typical undercut core material cavity.

Par 3-10 Page 3-87

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Par 3-10 Page 3-88 It may be necessary when routing a tapered section such as an aileron to use a wedgeshaped block between the routing template and the upper surface. This will allow the router to cut the core material parallel with the lower surface. (See figure 3-14.)

Select the appropriate potting adhesive as recommended by the manufacturer. Mix a sufficient quantity of filler to fill the hole and add microballoons if they are needed to serve as a filler. When the resin and filler are thoroughly mixed according to the manufacturer’s recommendations, pour the mixture into the hole filling all of the cells, then work out all of the bubbles with a toothpick. If performing an overlay repair fill the core cavity to slightly above the part’s surface. If performing a flush repair, fill the core cavity to slightly above the original core.

Cure the compound according to the manufacturer’s directions. Trim the top of the cured potting compound flush with the surface, for the type of repair you are performing.

A core plug repair replaces damaged core material with a shaped piece of similar core material.

Complete removal of core material to the opposite face generally requires some handcutting with a core knife. Figure 3-15 shows core material being removed with a core knife.

The core can be peeled away from the skin bond using duckbill pliers. Sanding is then required to remove irregular accumulations of adhesive from the undamaged inner face.

Remove only enough adhesive to produce a smooth finish.

CAUTION: Care should be used when peeling core material from thin-skin sandwich face sheets, because the skin can be damaged by pulling on the core.

FIGURE 3-12. Cross section of bonded and bolted overlap repairs.

FIGURE 3-13. Honeycomb core removal.

FIGURE 3.14  Removing honeycomb core from a tapered control surface.

FIGURE 3-15. Removal of core with core knife.

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Par 3-11 Page 3-89 (and 3-14) A core plug should be fabricated as follows.

Select the core density. Cell size and ribbon direction and thickness should be at least the same as that used in the original construction.

Trim the sides of the plug to a loose fit in the routed cavity. Trim the plug height so the top of the plug sits .001 inch higher than the level of the original surface. The core will compress and set during cure thereby requiring the extra height. Carefully remove the trimmed core plug from the machined cavity. Use a vacuum cleaning device to remove any dust or particles remaining on the core plug or in the repair area. Clean the core plug by rinsing with an approved solvent and wrap the plug in a clean polyethylene bag until needed for assembly.

CAUTION: When handling film adhesives, prepreg fabrics, or parts with prepared surfaces, latex gloves must be worn.

Using a film adhesive, the core plug should be installed as follows. Select the appropriate adhesive film. Cut one disk of adhesive to the same shape and size as the perimeter of the repair cutout if the repair extends through the entire core thickness. Cut two disks if a partial depth core repair is being made. Cut one strip of core splice adhesive to wrap around the core perimeter to its full depth. For a partial depth core repair, also cut out a fiberglass or aluminum disk, again matching the size and the shape of the repair cutout. Figure 3-16 shows the details of a partial depth core repair. Preassemble the pieces.

Wipe the bottom and sides of the cutout area with solvent. Allow the area to dry. Insert the core plug assembly with splice adhesive applied to the perimeter into the core cutout.

Ensure the core plug ribbon direction matches that of the parent core. In the case of a partial depth core repair, the plug and disk

FIGURE 3-16. Details of core repair.

may be cured to save time. Some restraining method, such as vacuum bagging, may be desirable over the core splice adhesive as this material usually foams up and out during cure. Cure the adhesive according to the manufacturer’s directions and allow the area to cool.

Trim the top of the cured core plug flush with either the original core or the mold line, depending on the type of laminate repair to be performed. Proceed with laminate repair.

FINISHING. The type of finish coating applied to a metallic sandwich repair will normally be determined by the exposed material and the application of the part or assembly. Rain erosion of plastic parts, the need for electrical or dielectric properties, and/or the necessity for anti-corrosion coatings must be considered when the choice of finish is made. Plasticfaced parts such as radomes are finished primarily for rain erosion while aluminum- or other metal-faced laminates are

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Par 3-11 Page 3-90 (and 3-14) finished for corrosion protection. For coatings to perform their function properly, it is essential that they be applied to surfaces that are clean, free of voids, and smooth. The edges of all parts not protected by a bonding of aluminum or glass-fabric laminate must be sealed to reduce the rate of moisture absorption.

3-12.3-17. [RESERVED.]

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Par 3-18 Page 3-91

**SECTION 3. TRANSPARENT PLASTICS**

GENERAL. Plastics cover a broad field of organic synthetic resins and may be divided into two main classifications-thermoplastic and thermosetting plastics.

Thermoplastics. Thermoplastics may be softened by heat and can be dissolved in various organic solvents.

Two kinds of transparent th ermoplastic materials are commonly employed in windows, canopies, etc. These materials are known as acrylic plastics and cellulose acetate plastics.

Cellulose acetate was used in the past but since it is dimensionally unstable and turns yellow after it has been installed for a time, it has just about passed from the scene and is not considered an acceptable substitute for acrylic.

Acrylic plastics are known by the trade names of Lucite or Plexiglas and by the British as Perspex and meet the military specifications of MIL-P-5425 for regular acrylic, MIL-P-8184 for craze-resistant acrylic.

Thermosetting Plastics. These plastics do not soften appreciably under heat but may char and blister at temperatures of 240 to 260 °C (400 to 500 °F). Most of the molded products of synthetic resin composition, such as phenolic, urea-formaldehyde, and melamine formaldehyde resins, belong to the thermosetting group.

Once the plastic becomes hard, additional heat will not change it back into a liquid as it would with a thermoplastic.

STORAGE AND HANDLING. Because transparent thermoplastic sheets soften and deform when they are heated, they must be stored where the temperature will never become excessive. Store them in a cool, dry location away from heating coils, radiators, or steam pipes, and away from such fumes as are found in paint spray booths or paint storage areas.

Paper-masked transparent sheets must be kept out of the direct rays of the s un, because sunlight will accelerate deterioration of the adhesive, causing it to bond to the plastic and making it difficult to remove.

Plastic sheets should be stored with the masking paper in place, in bins that are tilted at a ten-degree angle from the vertical. This will prevent their buckling. If the sheets are stored horizontally, take care to avoid getting dirt and chips between them. Stacks of sheets must never be over 18 inches high, with the smallest sheets stacked on top of the larger ones so there will be no unsupported overhang. Leave the masking paper on the sheets as long as possible, and take care not to scratch or gouge the sheets by sliding them against each other or across rough or dirty tables.

Formed sections should be stored with ample support so they will not lose their shape. Vertical nesting should be avoided. Protect formed parts from temperatures higher than 120 °F (49 °C), and leave their protective coating in place until they are installed on the aircraft.

**FORMING PROCEDURES AND TECHNIQUES.**

Transparent acrylic plastics get soft and pliable when they are heated to their forming temperatures and can be formed to almost any shape. When they cool, they retain the shape to which they were formed. Acrylic plastic may be cold-bent into a single curvature if the material is thin and the bending radius is at least 180 times the thickness of the sheet.

Cold bending beyond these limits

### Page 156

Par 3-18 Page 3-92 will impose so much stress on the surface of the plastic that tiny fissures or cracks, called crazing, will form.

HEATING. Before heating any transparent plastic material, remove all of the masking paper and adhesive from the sheet. If the sheet is dusty or dirty, wash it with clean water and rinse it well. Dry the sheet thoroughly by blotting it with soft absorbent paper towels.

NOTE: Wear cotton gloves when handling the plastic to eliminate finger marks on the soft surface.

For the best results when hot-forming acrylics, use the temperatures recommended by the manufacturer. A forced-air oven should be used −one that is capable of operating over a temperature range of 120 to 374 °F (49 to 190 °C). If the part gets too hot during the forming process, bubbles may form on the surface and impair the optical qualities of the sheet.

For uniform heating, it is best to hang the sheets vertically by grasping them by their edges with spring clips and suspending the clips in a rack. (See figure 3-17.) If the piece is too small to hold with clips, or if there is not enough trim area, lay the sheets on shelves or racks covered with soft felt or flannel. Be sure there is enough open space to allow the air to circulate around the sheet and heat it evenly.

Small forming jobs, such as landing light covers, may be heated in a kitchen baking oven. Infrared heat lamps may be used if they are arranged on 7-or 8-inch centers and enough of them are used in a bank to heat the sheet evenly. Place the lamps about 18 inches from the material.

FIGURE 3-17. Hanging of acrylic sheets.

Never use hot water or steam directly on the plastic to heat it, because this will likely cause the acrylic to become milky or cloudy.

FORMS. Heated acrylic plastic will mold with almost no pressure, so the forms used can be of very simple construction. Forms made of pressed wood, plywood, or plaster are adequate to form simple curves, but reinforced plastic or plaster may be needed to shape complex or compound curves.

Since hot plastic conforms to any waviness or unevenness, the form used must be completely smooth.

To ensure this, sand the form and cover it with soft cloth such as outing flannel or billiard felt.

The mold should be large enough to extend beyond the trim line of the part, and provisions should be made for holding the hot plastic snug against the mold as it cools.

A mold can be made for a complex part by using the damaged part itself. If the part is broken, tape the pieces together, wax or grease the inside so the plaster will not stick to it, and support the entire part in sand.

Fill the part with plaster and allow it to harden, and

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Par 3-22 Page 3-93 then remove it from the mold. Smooth out any roughness and cover it with soft cloth. It is now ready to use to form the new part.

FORMING METHODS. (See table 3-1.)

Simple Curve Forming. Heat the plastic material to the recommended temperature, remove it from the heat source, and carefully drape it over the prepared form.

Carefully press the hot plastic to the form and either hold or clamp the sheet in place until it cools. This process may take from ten minutes to one-half hour.

Do not force-cool it.

Compound-Curve Forming. This type of forming is normally used for such parts as canopies or complex wingtip light covers, and it requires a great deal of specialized equipment. Th ere are four commonly used methods, each having its advantages and disadvantages.

Stretch Forming. Preheated acrylic sheets are stretched mechanically over the form in much the same way as is done with the simp le curved piece. Special care must be taken to preserve uniform thickness of the material, since some parts will have to stretch more than others.

Male And Female Die Forming. This requires expensive matching male and female dies. The heated plastic sheet is placed between the dies which are then mated. When the plastic cools, the dies are opened.

Vacuum-Forming Without Forms. Many aircraft canopies are formed by this method. In this process a clamp with an opening of the desired shape is placed over a vacuum box and the heated sheet of plastic is clamped in place. When the air in the box is evacuated, the outside air pressure will force the hot plastic through the opening and form the concave canopy. It is the surface tension of the plastic that shapes the canopy.

Vacuum-Forming With A Female Form. If the shape needed is other than that which would be formed by surface tension, a female mold, or form must be used. It is placed below the plastic sheet and the vacuum pump is connected. When air from the form is evacuated, the outside air pressure will force the hot plastic sheet into the mold and fill it.

Sawing And Drilling.

Several types of saws can be used with transparent plastics, however circular saws are the best for straight cuts. The blades TABLE 3-1. Typical temperatures for forming acrylic sheets.

Thickness of sheet (in.)

0.125

0.250

0.125

0.250

Type of forming

Regular acrylic plastic.

MIL-P-6886 Heat-resistant acrylic plastic, MIL-P-5425, and craze-resistant acrylic plastic, MIL-P-8184 °C °F °C °F °C °F °C °F Simple curve 113 235 110 230 135 275 135 275 Stretch forming (dry mold cover)

302 Male and female forming 140 284 135 275 180 356 170 338 Vacuum forming without form

293 Vacuum forming with female form

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Par 3-23 Page 3-94 should be hollow -ground or have some set to prevent binding. After the teeth are set, they should be side-dressed to produce a smooth edge on the cut. Band saws are recommended for cutting flat acrylic sheets when the cuts must be curved or where the sheet is cut to a rough dimension to be trimmed later. Close control of size and shape may be obtained by band sawing a piece to within 1/16 inch of the desired size, as marked by a scribed line on the plastic, and then sanding it to the correct size with a drum or belt sander.

Unlike soft metal, acrylic plastic is a very poor conductor of heat. Make provisions for removing the heat when drilling. Deep holes need cooling, and a water-soluble cutting oil is a satisfactory coolant since it has no tendency to attack the plastic.

The drill used on acrylics must be carefully ground and free from nicks and burrs that would affect the surface finish. Grind the drill with a greater included angle than would be used for soft metal. The rake angle should be zero in order to scrape, not cut. (See figure 3-18.)

FIGURE 3-18. Drill having an included angle of approximately 150°, used to drill acrylic plastics.

The patented Unibit (see figure 3 -19) is good for drilling small holes in aircraft windshields and windows. It can cut holes from 1/8-to 1/2 −inch in 1/32 −inch increments and produces good smooth holes with no stress cracks around their edges.

FIGURE 3-19. Unibit drill for drilling acrylic plastics.

Polymerizable Cements. Polymerizable cements are those in which a catalyst is added to an already thick monomerpolymer syrup to promote rapid hardening.

Cement PS-30 and Weld -On 40 are polymerizable cements of this type. Th ey are suitable for cementing all types of PLEXIGLAS acrylic cast sheet and parts molded from PLEXIGLAS molding pellets. At room temperature, the cements ha rden (polymerize) in the container in about 45 minutes after mixing the components. They will hard en more rapidly at higher temperatures. The cement joints are usually hard enough for handling within 4 hours after assembly. The joints may be machined within 4 hours after assembly, but it is better to wait 24 hours.

PS-30 and Weld -On 40 joints retain excellent appearance and color stability after outdoor exposure. These cements produce clear, transparent joints and should be used when the color and appearance of the joints are important.

PS-30 and Weld -On 40 should be used at temperatures no lower than 65 °F. If cementing is done in a room cooler than 65 °F, it will require a longer time to harden and the joint strength will be reduced.

The cement should be prepared with the correct proportions of components as given in the manufacturer’s instructions and thoroughly mixed, making sure neither the mixing container nor mixing paddle adds color or affects the hardening of the cement.

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Par 3-23 Page 3-95 Clean glass or polyethylene mixing containers are preferred.

Because of their short pot life (approximately 45 minutes) Cement PS-30 and Weld-On 40 must be used quickly once the components are mixed. Time consumed in preparation shortens the effective working time, making it necessary to have everything ready to be cemented before the cements are mixed. For better handling pour cement within 20 minutes of mixing.

For maximum joint strength, the final cement joint should be free of bubbles. It will usually be sufficient to allow the mixed cement to stand for 10 minutes before cementing to allow bubbles to rise to the surface. (See figure 3-20.)

FIGURE 3-20. Applying pressure to acrylic plastics.

The gap joint technique can only be used with colorless PLEXIGLAS acrylic or in cases in which joints will be hidden. If inconspicuous joints in colored PLEXIGLAS acrylic are needed, the parts must be fitted closely, using closed V groove, butt, or arc joints.

Cement forms or dams may be made with masking tape as long as the adhesive surface does not contact the cement. This is easily done with a strip of cellophane tape placed over the masking tape adhesive. The tape must be chosen carefully. The adhesive on ordinary cellophane tape prevents the cure of PS-30 and Weld-On 40. Before actual fabrication of parts, sample joints should be tried to ensure that the tape system used will not harm the cement. Since it is important for all of the cement to remain in the gap, only contact pressure should be used.

Bubbles will tend to float to the top of the cement bead in a gap joint after the cement is poured. These cause no problem if the bead is machined off. A small wire (not copper), or similar objects may be used to lift some bubbles out of the joint; however, the cement joint should be disturbed as little as possible.

Polymerizable cements shrink as the cement hardens. Therefore, the freshly poured cement bead should be left above the surfaces being cemented to compensate for the shrinkage. If it is necessary for appearances, the bead may be machined off after the cement has set.

REPAIR OF PLASTICS. Replace, rather than repair extensively damaged transparent plastic, whenever possible, since even a carefully patched part is not the equal of a new section, either optically or structurally. At the first sign of crack development, drill a small hole with a # 30 or a 1/8-inch drill at the extreme ends of the cracks as shown in figure 3-21. This serves to localize the cracks and to prevent further splitting by distributing the strain over a large area. If the cracks are small, stopping them with drilled holes will usually suffice until replacement or more permanent repairs can be made. The following repairs are permissible;

however, they are not to be located in the pilot’s line of vision during landing or normal flight.

### Page 160 (OCR)

**ALL THE STRAINS WHICH ORIGINALLY CAUSED**

THE CRACK ARE CONCENTRATED AT POINT A - TENDING TO EXTEND THE CRACK. THEREFORE DRILL, WITH A #30 OR 1/8" DRILL BIT, A SMALL HOLE AT END OF THE CRACK POINT A1 TO DISTRIBUTE THE STRAIN OVER A WIDER AREA.

**EACH CRACK OCCURRING AT ANY HOLE OR TEAR IS DRILLED IN SAME MANNER.**

FIGURE 3-21. Stop-drilling cracks.

Surface Patch. If a surface patch is to be installed, trim away the damaged area and round all corners. Cut a RFACE PATCH piece of plastic of sufficient size to cover the damaged area and extend at least 3/4 inch on each side of the STOI crack or hole. Bevel the edges as shown in figure 3-22. CRÁCK DRILLEC If the section to be repaired is curved, shape the patch to the same contour by heating it in an oil bath at a temperature of 248 to 302 °F, or it may be heated on a hot-plate until soft. Boiling water should not be used SURFACE PATCH FOR CRACKS BEVELED EDGE for heating. Coat the patch evenly with plastic solvent adhesive and immediately place it over the hole. Maintain a uniform pressure of 5 to 10 psi on the patch TRAM AL CORNERND SURFACE PATCH for a minimum of 3 hours. Allow the patch to dry CORNE SHARP 24 to 36 hours before sanding or polishing → BEVELED Plug Patch. When using inserted patches to repair SURFACE PATCH FOR IRREGULAR SHAPED DAMAGE / "EDGE holes in plastic structures, trim the holes to a perfect circle or oval and bevel the edges slightly. Make the patch slightly thicker than the material being ACE PATCH ROUND HOLE SÉVELED EDGE

SURFACE PATCH FOR ROUND HOLES FIGURE 3-22. Surface patches.

Par 3-23 Page 3-96

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Par 3-24 Page 3-97 repaired and similarly bevel its edges. Install patches in accordance with figure 3-23. Heat the plug until soft and press it into the hole without cement and allow to cool to make a perfect fit. Remove the plug, coat the edges with adhesive, and then reinsert in the hole.

Maintain a firm light pressure until the cement has set.

Sand or file the edges level with the surface, then buff and polish.

FIGURE 3-23. Plug patch repair.

CLEANING AND POLISHING TRANSPARENT PLASTIC. Plastics have many advantages over glass for aircraft use, but they lack the surface hardness of glass and care must be exercised while servicing the aircraft to avoid scratching or otherwise damaging the surface.

Clean the plastic by washing it with plenty of water and mild soap, using a cl ean, soft, grit-free cloth, sponge, or bare hands. Do not use gasoline, alcohol, benzene, acetone, carbon tetrachloride, fire extinguisher or deicing fluids, lacquer thinners, or window cleaning sprays. These will soften the plastic and cause crazing.

Plastics should not be rubbed with a dry cloth since this is likely to cause scratches, and also to build up an electrostatic charge that 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 coat and bring to a high polish by rubbing lightly with a soft cloth.

Do not attempt hand polishing or buffing until the surface is clean. A soft, ope n-type cotton or flannel buffing wheel is suggested. Minor scratches may be removed by vigorously rubbing the affected area by hand, using a soft clean cloth dampened with a mixture of turpentine and chalk, or by applying automobile cleanser with a damp cloth. Remove the cleaner and polish with a soft, dry cloth. Acrylic and cellulose acetate plastics are thermoplastic. Friction created by buffing or polishing too long in one spot can generate sufficient heat to soften the surface. This condition will produce visual distortion and should be avoided.

REPLACEMENT PANELS. Use material equivalent to that originally used by the manufacturer of the aircraft for replacement panels. There are many types of transparent plastics on the market. Their properties vary greatly, particularly in regard to expansion characteristics, brittleness under low

### Page 162

Par 3-40 Page 3-98 temperatures, resistance to discolora tion when exposed to sunlight, surface checking, etc. Information on these properties is in MIL-HDBK-17A, Plastics for Flight Vehicles, Part II Transparent Glazing Materials, available from the Government Printing Office (GPO). These properties are considered by aircraft manufacturers in selecting materials to be used in their designs and the use of substitutes having different characteristics may result in subsequent difficulties.

INSTALLATION PROCEDURES. When installing a replacement panel, use the same mounting method employed by the manufacturer of the aircraft. While the actual installation will vary from one type of aircraft to another, consider the following major principles when installing any replacement panel.

Never force a plastic panel out of shape to make it fit a frame. If a replacement panel does not fit easily into the mounting, obtain a new replacement or heat the whole panel and reform. When possible, cut and fit a new panel at ordinary room temperature.

In clamping or bolting plastic panels into their mountings, do not place the plastic under excessive compressive stress. It is easy to develop more than 1,000 psi on the pl astic by over-torquing a nut and bolt. Tighten each nut to a firm fit, then back the nut off one full turn (until they are snug and can still be rotated with the fingers).

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Par 3-40 Page 3-99 In bolted installations, use spacers, collars, shoulders, or stop-nuts to prevent tightening the bolt excessively.

Whenever such devices are used by th e aircraft manufacturer, retain them in the replacement installation. It is important that the original number of bo lts, complete with washers, spacers, etc., be used. When rivets are used, provide adequate spacers or other satisfactory means to prevent excessive tightening of the frame to the plastic.

Mount plastic panels between rubber, cork, or other gasket material to make the installation waterproof, to reduce vibration, and to help to distribute compressive stresses on the plastic.

Plastics expand and contract considerably more than the metal channels in which they are mounted. Mount windshield panels to a sufficient depth in the channel to prevent it from falling out when the panel contracts at low temperatures or deforms under load. When the manufacturer’s original design permits, mount panels to a minimum depth of 1-1/8 inch, and with a clearance of 1/8 inch between the plastic and bottom of the channel.

In installations involving bolts or rivets, make the holes through the plastic oversize by 1/8-inch diameter and center so that the plastic will not bind or crack at the edge of the holes. The use of slotted holes is also recommended.

3-28.3-39. [RESERVED.]

### Page 164

Par 3-40 Page 3-100 SECTION 4. WINDSHIELDS, ENCLOSURES, AND WINDOWS

GENERAL. These repairs are applicable to plastic windshields, enclosures, and windows in nonpressurized airplanes. For pressurized airplanes, replace or repair plastic elements in accordance with the manufacturer’s recommendation. When windshields and side windows made of acrylic plastics are damaged, they are usually replaced unless the damage is minor and a repair would not be in the line of vision. Repairs usually require a great deal of labor. Replacement parts are readily available, so replacement is normally more economical than repair.

Minor Repairs. There are times, however, when a windshield may be cracked and safety is not impaired.

In that case, repairs can be made by stop-drilling the ends of the crack with a # 30 drill (1/8 inch) to prevent the concentration of stre sses causing the crack to continue. Drill a series of number 40 holes a half-inch from the edge of the crack about a half-inch apart, and lace through these holes with brass safety wire (see figure 3-24) and seal with clear silicone to waterproof.

Temporary Repairs. One way to make a temporary repair is to stop-drill the ends of the crack, and then drill number 27 holes every inch or so in the crack. Use AN515-6 screws and AN365-632 nuts with AN960-6 washers on both sides of the plastic. This will hold the crack together a nd prevent further breakage until the windshield can be properly repaired or replaced. (See figure 3-24.)

Permanent Repairs. Windshields or side windows with small cracks that affect only the appearance rather than the airworthiness of a sheet, may be repaired by first stop-drilling the ends of the crack with a # 30 or a 1/8-inch drill. Then us e a hypodermic syringe and needle to fill the crack with polymerizable cement such as PS-30 or Weld-On 40, and allow capillary action to fill the crack completely.

Soak the end of a 1/8-inch acrylic rod in cement to form a cushion and insert it in the stop-drilled hole. Allow the repair to dry for about 30 minutes, and then trim the rod off flush with the sheet.

Polishing and Finishing. Scratches and repair marks, within certain limitations, can be removed from acrylic plastic. No sanding that could adversely affect the plastic’s optical properties and distort the pilot’s vision should be done on any portion of a windshield.

If there are scratches or repair marks in an area that can be sanded, they may be removed by first sanding the area. Use 320- or 400-grit abrasive paper that is wrapped around a felt or rubber pad.

Use circular rubbing motions, light pressure, and a mild liquid soap solution as a lubricant.

After the sanding is complete, rinse the surface thoroughly with running water. Then, using a 500-grit paper, continue to sand lightly. Keep moving to higher grit paper and sand and rinse until all of the sanding or repair marks have been removed.

After using the finest abrasive paper, use rubbing compound and buff in a circular motion to remove all traces of the sanding.

Cleaning. Acrylic windshields and windows may be cleaned by washing them with mild soap and running water. Rub the surface with your bare hands in a stream of water. Follow with the same procedure but with soap and water. After the soap and dirt have been flushed away, dry the surface with a soft, clean cloth or tissue and polish it with a

### Page 165

Par 3-40 Page 3-101

FIGURE 3-24. Temporary repairs to cracked windshields or windows.

windshield cleaner especially approved for use on aircraft transparent plastic s. These cleaners may be purchased through aircraft supply houses.

Waxing. A thin coating of wax will fill any minute scratches that may be present and will cause rain to form droplets that are easily blown away by the wind.

PROTECTION. Acrylic windshields are often called “lifetime” windshields, to distinguish them from those made of the much shorter-lived acetate material. However, even acrylic must be protected from the ravages of the elements.

When an aircraft is parked in direct sunlight, the windshield will absorb heat and will actually become hotter than either the inside of the aircraft or the outside air. The sun will cause the inside of a closed aircraft to become extremely hot, and this heat is also absorbed by the plastic windshield.

To protect against this damage, it is wise to keep the aircraft in a hangar. If this is not possible, some type of shade should be provided to keep the sun from coming in direct contact with the windshield. Some aircraft owners use a close-fitting, opaque, reflective cover over the windshield. In many cases, this has done more harm than good. This cover may absorb moisture from the air and give off harmful vapors, and if it touches the surface of the pl astic it can cause crazing or minute cracks to form in the windshield. Another hazard in using such a cover is that sand can blow up under the cover and scratch the plastic.

WINDSHIELD INSTALLATION. Aircraft windshields may be purchased either from the original aircraft manufacturer or from any of several FAA-PMA sources. These windshields are formed to the exact shape required, but are slightly larger than necessary so they may be trimmed to the exact size.

After removing the damaged windshield, clean all of the sealer from the grooves and cut the new windshield to fit. New windshields are covered with either protective paper or film to prevent damage during handling or installation. Carefully peel back just enough of this covering to make the installation. The windshield must fit in its channels with about 1/8- to 1/4-inch clearance to allow for expansion and contraction. If any holes are drilled in the plastic for screws, they should be about 1/8 inch oversize.

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Par 3-42 Page 3-102 (and 3-26) Place the sealing tape around the edges of the windshield and install the windshield in its frame.

Screws that go through the windshield should be tightened down snug and then backed out a full turn, so the plastic can shift as it expands and contracts.

Do not remove the protective paper or film until the windshield is installed and all of the securing screws are in place.

3-43.3-47. [RESERVED.]

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Par 4-1 Page 4-103

**CHAPTER 4. METAL STRUCTURE, WELDING, AND BRAZING**

**SECTION 1. IDENTIFICATION OF METALS**

GENERAL. Proper identification of the aircraft structural material is the first step in ensuring that the continuing airworthiness of the aircraft will not be degraded by making an improper repair using the wrong materials.

Ferrous (iron) alloy materials are generally classified according to carbon content. (See table 4-1.)

TABLE 4-1. Ferrous (iron) alloy materials.

MATERIALS CARBON CONTENT Wrought iron Trace to 0.08% Low carbon steel 0.08% to 0.30% Medium carbon steel 0.30% to 0.60% High carbon steel 0.60% to 2.2% Cast iron 2.3% to 4.5%

The strength and ductility, or toughness of steel, is controlled by the kind and quantity of alloys used and also by cold-working or heat-treating processes used in manufacturing. In general, any process that increases the strength of a material will also decrease its ductility.

IDENTIFICATION OF STEEL STOCK. The Society of Automotive Engineers (SAE) and the American Iron and Steel Institute (AISI) use a numerical index system to identify the composition of various steels. The numbers assigned in the combined listing of standard steels issued by these groups represent the type of steel and make it possible to readily identify the principal elements in the material.

The basic numbers for the four digit series of the carbon and alloy steel may be found in table 4-2. The first digit of the four number designation indicates the type to which the steel belongs. Thus, “1” indicates a carbon steel, “2” a nickel steel, “3” a nickel chromium steel, etc. In the case of simple alloy steels, the second digit indicates the approximate percentage of the predominant alloying element. The last two digits usually indicate the mean of the range of carbon content. Thus, the designation “1020” indicates a plain carbon steel lacking a principal alloying element and containing an average of 0.20 percent (0.18 to 0.23) carbon. The designation “2330” indicates a nickel steel of approximately 3 percent (3.25 to 3.75) nickel and an average of 0.30 percent, (0.28 to 0.33) carbon content.

The designation “4130” indicates a chromiummolybdenum steel of approximately 1 percent

(0.80 to 1.10) chromium, 0.20 percent (0.15 to 0.25) molybdenum, and 0.30 percent (0.28 to 0.33) carbon.

There are numerous steels with higher percentages of alloying elements that do not fit into this numbering system. These include a large group of stainless and heat resisting alloys in which chromium is an essential alloying element. Some of these alloys are identified by three digit AISI numbers and many others by designations assigned by the steel company that produces them. The few examples in table 4-3 will serve to illustrate the kinds of designations used and the general alloy content of these steels.

“1025” welded tubing as per Specification MIL-T-5066 and “1025” seamless tubing conforming to Specification MIL-T-5066A are interchangeable.

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Par 4-1 Page 4-104

**INTERCHANGEABILITY OF STEEL TUBING .**

“4130” welded tubing conforming to Specification MIL-T-6731, and “4130” seam-less tubing conforming to Specification MIL-T-6736 are interchangeable.

NE-8630 welded tubing conforming to Specification MIL-T-6734, and NE-8630 seamless tubing conforming to Specification MIL-T-6732 are interchangeable.

IDENTIFICATION OF ALUMINUM. To provide a visual means for identifying the various grades of aluminum and aluminum alloys, such metals are usually marked with symbols such as a Government Specification Number, the temper or condition furnished, or the commercial code marking. Plate and sheet are usually marked with specification numbers or code markings in rows approximately 5 inches apart. Tubes, bars, rods, and extruded shapes are marked with specification numbers or code markings at intervals of 3 to 5 feet along the length of each piece.

The commercial code marking consists of a number which identifies the particular composition of the alloy. In addition, letter suffixes (see table 4-4) designate the basic temper designations and subdivisions of aluminum alloys.

TABLE 4-2. Numerical system for steel identification.

**TYPES OF STEELS NUMERALS**

AND DIGITS Plain carbon steel 10XX Carbon steel with additional sulfur for easy machining.

11XX Carbon steel with about 1.75% manganese 13XX .25% molybdenum. 40XX 1% chromium, .25% molybdenum 41XX 2% nickel, 1% chromium, .25% molybdenum 43XX

1.7% nickel, .2% molybdenum 46XX

3.5% nickel, .25% molybdenum 48XX 1% chromium steels 51XX 1% chromium, 1.00% carbon 51XXX

1.5% chromium steels 52XX

1.5% chromium, 1.00% carbon 52XXX 1% chromium steel with .15% vanadium 61XX .5% chromium, .5% nickel, .20% molybdenum 86XX .5% chromium, .5% nickel, .25% molybdenum 87XX 2% silicon steels, .85% manganese 92XX

3.25% nickel, 1.20% chromium, .12% molybdenum 93XX

TABLE 4-3. Examples of stainless and heat-resistant steels nominal composition (percent)

ALLOY DESIGNATION CARBON CHROMIUM NICKEL OTHER GENERAL CLASS OF STEEL 302 0.15 18 9 Austenitic 310 0.25 25 20 Austenitic 321 0.08 18 11 Titanium Austenitic 347 0.08 18 11 Columbium or Tantalum Austenitic 410 0.15 12.5 Martensitic, Magnetic 430 0.12 17 Ferritic, Magnetic 446 0.20 25 Nitrogen Ferritic, Magnetic PH15-7 Mo 0.09 15 7 Molybdenum, Aluminum Precipitation Hardening 17-4 PH 0.07 16.5 4 Copper, Columbium or Tantalum Precipitation Hardening

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Par 4-4 Page 4-105 (and 4-4) TABLE 4-4. Basic temper designations and subdivisions from aluminum alloys.

NON HEAT-TREATABLE ALLOYS HEAT-TREATABLE ALLOYS Temper Designation

Definition Temper Designation

Definition -0 Annealed recrystallized (wrought products only) applies to softest temper of wrought products.

-0 Annealed recrystallized (wrought products only) applies to softest temper of wrought products.

-H1 Strain-hardened only. Applies to products which are strain-hardened to obtain the desired strength without supplementary thermal treatment.

-T1 Cooled from an elevated temperature shaping process (such as extrusion or casting) and naturally aged to a substantially stable condition.

-H12 Strain-hardened one-quarter-hard tem per. -T2 Annealed (castings only).

-H14 Strain-hardened half-hard temper. -T3 Solution heat-treated and cold-worked by the flattening or straightening operation.

-H16 Strain-hardened three-quarters-hard temper. -T36 Solution heat-treated and cold-worked by reduction of 6 percent -H18 Strain-hardened full-hard temper. -T4 Solution heat-treated.

-H2 Strain-hardened and then partially annealed.

Applies to products which are strain-hardened more than the desired final amount and then reduced in strength to the desired level by partial annealing.

-T42 Solution heat-treated by the user regardless of prior temper (applicable only to 2014 and 2024 alloys).

-H22 Strain-hardened and partially annealed to one-quarter-hard temper.

-T5 Artificially aged only (castings only).

-H24 Strain-hardened and partially annealed to half-hard temper.

-T6 Solution heat-treated and artificially aged.

-H26 Strain-hardened and partially annealed to three-quarters-hard temper.

-T62 Solution heat-treated and aged by user regardless of prior temper (applicable only to 2014 and 2024 alloys).

-H28 Strain-hardened and partially annealed to full-hard temper.

-T351, -T451, -T3510, -T3511, -T4510, -T4511.

Solution heat-treated and stress relieved by stretching to produce a permanent set of 1 to 3 percent, depending on the product.

-H3 Strain-hardened and then stabilized. Applies to products which are strain-hardened and then stabilized by a low temperature heating to slightly lower their strength and increase ductility.

-T651, -T851, -T6510, -T8510, -T6511, -T8511.

Solution heat-treated, stress relieved by stretching to produce a permanent set of 1 to 3 percent, and artificially aged.

-H32 Strain-hardened and then stabilized. Final temper is one-quarter hard.

-T652 Solution heat-treated, compressed to produce a permanent set and then artificially aged.

-H34 Strain-hardened and then stabilized. Final temper is one-half hard.

-T8 Solution heat-treated, cold-worked and then artificially aged.

-H36 Strain-hardened and then stabilized. Final temper is three-quarters hard.

-T/4 Solution heat-treated, cold-worked by the flattening or straightening operation, and then artificially aged.

-H38 Strain-hardened and then stabilized. Final temper is full-hard.

-T86 Solution heat-treated, cold-worked by reduction of 6 percent, and then artificially aged.

-H112 As fabricated; with specified mechanical property limits.

-T9 Solution heat-treated, artificially aged and then cold-worked.

-F For wrought alloys; as fabricated. No mechanical properties limits. For cast alloys; as cast.

-T10 Cooled from an elevated temperature shaping process artificially aged and then cold-worked.

-F For wrought alloys; as fabricated. No mechanical properties limits. For cast alloys; as cast.

4-5.4-15. [RESERVED.]

### Page 170

Par 4-16 Page 4-106

**SECTION 2. TESTING OF METALS**

HARDNESS TESTING. If the material type is known, a hardness test is a simple way to verify that the part has been properly heattreated. Hardness testers such as Rockwell, Brinell, and Vickers can be useful to check metals for loss of strength due to exposure to fire or abusive heating. Also, under-strength bolts can be found and removed from the replacement part inventory by checking the hardness of the bolt across the hex flats.

Although hardness tests are generally considered nondestructive, hardness testing does leave a small pit in the surface; therefore, hardness tests should not be used on sealing surfaces, fatigue critical parts, load bearing areas, etc., that will be returned to service.

These hardness tests provide a convenient means for determining, within reasonable limits, the tensile strength of steel. It has several limitations in that it is not suitable for very soft or very hard steels. Hardness testing of aluminum alloys should be limited to distinguishing between annealed and heattreated material of the same aluminum alloy. In hardness testing, the thickness and the edge distance of the specimen being tested are two factors that must be considered to avoid distortion of the metal. Several readings should be taken and the results averaged. In general, the higher the tensile strength, the greater its hardness. Common methods of hardness testing are outlined in the following paragraphs. These tests are suitable for determining the tensile properties resulting from the heat treatment of steel. Care should be taken to have case-hardened, corroded, pitted, decarburized, or otherwise nonuniform surfaces removed to a sufficient depth.

Exercise caution not to cold-work, and consequently harden, the steel during removal of the surface.

ROCKWELL HARDNESS TEST. The Rockwell hardness test is the most common method for determining hardness of ferrous and many nonferrous metals. (See table 4-5.) It differs from Brinell hardness testing in that the hardness is determined by the depth of indentation made by a constant load impressing on an indenter. In this test, a standard minor load is applied to set a hardened steel ball or a diamond cone in the surface of the metal, followed by the application of a standard major load. The hardness is measured by depth of penetration.

Rockwell superficial hardness tests are made using light minor and major loads and a more sensitive system for measuring depth of indentation. It is useful for thin sections, very small parts, etc. Calibration of Rockwell hardness testers is done in accordance with American Society of Testing Materials (ASTM E-18) specifications.

BRINELL HARDNESS TEST. In this test a standard constant load, usually 500 to 3,000 kg, is applied to a smooth flat metal surface by a hardened steel-ball type indenter, 10 mm in diameter. The 500-kg load is usually used for testing nonferrous metals such as copper and aluminum alloys, whereas the 3,000-kg load is most often used for testing harder metals such as steels and cast irons. The numerical value of Brinell Hardness (HB), is equal to the load, divided by the surface area of the resulting spherical impression.

Where P is the load, in kg; D is the diameter of the ball, in mm; and d is the diameter of the indentation, in mm.

### Page 171

Par 4-20 Page 4-107 General Precautions. To avoid misapplication of Brinell hardness testing, the fundamentals and limitations of the test procedure must be clearly understood. To avoid inaccuracies, the following rules should be followed.

Do not make indentations on a curved surface having a radius of less than 1 inch.

Do make the indentations with the correct spacing. Indentations should not be made too close to the edge of the work piece being tested.

Apply the load steadily to avoid overloading caused by inertia of the weights.

Apply the load so the direction of loading and the test surface are perpendicular to each other within 2 degrees.

The thickness of the work piece being tested should be such that no bulge or mark showing the effect of the load appears on the side of the work piece opposite the indentation.

The indentation diameter should be clearly outlined.

Limitations. The Brinell hardness test has three principal limitations.

The work piece must be capable of accommodating the relatively large indentations.

Due to the relatively large indentations, the work piece should not be used after testing.

The limit of hardness, 15 HB with the 500-kg load to 627 HB with the 3,000-kg load, is generally considered the practical range.

Calibration. A Brinell Hardness Tester should be calibrated to meet ASTM standard E10 specifications.

VICKERS HARDNESS TEST. In this test, a small pyramidal diamond is pressed into the metal being tested. The Vickers Hardness number (HV) is the ratio of the load applied to the surface area of the indention. This is done with the following formula.

The indenter is made of diamond, and is in the form of a square-based pyramid having an angle of 136 degrees between faces. The facets are highly-polished, free from surface imperfections, and the point is sharp. The loads applied vary from 1 to 120 kg; the standard loads are 5, 10, 20, 30, 50, 100, and 120 kg. For most hardness testing, 50 kg is maximum.

A Vickers hardness tester should be calibrated to meet ASTM standard E10 specifi cations, acceptable for use over a loading range.

MICROHARDNESS TESTING. This is an indentation hardness test made with loads not exceeding 1 kg (1,000 g). Such hardness tests have been made with a load as light as 1 g, although the majority of microhardness tests are made with loads of 100 to 500 g. In general, the term is related to the size of the indentation rather than to the load applied.

### Page 172

Par 4-20 Page 4-108 Fields of Application . Microhardness testing is capable of providing information regarding the hardness characteristics of materials which cannot be obtained by hardness tests such as the Br inell or Rockwell, and are as follows.

Measuring the hardness of precision work pieces that are too small to be measured by the more common hardness-testing methods.

Measuring the hardness of product forms such as foil or wire that are too thin or too small in diameter to be measured by the more conventional methods.

Monitoring of carburizing or nitriding operations, which is sometimes accomplished by hardness surveys taken on cross sections of test pieces that accompanied the work pieces through production operations.

Measuring the hardness of individual microconstituents.

Measuring the hardness close to edges, thus detecting undesirable surface conditions such as grinding burn and decarburization.

Measuring the hardness of surface layers such as plating or bonded layers.

Indenters. Microhardness testing can be performed with either the Knoop or the Vickers indenter. The Knoop indenter is used mostly in the United States; the Vickers indenter is the more widely used in Europe.

Knoop indentation testing is performed with a diamond, ground to pyramidal form, that produces a diamond-shaped indentation with an approximate ratio between long and short diagonals of 7 to 1. The indentation depth is about one-thirtieth of its length. Due to the shape of the indenter, indentations of accurately measurable length are obtained with light loads.

The Knoop hardness number (HK) is the ratio of the load applied to the indenter to the unrecovered projected area of indentation.

The formula for this follows.

H K=P/A =P/ Cl2

Where P is the applied load, in kg; A is the unrecovered projected area of indentation, in square mm; l is the measured length of the long diagonal, in mm; and C is 0.07028, a constant of the indenter relating projected area of the indentation to the square of the length of the long diagonal.

INDENTATIONS. The Vickers indenter penetrates about twice as far into the work piece as does the Knoop indenter. The diagonal of the Vickers indentation is about one-third of the total length of the Knoop indentation. The Vickers indenter is less sensitive to minute differences in surface conditions than is the Knoop indenter.

However, the Vickers indentation, because of the shorter diagonal, is more sensitive to errors in measuring than is the Knoop indentation. (See figure 4-1.)

FIGURE 4-1. Comparison of indentation made by Knoop and Vickers indenters in the same work metal and at the same loads.

### Page 173

Par 4-36 Page 4-109 MAGNETIC TESTING. Magnetic testing consists of determining whether the specimen is attracted by a magnet. Usually, a metal attracted by a magnet is iron, steel, or an iron-base alloy containing nickel, cobalt, or chromium. However, there are exceptions to this rule since some nickel and cobalt alloys may be either magnetic or nonmagnetic. Never use this test as a final basis for identification.

The strongly attracted metals could be pure iron, pure nickel, cobalt, or iron-nickel-cobalt alloys. The lightly attracted metals could be cold-worked stainless steel, or monel. The nonmagnetic metals could be aluminum, magnesium, silver, or copper-base alloy, or an annealed 300-type stainless steel.

ALUMINUM TESTING. Hardness tests are useful for testing aluminum alloy chiefly as a means of distinguishing between annealed, cold-worked, heat-treated, and heat-treated and aged material. It is of little value in indicating the strength or quality of heat treatment. Typical hardness values for aluminum alloys are shown in table 4-5.

Clad aluminum alloys have surface layers of pure aluminum or corro sion-resistant aluminum alloy bonded to the core material to inhibit corrosion. Presence of such a coating may be determined under a magnifying glass by examination of the edge surface which will show three distinct layers. In aluminum alloys, the properties of an y specific alloy can be altered by work hardening (often called strain-hardening), heat treatment, or by a combination of these processes.

Test for distinguishing heat-treatabl e and nonheat-treatable aluminum alloys . If for any reason the identification mark of the alloy is not on the material, it is possible to distinguish between some heat-treatable alloys

### Page 174

Par 4-36 Page 4-110 and some nonheat-treatable alloys by immersing a sample of the material in a 10 percent solution of caustic soda (sodium hydroxide). Those h eat-treated alloys containing several per cent of copper (2014, 2017, and 2024) will turn black due to the copper content. High- copper alloys when clad will not turn black on the surface, but the edges will turn black at the center of the sheet where the core is e xposed. If the alloy does not turn black in the caustic soda solution it is not evidence that the alloy is nonheat-treatable, as various high-strength heat-treatable alloys are not based primarily on the use of copper as an alloying agent. These include among others 6053, 6061, and 7075 alloys. The composition and heat-treating ability of alloys which do not turn black in a caustic soda solution can be established only by chemical or spectro-analysis.

TABLE 4-5. Hardness values for aluminum alloys. (Reference MIL-H-6088G .)

Material

Hardness

Brinell number Commercial Temper 500 kg. load Designation 10 mm. ball

1100 0 23 H18 44 3003 0 28 H16 47 2014 0 45 T6 135 2017 0 45 T6 105 2024 0 47 T4 120 2025 T6 110 6151 T6 100 5052 0 47 H36 73 6061 0 30 T4 65 T6 95 7075 T6 135 7079 T6 135 195 T6 75 220 T4 75 C355 T6 80 A356 T6 70

4-24.4-35. [RESERVED.]

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Par 4-36 Page 4-111 SECTION 3. PRECAUTIONARY MEASURES

FLUTTER AND VIBRATION PRECAUTIONS. To prevent the occurrence of severe vibration or flutter of flight control surfaces during flight, precautions must be taken to stay within the design balance limitations when performing maintenance or repair.

Balance Changes. The importance of retaining the proper balance and rigidity of aircraft control surfaces cannot be overemphasized. The effect of repair or weight change on the balance and center of gravity is proportionately greater on lighter surfaces than on the older heavier designs. As a general rule, repair the control surface in such a manner that the weight distribution is not affected in any way, in order to preclude the occurrence of fl utter of the control surface in flight. Under certain conditions, counter-balance weight is added forward of the hinge line to maintain balance. Add or remove balance weights only when necessary in accordance w ith the manufacturer’s instructions. Flight tes ting must be accomplished to ensure flutter is not a problem. Failure to check and retain control surface balan ce within the original or maximum allowable value could result in a serious flight hazard.

Painting and Refinishing. Special emphasis is directed to the effect of too many extra coats of paint on balanced control surfaces. Mechanics must avoid adding additional coats of paint in excess of what the manufacturer originally applied. If available consult the aircraft manufacturer’s instru ctions relative to finishing and balance of control surfaces.

Trapped Water or Ice. Instances of flutter have occurred from unbalanced conditions caused by the collection of water or ice with in the surface. Therefore, ventilation and drainage provisions must be checked and retained when maintenance is being done.

Trim Tab Maintenance. Loose or vibrating trim tabs will increase wear of actuating mechanisms and hinge points which may develop into serious flutter conditions. When this happens, primary control surfaces are highly susceptible to wear, deformation, and fatigue failures because of the buffeting nature of the airflow over the tab mechanism. Trailing-edge play of the tab may increase, creating an unsafe flutter condition. Careful inspection of the tab and its mechanism should be conducted during overhaul and annual inspection periods. Compared to other flight control systems on the aircraft, only a minor amount of tab-mechanism wear can be tolerated.

Free play and stiffness may best be measured by a simple static test where “upward” and “downward” (or “leftward” and “rightward”) point forces are applied near the trailing edge of the tab at the span-wise attachment of the actuator (so as not to twist the tab). The control surface to which the trim tab is attached should be locked in place.

Rotational deflection readings are then taken near the tab trailing edge using an appropriate measuring device, such as a dial gauge.

Several deflection readings should be taken using loads first applied in one direction, then in the opposite. If the tab span does not exceed 35 percent of the span of the supporting control surface, the total free play at the tab trailing edge should not exceed 2 percent of the tab chord. If the tab span equals or exceeds 35 percent of the span of the supporting control surface, the total free play at the tab trailing edge should not exceed 1 percent of the distance from the tab hinge line to the trailing edge of the tab perpendicular to the tab hinge line. For example, a tab that has a chord of

### Page 176

Par 4-50 Page 4-112 4 inches and less than or equal to 35 percent of the cont rol surface span would have a maximum permissible free play of 4 inches x 0.020 or 0.080 inches (total motion up and down) measured at the tr ailing edge. Correct any free play in excess of this amount.

Care must also be exercised during repair or rework to prevent stress concentration points or areas that could increase the fatigue susceptibility of the trim tab system. Advisory Circular (AC) 23.629-1A, Means of Compliance with Section 23.629, “Flutter,” contains additional information on this subject.

NOTE: If the pilot has experi enced flutter, or thinks he/she has, then a complete inspection of the aircraft flight control system and all related components including rod ends, bearings, hinges, and bellcranks must be accomplished. Suspected parts should be replaced.

LOAD FACTORS FOR REPAIRS. In order to design an effective repair to a sheet metal aircraft, the stresses that act on the structure must be understood.

Six types of major stresses are known and should be considered when making repairs. These are tension, compression, bending, torsion, shear, and bearing

The design of an aircraft repair is complicated by the requirement that it be as light as possible. If weight were not critical, repairs could be made with a large margin of safe ty. But in actual practice, repairs must be strong enough to carry all of the loads with the required factor of safety, but they must not have too much extra strength. A joint that is too weak cannot be tolerated, but neither can one that is too strong because it can create stress risers that may cause cracks in other locations.

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Par 4-50 Page 4-113 TRANSFER OF STRESSES WITH-IN A STRUCTURE. An aircraft structure must be designed in such a way that it will accept all of the stresses imposed upon it by the flight and ground loads without any permanent deformation. Any repair made must accept the stresses, carry them across the repair, and then transfer them back into the original structure. These stresses are considered as flowing through the structure, so there must be a continuous path for them, with no abrupt changes in crosssectional areas along the way. Abrupt changes in cross-sectional areas of aircraft structure that are subject to cycle loading/stresses will result in stress concentration that may induce fatigue cracking and eventual failure. A scratch or gouge in the surface of a highly-stressed piece of metal will cause a stress concentration at the point of damage.

Multirow Fastener Load Transfer. When multiple rows of rivets are used to secure a lap joint, the transfer of stresses is not equal in each row. The transfer of stress at each row of rivets may be thought of as transferring the maximum amount capable of being transferred without experiencing rivet shear failure.

Use Of Stacked Doublers. A stacked doubler is composed of two or more sheets of material that are used in lieu of a single, thicker sheet of material. B ecause the stress transferred at each row of rivets is dependent upon the maximum stress that can be transferred by the rivets in that row, the thickness of the sheet material at that row need only be thick enough to transfer the stress applied. Employing this principle can reduce the weight of a repair joint.

4-39.4-49. [RESERVED.]

### Page 178

Par 4-50 Page 4-114 SECTION 4. METAL REPAIR PROCEDURES

GENERAL. The airframe of a fixed-wing aircraft is generally considered to consist of five principal units; the fuselage, wings, stabilizers, flight control surfaces, and landing gear.

Aircraft principal structural elements (PSE) and joints are designed to carry loads by distributing them as stresses. The elements and joints as originally fabricated are strong enough to resist these stresses, and must remain so after any repairs. Long, thin elements are called members. Some examples of members are the metal tubes that form engine mount and fuselage trusses and frames, beams used as wing spars, and longerons and stringers of metal-skinned fuselages and wings. Longerons and stringers are designed to carry principally axial loads, but are sometimes required to carry side loads and bending moments, as when they frame cutouts in metal-skinned structures. Truss members are designed to carry axial (tension and compression) loads applied to their ends only. Frame members are designed to carry side loads and bending moments in addition to axial loads. Beam members are designed to carry side loads and bending moments that are usually large compared to their axial loads. Beams that must resist large axial loads, particularly compression loads, in combination with side loads and bending moments are called beam-columns. Other structural elements such as metal skins, plates, shells, wing ribs, bulkheads, ring frames, intercostal members, gussets, and other reinforcements, and fittings are designed to resist complex stresses, sometimes in three dimensions.

Any repair made on an aircraft structure must allow all of the stresses to enter, sust ain these stresses, and then allow them to return into the structure. The repair must be equal to the original structure, but not stronger or stiffer, which will cause stress concentrations or alter the resonant frequency of the structure.

All-metal aircraft are made of very thin sheet metal, and it is possible to restore the strength of the skin without restoring its rigidity. All repairs should be made using the same type and thickness of material that was used in the original structure. If the original skin had corrugations or flanges for rigidity, these must be preserved and strengthened. If a flange or corrugation is dented or cracked, the ma terial loses much of its rigidity; and it must be repaired in such a way that will restore its rigidity, stiffness, and strength.

RIVETED (OR BOLTED) STEEL TRUSS-TYPE STRUCTURES. Repairs to riveted structures may be made employing the general principles outlined in the following paragraphs on aluminum alloy structures. Repair methods may also be found in text books on metal structures. Methods for repair of the major structural members must be specifically approved by the Federal Aviation Administration (FAA).

ALUMINUM ALLOY STRUCTURES. Extensive repairs to damaged stressed skin on monocoque-types of aluminum alloy structures should preferably be made in accordance with specific manufacturer’s instructions or other FAA-approved source.

Rivet Holes. Rivet holes are slightly larger than the diameter of the rivet. When driven, solid rivets expand to fill the hole. The strength of a riveted joint is based upon the expanded diameter of the rivet. Therefore, it is important that the proper drill size be used for each rivet diameter.

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Par 4-54 Page 4-115 The acceptable drill size for rivets may be found in Metallic Materials and Elements for Flight Vehicle Structure (MIL-HDBK-5).

Avoid drilling oversized holes or otherwise decreasing the effective tensile areas of wing-spar capstrips, wing, fuselage, finlongitudinal stringers, or highly-stressed tensile members. Make all repairs, or reinforcements, to such members in accordance with factory recommendations or with the specific approval of an FAA representative.

Disassembly Prior to Repairing. If the parts to be removed are essential to the rigidity of the complete structure, support the structure prior to disassembly in such a manner as to prevent distortion and permanent damage to the remainder of the structure. When rivets are removed, undercut rivet heads by drilling. Use a drill of the same size as the diameter of the rivet.

Drilling must be exactly centered and to the base of the head only. After drilling, break off the head with a pin punch and carefully drive out the shank. On thin or unsupported metal skin, support the sheet metal on the inside with a bucking bar. Removal of rivet heads with a cold chisel and hammer is not recommended because skin damage and distorted rivet holes will probably result. Inspect rivet jo ints adjacent to damaged structure for partial failure by removing one or more rivets to see if holes are elongated or the rivets have started to shear.

Effective Tools. Care must also be taken whenever screws must be removed to avoid damage to adjoining structure. When properly used, impact wrenches can be effective tools for removal of screws; however, damage to adjoining structure may result from excessive vertical loads applied through the screw axis. Excessive loads are usually related to improperly adjusted impact tools or attempting to remove screws that have seized from corrosion. Remove seized screws by drilling and use of a screw extractor. Once the screw has been removed, check for structural cracks that may appear in the adjoining skin doubler, or in the nut or anchor plate.

SELECTION OF ALUMINUM FOR REPLACEMENT PARTS. All aluminum replacement sheet metal must be identical to the original or properly altered skin. If another alloy is being considered, refer to the information on the comparative strength properties of aluminum alloys contained in MIL-HDBK-5.

Temper. The choice of temper depends upon the severity of the subsequent forming operations. Parts having single curvature and straight bend lines with a large bend radius may be advantageously formed from heat-treated material; while a part, such as a fuselage frame, would have to be formed from a soft, annealed sheet, and heat-treated after forming. Make sure sheet metal parts which are to be left unpainted are made of clad (aluminum coated) material. Make sure all sheet material and finished parts are free from cracks, scratches, kinks, tool marks, corrosion pits, and other defects which may be factors in subsequent failure.

Use of Annealed Alloys for Structural Parts. The use of annealed aluminum alloys for structural repair of an aircraft is not recommended. An equivalent strength repair using annealed aluminum will weigh more than a repair using heat-treated aluminum alloy.

HEAT TREATMENT OF ALUMINUM ALLOY PARTS. All structural aluminum alloy parts are to be heat-treated in accordance with the heat-treatment instruction issued by the manufacturers of the part. In the case of a specified temper, the sequence of heat-treating operations set forth in

### Page 180

Par 4-54 Page 4-116 MIL-HDBK-5 and corresponding specifications. If the heat-treatment produces warping, straighten the parts immediately after quenching. Heat-treat riveted parts before riveting, to preclude warping and corrosion.

Quenching. Quench material from the solution heat-treating temperature as rapidly as possible after removal from the furnace. Qu enching in cold water is preferred, although less dras tic chilling (hot or boiling water, or airblast) is sometimes employed for bulk sections, such as forgings, to minimize quenching stresses.

Reheating at Temperatures Above Boiling Water.

Reheating of 2017 and 2024 alloys above 212 °F tend to impair the original heat treat ment. Therefore, reheating above 212 °F, including the baking of primers, is not acceptable without subsequent complete and correct heat treatment.

BENDING METAL. When describing a bend in aviation, the term “bend radii” is used to refer to the inside radius. Requirements for bending the metal to various shapes are frequently encountered. When a metal is bent, it is subjected to changes in its grain structure, causing an increase in its hardness.

The minimum radius is determined by the composition of the metal, its temper, and thickness. Table 4-6 shows the recommended radius for different types of aluminum. Note that the smaller the thickness of the material, the smaller the recommended minimum bend radius, and that as the material increases in hardness, the recommended bend radii increases.

When using layout techniques, the mechanic must be able to calculate exactly how much material will be required for the bend. It is easier to lay out the part on a flat sheet before the bending or shaping is performed.

Before bending, smooth all rough edges, remove burrs, and drill relief holes at the ends of bend lines and at corners; to prevent cracks from starting. Bend lines should preferably be made to lie at an angle to the grain of the metal (preferably 90 degrees).

Bend radii (BR) in inches for a specific metal composition (alloy) and temper is determined from table 4-6. For example, the minimum bend radii for

0.016 thick 2024-T6 (alloy and temper) is found is found to be 2 to 4 times the material thickness or

0.032 to 0.064.

**SETBACK.**

Setback is a measurement used in sheet metal layout. It is the distance the jaws of a brake must be setback from the mold line to form a bend. For a 90 degree bend, the point is back from the mold line to a distance equal to the bend radius plus the metal thickness. The mold line is an extension of the flat side of a part beyond the radius. The mold line dimension of a part, is the dimension made to the intersection of mold lines, and is the dimension the part would have if its corners had no radius. (See figure 4-2.)

FIGURE 4-2. Setback for a 90-degree bend.

### Page 181

Par 4-57 Page 4-117 TABLE 4-6. Recommended radii for 90-degree bends in aluminum alloys.

Alloy and Approximate sheet thickness (t) (inch) t e m p e r

0.

0 1 6

0.

0 3 2

0.

0 6 4

0.

1 2 8

0.

1 8 2

0.

2 5 8 2024-01 0 0-1t 0-1t 0-1t 0-1t0-1t 0-1t 2024-T31, 2 1½t-3t 2t-4t 3t-5t 4t-6t 4t-6t 5t-7t 2024-T61 2t-4t 3t-5t 3t-5t 4t-6t 5t-7t 6t-10t 5052-0 0 0 0-1t 0-1t 0-1t 0-1t 5052-H32 0 0 ½t-1t ½t-1½t ½t-1½t ½t-1½t 5052-H34 0 0 ½t-1½t 1½t-2½t 1½t-2½t 2t-3t 5052-H36 0-1t ½t-1½t 1t-2t 1½t-3t 2t-4t 2t-4t 5052-H38 ½t-1½t 1t-2t 1½t-3t 2t-4t 3t-5t 4t-6t 6061-0 0 0-1t 0-1t 0-1t 0-1t 0-1t 6061-T4 0-1t 0-1t ½t-1½t 1t-2t 1½t-3t 2½t-4t 6061-T6 0-1t ½t-1½t 1t-2t 1½t-3t 2t-4t 3t-4t 7075-0 0 0-1t 0-1t ½t-1½t 1t-2t 1½t-3t 7075-T61 2t-4t 3t-5t 4t-6t 5t-7t 5t-7t 6t-10t 1 Alclad sheet may be bent over slightly smaller radii than the corresponding tempers of uncoated alloy.

2 Immediately after quenching, this alloy may be formed over appreciably smaller radii.

To determine setback for a bend of more or less than 90 degrees, a correction known as a K-factor must be applied to find the setback.

Table 4-7 shows a chart of K-factors. To find the setback for any degree of bend, multiply the sum of the bend radius and metal thickness by the K-value for the angle through which the metal is bent.

Figure 4-3 shows an example of a piece of

0.064 inch sheet metal bent through 45 degrees to form an open angle of 135 degrees. For 45 degrees, the K-factor is 0.41421. The setback, or the distance from the mold point to the bend tangent line, is:

Setback = K(BR + MT) = 0.41421 (0.25 + 0.064) = 0.130 inches

If a closed angle of 45 degrees is formed, the metal must be bent through 135 degrees. The K-factor for 135 de-grees is 2.4142, so the setback, or distance from the mold point to the bend tangent line, is 0.758 inch.

**RIVETING.**

The two major types of rivets used in aircraft are the common solid shank rivet, which must be driven using an air-driven rivet gun and bucking bar; and special (blind) rivets, which are installed with special installation tools. Design allowables for riveted assemblies are specified in MIL-HDBK-5.

Solid shank rivets are used widely during assembly and repair work. They are identified by the material of which they are made, the head type, size of shank, and temper condition.

### Page 182

Par 4-57 Page 4-118 TABLE 4-7. K-chart for determining setback for bends other than 90 degrees.

Deg. K Deg.

K Deg. K Deg.

K Deg:

K 1 0.0087 37 0.3346 73 0.7399 109 1.401 145 3.171 2 0.0174 38 0.3443 74 0.7535 110 1,428 146 3.270 3 0.0261 39 0.3541 75 0.7673 111 1.455 147 3.375 4 0.0349 40 0.3639 76 0.7812 112 1.482 148 3.487 5 0.0436 41 0.3738 77 0.7954 113 1.510 149 3.605 6 0.0524 42 0.3838 78 0.8097 114 1.539 150 3.732 7 0.0611 43 0.3939 79 0.8243 115 1.569 151 3.866 8 0.0699 44 0.4040 80 0.8391 116 1.600 152 4.010 9 0.0787 45 0.4142 81 0.8540 117 1.631 153 4.165 10 0.0874 46 0.4244 82 0.8692 118 1.664 154 4.331 11 0.0963 47 0.4348 83 0.8847 119 1.697 155 4.510 12 0.1051 48 0.4452 84 0.9004 120 1.732 156 4.704 13 0.1139 49 0.4557 85 0.9163 121 1.767 157 4.915 14 0.1228 50 0.4663 86 0.9324 122 1.804 158 5.144 15 0.1316 51 0.4769 87 0.9489 123 1.841 159 5.399 16 0.1405 52 0.4877 88 0.9656 124 1.880 160 5.671 17 0.1494 53 0.4985 89 0.9827 125 1.921 161 5.975 18 0.1583 54 0.5095 90 1.000 126 1.962 162 6.313 19 0.1673 55 0.5205 91 1.017 127 2.005 163 6.691 20 0.1763 56 0.5317 92 1.035 128 2.050 164 7.115 21 0.1853 57 0.5429 93 1.053 129 2.096 165 7.595 22 0.1943 58 0.5543 94 1.072 130 2.144 166 8.144 23 0.2034 59 0.5657 95 1.091 131 2.194 167 8.776 24 0.2125 60 0.5773 96 1.110 132 2.246 168 9.514 25 0.2216 61 0.5890 97 1-130 133 2.299 169 10.38 26 0.2308 62 0.6008 98 1.150 134 2.355 170 11.43 27 0.2400 63 0.6128 99 1.170 135 2.414 171 12.70 28 0.2493 64 0.6248 100 1.191 136 2.475 172 14.30 29 0.2586 65 0.6370 101 1.213 137 2.538 173 16.35 30 0.2679 66 0.6494 102 1.234 138 2.605 174 19.08 31 0.2773 67 0.6618 103 1.257 139 2.674 175 22.90 32 0.2867 68 0.6745 104 1.279 140 2.747 176 26.63 33 0.2962 69 0.6872 105 1.303 141 2.823 177 38.18 34 0.3057 70 0.7002 106 1.327 142 2.904 178 57.29 35 0.3153 71 0.7132 107 1.351 143 2.988 179 114.59 36 0.3249 72 0.7265 108 1.376 144 3.077 180 Inf.

The material used for the majority of solid shank rivets is aluminum alloy. The strength and temper conditions of aluminum alloy rivets are identified by digits and letters similar to those used to identify sheet stock. The 1100, 2017-T, 2024-T, 2117-T, and 5056 rivets are the six grades usually available. AN-type aircraft solid rivets can be identified by code markings on the rivet heads. A rivet made of 1100 material is designated as an “A” rivet, and has no head marking. The 2017-T alloy rivet is designated as a “D” rivet and has a raised teat on the head. Two dashes on a rivet head indicate a 2024-T alloy designated as a “DD” rivet. The 2117-T rivet is designated as an “AD” rivet, and has a dimple on the head. A “B” designation is given to a rivet of 5056 material and is marked with a raised cross on the rivet head. Each type of rivet is identified by a part number to allow the user to select the correct rivet. The numbers are in series and each series represents a particular type of head. (See figure 4-4 and table 4-8.)

An example of identification marking of rivet follows.

MS 20470AD3-5 Complete part number MS Military standard number 20470 Universal head rivet AD 2117-T aluminum alloy 3 3/32nds in diameter 5 5/16ths in length

### Page 183 (OCR)

RIVET IDENTIFICATION T = 0.064" The material can be identified by the head marking Material R = 0.250 Rivet Code Head Marking Material SETBACK PLAIN (Dyed) 1100 ° (- AD DIMPLED 2117

- SETBACK D RAISED DOT 2017T This is a 90-degree angle.

Setback = BR + MT DD TWO RAISED DASHES 2024 = 0.250 + 0.064 = 0.314" B RAISED CROSS (Dyed) 5056 T = 0.064" RAISED CIRCLE 7050 M TWO DOTS Monel R = 0.250

**PART NUMBER BREAKDOWN**

• MS20470 AD 4-7

SETBACK This is a 135-degree open angle. The metal has only been bent 45°.

SB = (BR + MT)K • MS20470 •• AD = (0.250 + 0.064) 0.4121 = 0.130" Type Head Material Diameter Length Code Measured Measured T = 0.064 in 32nds in 16ths of an inch of an inch Length is measured from the SETBACK Length top of the flush head and the R = 0.250 underside of the universal head.

SETBACK This is a 45-degree closed angle.

Setback = (BR + MT)K = (0.250 + 0.064)2.4142 = 0.758" FIGURE 4-3. Methods of determining setback for bends other than 90 degree.

Countersunk head rivets (MS20426 supersedes AN426 100-degree) are used where a smooth finish is desired. The 100-degree countersunk head has been adopted as the standard in the United States.

The universal head rivet (AN470 superseded by MS20470) has been adopted as the standard for protruding-head rivets, and may be used as a replacement for the roundhead, flathead, and brazier head rivet. These rivets can also be purchased in half sizes by designating a "0.5" after the main length (i.e., MS20470 AD4-3.5).

FIGURE 4-4. Rivet identification and part number breakdown.

Replace rivets with those of the same size and strength whenever possible. If the rivet hole becomes enlarged, deformed, or otherwise damaged; drill or ream the hole for the next larger size rivet. However, make sure that the edge distance and spacing is not less than minimums listed in the next paragraph. Rivets may not be replaced by a type having lower strength properties, unless the lower strength is adequately compensated by an increase in size or a greater number of rivets. It is acceptable to replace 2017 rivets of 3/16 inch diameter or less, and 2024 rivets of 5/32 inch diameter or less with 2117 rivets for general repairs, provided the replacement rivets are 1/32 inch greater in diameter than the rivets they replace.

Par 4-57 Page 4-119

### Page 184

Par 4-57 Page 4-120 TABLE 4-8. Aircraft rivet identification.

Material 1100 2117T 2017T 2017T-HD 2024T 5056T 7075-T73 Head Marking Plain Dimpled Raised Dot Raised Dot Raised Double Dash Raised Cross Three Raised Dashes

AN Material Code

A AD D D DD B AN425 78º Counter- Sunk Head

X X X X X X

AN426 100º Counter- Sunk Head MS20426

X X X X X X X

AN427 100º Counter- Sunk Head MS20427

AN430 Round Head MS20470

X X X X X X X AN435 Round Head MS20613 MS20615

AN 441 Flat Head

AN 442 Flat Head MS20470

X X X X X X X

AN 455 Brazier Head MS20470

X X X X X X X

AN 456 Brazier Head MS20470

X X X X X X X

AN 470 Universal Head MS20470

X X X X X X X Heat Treat Before Using No No Yes No Yes No No Shear Strength psi 10000 30000 34000 38000 41000 27000 Bearing Strength psi 25000 100000 113000 126000 136000 90000

### Page 185

Par 4-57 Page 4-121 TABLE 4-8. Aircraft rivet identification. (continued) Material Carbon Steel Corrosion- Resistant Steel Copper Monel Monel Nickel- Copper Alloy Brass Titanium Head Marking Recessed Triangle Recessed Dash Plain Plain Recessed Double Dots Plain Recessed Large and Small Dot

AN Material Code

F C M C AN425 78º Counter- Sunk Head

AN426 100º Counter- Sunk Head MS20426 MS 20426 AN427 100º Counter- Sunk Head MS20427

X X X X AN430 Round Head MS20470

AN435 Round Head MS20613 MS20615 X MS20613 X MS20613 X X MS20615 X MS20615

AN 441 Flat Head

X X X X AN 442 Flat Head MS20470

AN 455 Brazier Head MS20470

AN 456 Brazier Head MS20470

AN 470 Universal Head MS20470

Heat Treat Before Using No No No No No No No Shear Strength psi 35000 65000 23000 49000 49000 95000 Bearing Strength psi 90000 90000

### Page 186

Par 4-57 Page 4-122 Rivet edge distance is defined as the distance from the center of the rivet hole to the nearest edge of the sheet.

Rivet spacing is the distance from the center of the rivet hole to the center of the ad jacent rivet hole. Unless structural deficiencies ar e suspected, the rivet spacing and edge distance should duplicate those of the original aircraft structure. If st ructural deficiencies are suspected, the following may be used in determining minimum edge distance and rivet spacing.

For single row rivets, the edge distance should not be less than 2 times the diameter of the rivet and spacing should not be less than 3 times the diameter of the rivet.

For double row rivets, the edge distance and spacing should not be less than the minimums shown in figure 4-5.

For triple or multiple row rivets, the edge distance and spacing should not be less than the minimums shown in figure 4-5.

The 2117 rivets may be driven in the condition received, but 2017 rivets above 3/16 inch in diameter and all 2024 rivets are to be kept packed in dry ice or refrigerated in the “quenched” condition until driven, or be reheat treated just prior to driving, as they would otherwise be too hard for satisfactory riveting.

Dimensions for formed rivet heads are shown in figure 4-6(a), together with commonly found rivet imperfections.

When solid shank rivets are impractical to use, then special fasteners are used. Special fastening systems used for aircraft construction and repair are divided into two types, special and blind fa steners. Special fasteners are sometimes designed for a specific purpose in an aircraft structure. The name “special fasteners” refers to its job requirement and the tooling needed for in stallation. Use of special fasteners may require an FAA field approval.

Blind rivets are used when there is access to only one side of the structure. Typically, the locking characteristics of a blind rivet is not as good as a driven rivet. Therefore, blind rivets are usually not used when driven rivets can be installed.

CAUTION: Blind rivets should not be used on floats or amphibian hulls below the water line.

Self plugging friction-lock cherry rivets. This patented rivet may be installed when there is access to only one side of the structure. The blind head is formed by pulling the tapered stem into the hollow shank. This swells the shank and clamps the skins tightly together.

When the shank is fully upset, the stem pulls in two. The stem does not fracture flush with the rivet head and must be trimmed and filed flush for the installation to be complete.

Because of the friction-locking stem, these rivets are very sensitive to vibrations.

Inspection is visual, with a loose rivet standing out in the standard “smoking rivet” pattern.

Removal consists of punching out the frictionlocked stem and then treating it like any other rivet. (See figure 4-7.)

Mechanical-lock rivets have a device on the puller or rivet head which locks the center stem into place when installed. Many frictionlock rivet center stems fall out due to vibrations; this in turn, greatly reduces its shear strength. The mechanical-lock rivet was developed to prevent that problem.

Various manufacturers make mechanical-lock fasterners such as: Bulbed Cherrylock, CherryMax, Olympic-Loks, and Huck-Loks.

### Page 187 (OCR)

STRENGTH THROUGH THIS SECTION IS 75% OF SHEET WITHOUT HOLES.

D = DIAMETER OF 40 RIVET (PREFER- ABLY ABOUT 3x

**THICKNESS OF**

THICKER SHEET) 2D →4Dx - 20 (A) DOUBLE ROW

**CAN DO TA SA NOTES WITHOUT HOLES..**

6D 30 20-→ * 20 (B) TRIPLE OR MULTIPLE ROWS FIGURE 4-5. Rivet hole spacing and edge distance for single-lap sheet splices.

Par 4-57 Page 4-123

### Page 188 (OCR)

(a) DIMENSIONS FOR FORMED RIVET HEADS.

-1ž0 65D (ID MIN.) (MIN.) (=D MIN.)

(b) RIVETING TOOLS

**CORRECT**

**INCORRECT SHAPE FOR**

**RIVET SNAP OR SET.**

**BUCKING BAR OR DOLLY.**

(C) RIVET IMPERFECTIONS.

**RIVET DRIVEN AT SLANT, RIVET DRIVEN RIVET FLAT ON ONE**

**CORRECTLY, DOLLY SIDE OR DOLLY HELD HEAD AT SLANT. FLAT.**

BODY OF RIVET TOO SHORT. CLOSING HEAD RIVET NOT PULLED TIGHT, CLINCHES BULGED ON ACCOUNT RIVET TIGHT, PLATES SHAPED TOO MUCH BETWEEN PLATES, OF POOR FIT.

**WITH SNAP DIE. CLOSING HEAD TOO FLAT.**

**RIVETED TOO MUCH. RIVET BODY CLINCHED TOO MUCH, HEAD CRACKED.**

**PLATES CLINCHED AT RIVET MATERIAL TOO AND DRIVEN APART. HARD WHEN RIVETING TOOL FORMED.**

DAMAGED PLATE FIGURE 4-6. Riveting practice and rivet imperfections.

Par 4-57 Page 4-124

### Page 189

Par 4-57 Page 4-125

FIGURE 4-7. Self plugging friction-lock Cherry rivets.

Bulbed Cherrylock Rivets. One of the earlier types of mechanical-lock rivets developed were Bulbed Cherrylock blind rivets. These blind rivets have as their main advantage the ability to replace a solid shank rivet size for size. (See figure 4-8.)

A Bulbed Cherrylock consists of three parts;

a rivet shell, a puller, and a lockring. The puller or stem has five features which are activated during installation; a header, shank expanding section, lockring indent, weak or stem fracture point, and a serrated pulling stem. Carried on the pulling stem, near the manufactured head, is the stem lockring. When the rivet is pulled the action of the moving stem clamps together the sheets of metal and swells the shank to fill the drilled hole. When the stem reaches its preset limit of travel, the upper stem breaks away (just above the lockring) as the lockring snaps into the recess on the locking stem. The rough end of the retained stem in the center on the manufactured head must never be filed smooth, because it will weaken the strength of the lockring and the center stem could fall out. (See figure 4-8.)

The Bulbed Cherrylock rivets are available in two head styles: universal and 100° countersunk. Their lengths are measured in increments of 1/16 inch. It is important to select a rivet with a length related to the grip length of the metal being joined.

The Bulbed Cherrylock rivet can be installed using a G35 cherry rivet hand puller or a pneumatic Bulbed Cherrylock pulling tool.

The CherryMax (see figure 4-9) rivet uses one tool to install three standard rivet diameters and their oversize counterparts. This makes the use of CherryMax rivets very popular with many small general aviation repair shops.

CherryMax rivets are available in four nominal diameters 1/8, 5/32, 3/16, and 1/4 inch and three oversized diameters. CherryMax rivets are manufactured with two head styles, universal and countersunk. The CherryMax rivets consists of five parts; bulbed blind header, hollow rivet shell, locking (foil) collar, driving anvil, and pulling stem. The blind bulbed header takes up the extended shank and forms the bucktail on a CherryMax rivet stem. Rivet sleeves are made from 5056 aluminum, monel, and INCO 600. The stems are made from alloy steel, CRES, and INCO X-750 stem. CherryMax rivets have an ultimate shear strength ranging from 50 KSI to 75 KSI.

An Olympic-Lok (see figure 4-10) rivet is a light three-piece mechanically locked, spindle-type blind rivet. It carries its stem lock integral to the manufactured head. While installing, the lockring is pressed into a groove on the pulling stem just as the rivet completes drawing the metal together. After installation is completed, never file the stem of an Olympic-Lok rivet, because it will weaken the lockring attachment. The Olympic-Lok fastener is available in three head styles:

### Page 190 (OCR)

- STEM

**STEM IS PULLED**

COLLAR LOCKING BEFORE PULLING INTO RIVET SLEEVE AND STARTS TO FORM BULB BLIND HEAD BEGINS RIVET HEAD SHEET GAP > CLAMP-UP AND

**HOLE FILL**

**ACTION BEGIN SHEAR RING**

**FORMATION OF BLIND**

**CLAMP-UP COMPLETED HEAD AND HOLE FILLING**

**AS STEM CONTINUES TO ARE COMPLETED**

**BULB OUT BLIND HEAD**

**SHEAR RING NOW SHEARS**

**FROM STEM CONE TO ALLOW**

**STEM TO PULL FURTHER INTO**

**RIVET HEAD RIVET**

**FIRMLY SEATED**

**BLIND SIDE BULB HEAD**

- IS FORMED BELOW SHEAR RING GUARANTEES

**MINIMUM GRIP BLIND SIDE BULBED HEAD (IN MINIMUM GRIP SHEAR**

**RING MAY NOT SHEAR)**

**SHEAR RING HAS MOVED**

**DOWN STEM CONE UNTIL**

**PULLING HEAD AUTOMATICALLY STOPS STEM BREAK NOTCH COMPLETELY INSTALLED BULBED CHERRYLOCK**

**FLUSH WITH TOP OF**

**RIVET HEAD**

**PULLING HEAD HAS**

**LOCKING COLLAR IS NOV INSERTED LOCKING**

**READY TO BE INSERTED COLLAR AND STEM**

**HAS FRACTURED FLUSH**

**WITH RIVET HEAD (MAXIMUM GRIP**

**BLIND SIDE ILLUSTRATED)**

BULBED HEAD FIGURE 4-8. Mechanical-lock (Bulbed Cherrylock) Cherry rivet.

Par 4-57 Page 4-126

### Page 191

Par 4-57 Page 4-127

F IGURE 4-9. CherryMax rivet.

universal protruding, 100-degree flush countersink, and 100-degree flush shear; and three diameters 1/8, 5/32, and 3/16 inch. The three di ameters are available in eight different alloy combinations of 2017-T4, A-286, 5056, and monel. Olympic-Lok lock spindles are made from the same material as the sleeves.

Huck rivets (see figure 4-11) are available in two head styles, protruding and flush. They are available in four diameters 1/8, 5/32, 3/16, and 1/4 inch. Their diameters are measured in increments of 1/32 inch and lengths are measured in 1/16 inch increments. They are manufactured in three different combinations of alloys: 5056 aluminum sleeve with 2024 aluminum alloy pin, A-286 corrosion-resistant steel sleeve with an A-286 pin, and a monel 400 sleeve with an A-286 pin. The Huck fastener has the ability to tightly draw-up two or more sheets of metal together while being installed. After the take-up of the Huck fastener is completed, the lockring is squeezed into a groove on the pulling stem. The anvil or footer (of the installation tool) packs the ring into the groove of the pulling stem by bearing against the lockring.

Common pull-type Pop rivets, produced for nonaircraft related applications, are not approved for use on certificated aircraft structures or components.

Design a new or revised rivet pattern for strength required in accordance with one of the following:

The aircraft manufacturer’s maintenance manuals.

The techniques found in structural text books and using the mechanical properties found in MIL-HDBK-5.

The specific instructions in paragraphs 4-58g through 4-58n. When following the instruction in paragraphs 4-58g through 4-58n, the general rule for the diameter of the rivets used to join aluminum sheets is to use a diameter approximately three times the thickness of the thicker sheet. Do not use rivets where they would be placed in tension, tending to pull the heads off; and backup a lap joint of thin sheets with a stiffener section.

### Page 192 (OCR)

**SLEEVE REDUCED INTERNAL DIAMETER TO DOUBLE**

**PRECISELY ALIGN THE LOCKING GROOVE TAPER DESIGN**

**FULLY SERRATED STEM CONTINUOUS INTEGRAL**

**FOR INCREMENTAL PULL-UP LOCKING COLLAR RV 1201 SERIES**

**COUNTERSUNK**

**HEAD STYLE 1/16 INCH GRIP VARIATION CAPABILITY**

**DESIGNED INTO FASTENER**

**LARGER SECOND SHANK RV 1200 SERIES FLUSH BREAKING STEM DIAMETER FOR EXTRA**

**PROTRUDING SLEEVE EXPANSION AND**

**HEAD STYLE GREAT BLIND-SIDE**

**BEARING AREA**

**GREATER LOCKING COLLAR VOLUME PROTRUDING HEAD**

**DRIVEN INTO LOCKING GROOVE UNDERCUT ELIMINATES**

DEBURRING When the stem travel is Continued pulling breaks Olympic-Lok blind rivet is Tool jaws grip the stem and stopped by the sleeve's inter- the stem flush to the fasteninserted into a prepared pull it into the sleeve, closing nal step, the integral locking er head. Flushness is hole. Note the clearance the sheet gap, expanding the collar shears free and is assured by the positioning between the rivet and the sleeve to fill the hole, and forced into the stem locking of the break-groove when sides of the hole and the forming a consistently larger groove. This forms a stem travel is stopped by gap between the sheets. blind side bearing area. permanent mechanical lock. the sleeve's internal step.

FIGURE 4-10. Olympic-Lok rivet.

REPAIR METHODS AND PRECAUTIONS FOR ALUMINUM STRUCTURE. Carefully examine all adjacent rivets outside of the repair area to ascertain that they have not been harmed by operations in adjacent areas. Drill rivet holes round, straight, and free from cracks. Deburr the hole with an oversize drill or deburring tool.

The rivet-set used in driving the rivets must be cupped slightly flatter than the rivet head. (See figure 4-6.) Rivets are to be driven straight and tight, but not overdriven or driven while too hard, since the finished rivet must be free from cracks. Information on special methods of riveting, such as flush riveting, usually may be obtained from manufacturer's service manuals.

Splicing of Tubes. Round or streamline aluminum alloy tubular members may be repaired by splicing (See figure 4-12.) Splices in struts that overlap fittings are not acceptable. When solid rivets go completely through hollow tubes, their diameter must be at least one-eighth of the outside diameter of the outer tube.

Rivets which are loaded in shear should be hammered only enough to form a small head and no attempt made to form the standard roundhead. The amount of hammering required to form the standard roundhead often causes the rivet to buckle inside the tube. (Correct and incorrect examples of this type of rivet application are incorporated in figure 4-12.)

Par 4-57 Page 4-128

### Page 193

Par 4-58 Page 4-129

FIGURE 4-11. Huck rivet.

Repairs to Aluminum Alloy Members. Make repairs to aluminum alloy members with the same material or with suitable material of higher strength. The 7075 alloy has greater tensile strength than other commonly used aluminum alloys such as 2014 and 2024, but is subject to somewhat greater notch sensitivity. In order to take advantage of its strength characteristics, pay particular attention to design of parts to avoid notches, small radii, and large or rapid changes in cross-sectional areas. In fabrication, exercise caution to avoid processing and handling defects, such as machine marks, nicks, dents, burrs, scratches, and forming cracks. Cold straightening or forming of 7075-T6 can cause cracking; therefore, it may be advisable to limit this processing to minor cold straightening.

Wing and Tail Surface Ribs. Damaged aluminum alloy ribs either of the stam ped sheet-metal type or the built-up type employing special sections, square or round tubing, may be repair ed by the addition of suitable reinforcement. (A cceptable methods of repair are shown in figures 4-13 and 4-14.) These examples deal with types of ribs commonly found in small and medium size aircraft. Repair schemes developed by the aircraft manufacturer are acceptable, but any other methods of reinforcement are major repairs and require approved data.

Trailing and Leading Edges and Tip Strips. Repairs to wing, control surface tra iling edges, leading edges, and tip strips should be made by properly executed and reinforced splices. Acceptable methods of trailing edge repairs are shown in figure 4-15.

Repair of Damaged Skin. In cases where metal skin is damaged extensively, repair by replacing an entire sheet panel from one structural member to the next. The repair seams are to lie along stiffening members, bulkheads, etc.; and each seam must be made exactly the same in regard to rivet size, splicing, and rivet pattern as the manufactured seams at the edges of the original sheet. If the tw o manufactured seams are different, the stronger one will be copied. (See figure 4-16 for typical acceptable methods of repairs.)

Patching of Small Holes. Small holes in skin panels which do not involve damage to the stiffening members may be patched by covering the hole with a patch plate in the manner shown in figure 4-16. Flush patches also may be installed in stressed-skin type construction. An acceptable and easy flush patch may be made by trimming out the damaged area and then installing a conventional patch on the underneath side or back of the sheet being repaired. A plug patch plate of the same size and skin thickness as the opening may then be inserted and riveted to the patch plate. Other types of flush patches similar to those used for patching plywood may be used. The rivet pattern used, however, must follow standard practice to maintain satisfactory strength in the sheet.

Splicing of Sheets. The method of copying the seams at the edges of a sheet may not always be satisfactory.

For example, when the sheet has cutouts, or doubler plates at an edge seam, or when other members transmit loads into the sheet, the sp lice must be designed as illustrated in the following examples.

Material: Clad 2024 sheet, 0.032 inch thickness. Width of sheet (i.e., length at splice) = “W” = 10 inches.

Determine rivet size and pattern for a single-lap joint similar to figure 4-5.

Use rivet diameter of approximately three times the sheet thickness, 3 x 0.032 = 0.096-inch. Use 1/8-inch

The text continues at https://openairdocs.org/ac/43.13-1B.md?from=194

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