Structural Inspection and Repair Manual (Rev D6)

54-00-01 (Rev D0)

REPAIR OF DAMAGED NACELLE SKINS OR PANELS

1. General Information
NOTE: Repair of dented, cracked or gouged nacelle skins or panels may be accomplished in one of three different
repairs dependent upon the severity of the damage. Suggested repairs are as follows:
A. Repair Procedure (Ref. Figure 201, Sheet 1, 2 or 3)
a. If a skin or panel is cracked, stop drill all crack ends with a No. 40 drill.
b. If a skin or panel is gouged, trim out the hole smoothly and radius all corners to 1/8 inch.
c. Fabricate reinforcement doublers from the same thickness or the next size thicker material than the original as shown on
Details A, B or C of the illustrations. Material must be of the same type and temper as the original material.
d. If an insertion patch is used for the repair, the insertion must be of the same thickness and material as the original
material.
e. Drill No. 30 holes in all locations as shown in the applicable repair illustration.
f. Remove the doubler and/or insertion patch and deburr all holes.
g. Countersink all exterior surface holes. Refer to Chapter 20.
h. Secure the parts in place with clecos and install MS20426AD4 rivets of the appropriate length.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-01 (Rev D0)

Figure 201 : Sheet 1 : Repair of Damaged Skins or Panels

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 Structural Inspection and Repair Manual (Rev D6)
54-00-01 (Rev D0)

Figure 201 : Sheet 2 : Repair of Damaged Skins or Panels

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 Structural Inspection and Repair Manual (Rev D6)
54-00-01 (Rev D0)

Figure 201 : Sheet 3 : Repair of Damaged Skins or Panels

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 Structural Inspection and Repair Manual (Rev D6)
54-00-02 (Rev D0)

REINFORCEMENT/REPAIR OF UPPER FORWARD COWLING FLANGE

1. Reinforcement/Repair of Upper Forward Cowling Flange
A. Repair Procedure
a. If cracks are noted, stop drill all crack ends with a No. 40 drill.
b. If the flange is wrinkled or dented and cannot be straightened, trim out the area, smooth any sharp edges and radius all
corners to 0.25 inch as shown on Figure 201, Sheet 2.
c. Fabricate a doubler as shown on Figure 201, Sheet 1 or Sheet 2 to cover the repair area, or purchase P/N 118-910027-
53 reinforcement from Raytheon Aircraft Company. Material for the doubler, if fabricated, should be the same as the
original cowling and of the same thickness or one size thicker.
d. If material was trimmed from the cowling flange in step b, fabricate an insertion patch as shown on Figure 201, Sheet 2
of the same material and thickness as the cowling.
NOTE: The insertion patch can be cut from a purchased P/N 118-910027-53 reinforcement.
e. Position the reinforcement doubler (and insertion patch if to be installed) in place on the cowling flange and secure it in
place with clamps.
f. Drill No. 30 holes to an approximate pattern as shown on Figure 201, Sheet 1 or Figure 201, Sheet 2 (per the repair
method to be used). A minimum edge distance of 0.31 inch should be maintained on all holes.
g. Countersink all exterior surface holes. Refer to Chapter 20.
h. Remove all parts and deburr all holes.
i. Secure all parts in position with clecos and install MS20426AD4 rivets of the appropriate length on all countersink holes
and install MS20470AD4 on all other holes.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-02 (Rev D0)

Figure 201 : Sheet 1 : Reinforcement/Repair of Upper Forward Cowling Flange

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 Structural Inspection and Repair Manual (Rev D6)
54-00-02 (Rev D0)

Figure 201 : Sheet 2 : Reinforcement/Repair of Upper Forward Cowling Flange

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 Structural Inspection and Repair Manual (Rev D6)
54-00-03 (Rev D3)

REINFORCEMENT OF LOWER FORWARD COWLING ATTACH BULKHEADS

1. General Information
NOTE: Bulkhead assemblies may be purchased from Textron Aviation and installed in lieu of the following
reinforcement:
• The Bulkhead assembly part numbers are 129-910041-3, 129-910041-4, 129-910041-5 and 129-910041-
6.
A. Repair Procedure
a. Drill out and remove the existing clips at the top of each bulkhead member between the bulkhead and the lower forward
cowling fireseal. Drill out, remove and retain the existing receptacles on each clip and discard the clip.
b. Drill out and retain the existing four nut plates from the bulkhead members.
c. Stop drill the ends of all cracks with a No. 40 drill.
d. Fabricate two upper and two lower reinforcements as shown on Figure 201, Sheet 2 or purchase a nacelle bulkhead
reinforcement from Textron Aviation.
e. Position the reinforcements on the bulkhead as shown on Figure 201, Sheet 1 and clamp them in place.
f. Drill No. 30 holes through all reinforcements and the bulkheads as shown on Figure 201, Sheet 1.
g. Remove the reinforcements and deburr all holes.
h. Countersink (per Chapter 20) all holes in the aft side of the bulkheads as shown on Figure 201 , Sheet 1.
i. Secure the reinforcements to the bulkheads with clecos and rivet them in place with MS20426AD4 rivets of the
appropriate length.
j. Fabricate one LH and one RH clip as shown in Figure 201, Sheet 2, or purchase clips from Raytheon Aircraft Company
for each nacelle reinforcement.
k. Drill No. 30 holes in the clip to match the receptacles removed in Step a. Refer to Figure 201, Sheet 1.
l. Deburr and countersink (per Chapter 20) the No. 30 holes in the clips and install MS20426AD4 rivets of the appropriate
length.
m. Rivet the clips to the bulkheads with MS20426AD4 rivets and to the lower forward fireseal with MS20470AD4 rivets of
the appropriate length.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-03 (Rev D3)

Figure 201 : Sheet 1 : Reinforcement of Lower Forward Cowling Attach Bulkheads

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 Structural Inspection and Repair Manual (Rev D6)
54-00-03 (Rev D3)

Figure 201 : Sheet 2 : Reinforcement of Lower Forward Cowling Attach Bulkheads

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 Structural Inspection and Repair Manual (Rev D6)
54-00-04 (Rev D0)

ENGINE ANTI-ICE LIP CRACK REPAIR

1. General Information
A. In most cases, the anti-ice lip can be repaired without removing it from the cowling. There are precautions, however, that must
be taken to protect the aluminum cowling during welding repair of the stainless steel anti-ice lip as specified in the following
steps:
B. Repair Procedure
a. Stop drill the ends of all cracks with a No. 40 drill.
b. Purge the anti-ice lip with argon gas for 15 minutes (minimum gas flow of 5 cfm) prior to, and during the welding
procedure.
c. If the crack is in an area of the anti-ice lip that is not part of the hot air chamber, inlet or outlet tubes, the backside of the
area to be welded should be backed with a clean copper sheet 0.032-inch thick or thicker. Under no circumstances
should repair welding be performed on an area where the stainless steel anti-ice lip is lapped over the aluminum cowling
skin. The anti-ice lip must be removed from the cowling for repair of this nature.
NOTE: The argon gas purging will do much to dissipate the heat of welding, however, in some instances
it may be beneficial to use a protective heat sink compound (14, Chart 210, 91-00-00). Follow the
manufacturer's instructions for use of the product.
d. Weld the crack using the gas tungsten arc welding process (heli-arc). Use weld rod per AWS A 5.9, Class ER347.
e. Sand and buff the repaired area flush with the surrounding area.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-05 (Rev D0)

COWLING ICE VANE INSPECTION AND REPAIR

1. General Information
A. Repair Procedure
NOTE: Inspect the ice vane for looseness. If the ice vane assembly rotates at the leading or trailing edge more
than 0.10 inch, perform the following steps.
a. Remove the lower forward cowling from the airplane. Refer to the COWLING REMOVAL procedures in the applicable
maintenance manual.
b. Gain access to the ice vane mounting bolts through the ice vane access panel.
c. Remove the mounting bolts securing the ice vane and remove the vane from the cowling.
d. Install CR3242-4 rivets as indicated in Figure 201. It is permissible to install rivets one size larger if the existing hole is
oversized.
e. Install the ice vane on the cowling using the mounting bolts removed in Step c.
f. Install the ice vane access panel.
g. Install the lower forward cowling and all other cowling that was removed in Step a. Refer to the COWLING
INSTALLATION procedures in the applicable maintenance manual.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-05 (Rev D0)

Figure 201 : Sheet 1 : Cowling Ice Vane Repair

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 Structural Inspection and Repair Manual (Rev D6)
54-00-06 (Rev D0)

NACELLE STRUCTURAL DAMAGE

1. Nacelle Structural Damage
NOTE: This repair is to be performed when the nacelle has sustained structural damage due to fuel system
overpressurization.
A. Repair Procedure
a. Remove the fuel cell from the nacelle in accordance with the applicable maintenance manual.
b. Inspect the nacelle for damage and deformed structures. Chart 201 contains a list of potential repair parts.
c. Repair all damage as required.
d. Install the fuel cell in accordance with the applicable maintenance manual and rig all associated plumbing.
e. Document the repair and return aircraft to service.
Table 201. Chart 201 POTENTIAL REPAIR PARTS
QUANTITY PART NUMBER NOMENCLATURE
1 101-980000-5 Tank Liner, LH
1 101-980000-6 Tank Liner, LR
1 50-980002-179 Stiffener
2 50-980002-185 Clip
1 115-980000-227 Angle
2 101-980000-21 Stiffener
7 101-980000-29 Stiffener
1 101-980000-11 Angle, LH
1 101-980000-12 Angle, RH
1 101-980000-13 Angle, LH
1 101-980000-14 Angle, RH
1 No. 375 Tape
2 50-921587-3 Gasket
2 FCB 25719 Gasket
6 600-015-1/4 Stat-O-Seal
1 50-921587-3 Gasket
2 FCB 25722 Gasket
8 115-921041 Gasket
2 99-920032-1 Gasket
AR NAS1738B-4-2 Rivet
AR NAS1738B-4-3 Rivet
AR NAS1738B-4-4 Rivet
AR NAS1739B-4-2 Rivet
AR NAS1739B-4-3 Rivet
AR MS20470-4-4 Rivet
AR MS20470-4-6 Rivet

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 Structural Inspection and Repair Manual (Rev D6)
54-00-07 (Rev D0)

ICE DOOR HINGE LINE REPAIR

1. Ice Door Line Repair
NOTE: This procedure provides instructions to repair cracks in the ice door hinge line.
A. Repair Procedure
a. Fashion a 2.0 x 22.0 doubler from 0.040-inch thick 6061-T4 Al. Bend the doubler to form an angle to match the contour
as shown in Figure 201.
b. Remove the rivets from the bulkhead.
c. Refer to Chapter 20-10-02 and prepare the area for bonding. Install the doubler. Bond using adhesive (62, Chart 201,
91-00-00).
d. Rivet the doubler to the bulkhead using existing holes. Rivets on duct side of the doubler shall be spaced 1 inch apart.
NOTE: All rivets are to be driven from the inside out
e. Fill all cracks with aerodynamic smoother (Chart 202, 91-00-00).

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 Structural Inspection and Repair Manual (Rev D6)
54-00-07 (Rev D0)

Figure 201 : Sheet 1 : Inlet Duct

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 Structural Inspection and Repair Manual (Rev D6)
54-00-07 (Rev D0)

Figure 201 : Sheet 2 : Inlet Duct

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 Structural Inspection and Repair Manual (Rev D6)
54-00-08 (Rev D0)

UPPER FORWARD COWLING CUT-OUT REPAIR

1. Doubler Repair
A. Repair Procedure
a. Stop drill the end of the crack and slot to edge with a #40 drill bit. Smooth any sharp edges.
b. Fabricate a doubler from like material 6061-T6 Al. The doubler should be 0.046 or 0.051 inch thick, 3.40 inch x 1.25
inch, form 90°, contour to match. Lower flange to be 0.70 inch, trim to match cut-out profile in top cowl edge (Refer to
Figure 201). If the cut-out is not cracked, trim as shown in Figure 203.
c. Stop drill the end of the crack with a #40 drill bit. Smooth any sharp edges.
d. Remove the paint in the area of the patch and rough sand with #40 or #60 grit aluminum oxide paper.
e. Clean the area to be patched with solvent (2 or 12, Chart 208, 91-00-00).
f. Bond repair against lower (inside) side of upper cowl with adhesive (13, Chart 201, 91-00-00). Ensure repair is flush to
disc (flange) and centered. Install with NAS1097AD4 rivets spaced as shown in Figure 202.
2. Fiberglass Repair
A. Repair Procedure
a. Stop drill the end of the crack with a #40 drill bit. Smooth any sharp edges.
b. Remove the paint in the area of the patch and rough sand with #40 or #60 grit aluminum oxide paper.
c. Clean the area to be patched with solvent (2 or 12, Chart 208, 91-00-00).
d. Apply the fiberglass patch as follows:
1. The patch must extend at least 1 inch on each side and the end of the crack.
2. The patch must consist of three layers of glass fabric (6, Chart 209, 91-00-00) laid up with a 30-60-90 degree ply
orientation. glass fabric (7, Chart 209, 91-00-00) may be substituted.
3. Use adhesive (38, Chart 201, 91-00-00) and hardener (35, Chart 201, 91-00-00) as the adhesive for the patch.
e. Trim the cured fiberglass patch in the cut out to the dimensions in Figure 203.
f. Install MS20426AD4 rivets around the edge of the patch. Space rivets at 1 inch intervals with a 0.30 inch edge distance.
The total number of rivets required will be determined by the size of the patch.
NOTE: The purpose of the rivets is to provide additional security between the patch and the skin.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-08 (Rev D0)

Figure 201 : Sheet 1 : Doubler

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 Structural Inspection and Repair Manual (Rev D6)
54-00-08 (Rev D0)

Figure 202 : Sheet 1 : Trim Dimensions

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 Structural Inspection and Repair Manual (Rev D6)
54-00-08 (Rev D0)

Figure 203 : Sheet 1 : Fiberglass Patch Trim Dimensions

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 Structural Inspection and Repair Manual (Rev D6)
54-00-09 (Rev D0)

CRACKS IN INBOARD AND OUTBOARD SIDES OF ENGINE FORWARD AIR INLET DUCT REPAIR (MODEL C90, C90A,
C90GT, C90GTI, E90, AND F90)
1. Engine Forward Air Inlet Duct
A. Repair Procedure
NOTE: This repair describes the repair of the air inlet duct due to cracks on either or both the inboard and
outboard sides of the forward air inlet duct. The repair is applicable to P/N 109-910029-9, -121, -127, -149
and -155.
a. Eddy current inspect the crack location to find the ends of the crack.
1. Perform CALIBRATION (STANDARDIZATION) procedure (Ref. Chapter 20-00-00).
2. Perform GENERAL SURFACE INSPECTIONS procedure (Ref. Chapter 20-00-00).
3. Perform INDICATION EVALUATION FOR SURFACE INSPECTION procedure (Ref. Chapter 20-00-00).
b. Make a repair doubler from 0.050 inch thick 2024-T3 per QQ-A-250/5 to match the existing contour of the duct at the
cracked area.
NOTE: Fore and aft edges in the airstream must be chamfered.
c. Drill holes as shown in Figure 201. Drill 0.128 to 0.133 inch diameter holes for solid rivets and 0.129 to 0.132 inch
diameter for blind rivets (refer to Step g. for rivet details). Use the existing rivet locations as shown.
NOTE: Maintain two times diameter minimum edge distance on all fastener positions. Maintain 1.00 inch
spacing on all new fastener positions.
Maintain a minimum of two rows of fasteners in the repair doubler beyond the crack in all four
directions (forward, aft, inboard and outboard). It is permissible to add one additional row of
fasteners on forward and/or aft end of repair doubler to accommodate the above requirement for
cracks near or at maximum allowed lengths.
d. Deburr all sharp edges.
e. Cold bond the repair doubler in position with adhesive (62, Chart 201, 91-00-00) per manufacturer's instructions.
f. Protect all bare metal with chem film per AMS SAE-C-5541, class 1A.
g. Install MS20426AD4 rivets, length to suit. Optional rivets CR3212-4 can be used. Install wet with adhesive (15, Chart
201, 91-00-00).

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 Structural Inspection and Repair Manual (Rev D6)
54-00-09 (Rev D0)

Figure 201 : Sheet 1 : Right-Hand Forward Air Inlet Duct Assembly (Left-Hand Opposite)

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 Structural Inspection and Repair Manual (Rev D6)
54-00-10 (Rev D0)

CRACKS IN INBOARD AND OUTBOARD SIDES OF ENGINE FORWARD AIR INLET DUCT REPAIR (MODEL B300 SERIES)
1. Engine Forward Air Inlet Duct
A. Repair Procedure
NOTE: This repair is applicable to P/N 101-910119-151.
a. Eddy current inspect the crack location to find the ends of the crack.
1. Perform CALIBRATION (STANDARDIZATION) procedure (Ref. Chapter 20-00-00).
2. Perform GENERAL SURFACE INSPECTIONS procedure (Ref. Chapter 20-00-00).
3. Perform INDICATION EVALUATION FOR SURFACE INSPECTION procedure (Ref. Chapter 20-00-00).
b. Stop drill the ends of the crack with a 3/16 inch dia. drill.
c. Make a repair doubler from 0.050 inch thick 2024-T3 per QQ-A-250/5 to match the existing contour of the duct at the
cracked area.
NOTE: Fore and aft edges in the airstream must be chamfered.
d. Drill holes as shown in Figure 201. Drill 0.128 to 0.133 inch dia. holes for solid rivets and 0.129 to 0.132 inch dia. for
blind rivets (refer to Step h. for rivet details). Use the existing rivet locations as shown.
NOTE: Maintain 2 times dia. minimum edge distance on all fastener positions. Maintain 1.00 inch spacing
on all new fastener positions.
NOTE: Maintain a minimum of two rows of fasteners in the repair doubler beyond the crack in all four
directions (forward, aft, inboard and outboard). It is permissible to add one additional row of
fasteners on forward and/or aft end of repair doubler to accommodate the above requirement for
cracks near or at maximum allowed lengths.
e. Deburr all sharp edges.
f. Cold bond the repair doubler in position with adhesive (62, Chart 201, 91-00-00) per manufacturer's instructions.
g. Protect all bare metal with chem film per AMS SAE-C-5541, class 1A.
h. Install MS20426AD4 rivets, length to suit. Optional rivets CR3212-4 can be used. Install wet with adhesive (15, Chart
201, 91-00-00).

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 Structural Inspection and Repair Manual (Rev D6)
54-00-10 (Rev D0)

Figure 201 : Sheet 1 : Right-Hand Forward Air Inlet Duct Assembly (Left-Hand Opposite)

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 Structural Inspection and Repair Manual (Rev D6)
54-00-11 (Rev D0)

LOOSE AND PULLED THROUGH RIVETS ON LEFT AND RIGHT NACELLE UPPER AFT SKIN REPAIR (MODELS 200, B200,
300 AND B300 SERIES)
1. Loose and Pulled Through Rivets on Left and Right Nacelle Upper Aft Skin
A. Repair Procedure
a. Make a repair strap from QQ-A-250/5, 2024-T3 0.040 inch thick to the profile shown in Figure 201.
b. Drill holes to match the positions shown in Figure 201. Drill the holes to the minimum diameter necessary to remove the
damage. Maintain the minimum edge distance of 1.7 x diameter on the skin and underlying structure.
c. Deburr all sharp edges.
d. Fill redundant countersinks using drilled-off rivet heads. Cold bond in place with (15, Chart 201, 91-00-00) adhesive per
the manufacturer's instructions.
e. Cold bond the strap in place using adhesive (62, Chart 201, 91-00-00) per the manufacturer's instructions.
f. Install rivets wet, using (15, Chart 201, 91-00-00)) adhesive per the manufacturer's instructions as shown in Figure 201.
g. Apply a top coat to match the existing.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-11 (Rev D0)

Figure 201 : Sheet 1 : Repair Strap

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 Structural Inspection and Repair Manual (Rev D6)
54-00-12 (Rev D0)

ENGINE INERTIAL VANE SHOCK LINK ASSEMBLY REPAIR (FA-1 THRU FA-230, FL-1 THRU FL-344, FM-1 THRU FM-9, FN-1)
1. Shock Link Assembly Repair (P/N 101-910114-125)
NOTE: This section describes the repair of engine inertial vane shock link assembly (P/N 101-910114-123/-125) due
to loose or damaged pin.
a. Remove loose or damaged pin from shock link assembly (Ref. Figure 201).
CAUTION: Do not remove retainer material while grinding off dry film lubricant from the hole in the retainer.
b. Grind off dry film lubricant from the hole in the retainer.
c. Install replacement pin part number NASM16555-622 in the shock link assembly.
d. Pull spring out of the way to avoid damaging spring temper and tack weld, using best shop practices, over two holes in the
retainer.
2. Shock Link Assembly Repair (P/N 101-910114-123 and P/N 101-910114-125)
a. Remove loose or damaged pin from shock link assembly (Ref. Figure 201).
b. Install replacement pin part number NASM16555-624 in the shock link assembly.
c. Peen over each end of pin to permanently retain pin in retainer.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-12 (Rev D0)

Figure 201 : Sheet 1 : Engine Inertial Vane Shock Link Assembly Repair

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 Structural Inspection and Repair Manual (Rev D6)
54-00-13 (Rev D0)

CRACK: L/R NACELLE, COWLING BULKHEAD CANT STA 111.86 (MODEL 300, 300LW, B300 AND B300C) (SR-KA-00052)
1. Crack: Left and Right Nacelle, Cowling Bulkhead
A. Affected Part Numbers:
Bulkhead Frames: 101-910116-3 and 101-910116-9
Bulkhead Assemblies: 101-910116-1, 101-910116-7 and 101-910116-11
2. Repair Limitations:
This standard repair applies only for crack(s) that form in the cutline area as defined below and shown in Figure 201. This repair is
not applicable for crack(s) that form in a different area than defined in cutline area of Figure 201. If damage exceeds cutline area or a
different location, contact Beechcraft for further assistance.
3. Repair Instructions
1. Carefully de-rivet existing fasteners as required to complete the steps indicated below. Remove and discard existing clip
taking care not to oversize existing holes and not to damage any surrounding and/or underlying structures.
2. Use fluorescent penetrant or eddy current NDT technique in accordance with the King Air SIRM, 20-00-00-201, to locate and
mark the end of the crack.
3. Carefully trim crack out as shown in Figure 201. Removing minimum material necessary and taking care not to damage any
surrounding and/or underlying structure. Trim 0.10 min. beyond end of crack. Maintain 2D min. edge distance to all adjacent
fasteners and maintain minimum 0.16 radius.
4. Use fluorescent penetrant or eddy current NDT techniques in accordance with the King Air SIRM, 20-00-00-201, to ensure no
crack(s) remains and no cracks have formed. Report any adverse findings to Beechcraft RDO for further assessment before
continuing with this repair.
5. Fabricate a repair angle doubler (SR-KA-00052-1 (-2 opposite)), as shown in Figure 201. Form to nest into existing
bulkhead. Make from 0.050 thick, 2024-0 Alclad sheet per SAE AMS-QQ-A-250/Minimum bend radius 0.12. Heat treat to
T42 condition per SAE AMS2770 after forming. Alternate Material: Make from 2024-T3 alclad sheet per SAE AMS-QQ-A-
250/5, crack check using fluorescent penetrant NDT technique in accordance with the King Air SIRM, 20-00-00-201, to
ensure no cracks have formed. if cracks are found, discard and remake.
6. Fabricate a new repair clip (SR-KA-00052-3), size and profile to match existing clip except to allow for added angle doubler.
Min. bend radius 0.19. Make from 0.032 thick 2024-T3 Clad aluminum sheet per SAE AMS-QQ-A-250/5. Maintain min. 2D
edge distance to all picked up fasteners.
7. Drill and countersink all holes as shown in Figure 201, in accordance with the King Air SIRM, 20-50-02-201. Maintain 2D min.
edge distance at all fastener locations and 4D - 6D spacing on all new fastener locations.
8. Deburr all sharp edges.
9. Protect bare metal with chem film per MIL-DTL-5541, Class 1A, and prime with one coat of MIL-PRF-23377, Type 1, Class C
primer (8, Chart 206, 91-00-00-201) in accordance with the King Air SIRM, 20-30-00-201.
10. Install repair parts in place using EA 9309.3NA or EA 9359.3 cold bond adhesive (62, Chart 201, 91-00-00-201) mixed and
applied in accordance with the manufacturer's instructions.
11. Install fasteners as shown in Figure 201, in accordance with the King Air SIRM, 20-50-00-201. Length to suit.
12. Fill corners and cutout in repair area using SAE AMS-S-8802 PSR type sealant (PR1440 or equivalent (23, Chart 207, 91-
00-00-201) in accordance with the King Air SIRM, 20-10-08. Alternate Material: Permissible to fill corners and cutout with
RTV-109 (35, Chart 207, 91-00-00-201).
13. Topcoat to match as required.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-13 (Rev D0)

Figure 201 : Sheet 1 : Repair Angle Doubler

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 Structural Inspection and Repair Manual (Rev D6)
54-00-14 (Rev D0)

UPPER AFT ENGINE COWLING DOOR LOUVER PANEL REPLACEMENT (MODEL 200 SERIES AIRPLANES)
1. Upper Aft Engine Cowling Door Louver Panel Replacement
A. Affected Part Numbers
Louver Panel 101-910069-83
Upper Aft Engine Cowling Door Assemblies 101-910069-2, -81, -83, -99, -107, -109, -121, -123, -133 and -135
2. Repair Limitations
This standard repair applies only for the replacement of the compressor bleed valve cover louver panel which is spot welded to the
right upper aft engine cowling door assembly.
3. Repair Instructions
1. Perform the AFT UPPER COWLING REMOVAL procedure.
(Refer to 71-10-05, 401 of the Super King Air 200 Series Maintenance Manual.)
(Refer to 71-10-05, 401 of the Super King Air B200GT/B200CGT Fusion Maintenance Manual.)
2. Drill a 0.096 to 0.116 inch hole through each spot weld that attaches the louver panel to the aft upper cowling door (Ref.
Figure 201). Remove the louver panel from the aft upper cowling door.
3. Grind the spot weld locations in the aft upper cowling as necessary to provide a smooth mating surface for the new louver
panel.
CAUTION: Use care when grinding the spot weld areas. The material thickness in the area of the louver panel
must not be less than 0.032 inch.
4. Perform a liquid fluorescent penetrant inspection or an eddy current inspection on the area where the louver panel was
located. If cracks are found, contact Textron Aviation Technical Support before continuing this repair.
5. Enlarge the holes at the spot weld locations to 0.128 - 0.133 inch. Countersink the holes 100 degrees x 0.028 inch depth on
the inside surface of the cowling door.
6. Measure the distance between the holes at the spot weld locations. If the distance between any of the holes is more than 1.1
inch, drill an additional 0.128 - 0.133 inch diameter fastener hole between the two existing holes. Make sure that a minimum
2D edge distance is maintained on any new fastener holes.
7. Position the new louver panel on the cowling door. Make sure that the louver fins are in the correct orientation and that a gap
of 0.03 ± 0.03 inch exists between the edges of the louver panel and the adjacent cowling door skin.
8. Reposition or trim the edges of the louver panel as necessary to maintain a gap of 0.03 ± 0.03 inch between the edges of the
louver panel and the cowling door skin.
9. Temporarily attach the louver panel in this position and match drill fastener holes into the panel through the holes in the
cowling door.
10. Remove the louver panel and countersink the holes 100 degrees x 0.028 inch depth on the outside facing surface of the
panel.
11. Deburr all sharp edges and apply a chemical film conversion coating (4, Chart 203, 91-00-00) to all bare metal surfaces.
12. Apply epoxy primer (5 and 6, Chart 206, 91-00-00) to the louver panel. Allow the primer to fully cure.
13. Temporarily attach the louver panel to the cowling door with clecos. Install the rivets double flush. Remove clecos as the
riveting progresses.
14. Paint the louver panel to match the color and scheme of the airplane.
15. Perform the AFT UPPER COWLING INSTALLATION procedure.
(Refer to 71-10-05, 401 of the Super King Air 200 Series Maintenance Manual.)
(Refer to 71-10-05, 401 of the Super King Air B200GT/B200CGT Fusion Maintenance Manual.)

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 Structural Inspection and Repair Manual (Rev D6)
54-00-14 (Rev D0)

Figure 201 : Sheet 1 : Upper Aft Engine Cowling Door Louver Panel

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 Structural Inspection and Repair Manual (Rev D6)
54-00-15 (Rev D3)

STANDARD REPAIR - CRACK, L/H / R/H AFT LOWER ENGINE COWLING, INBOARD / OUTBOARD LONGERONS, FS 147.00
AT WL 84.75
1. Airplane Applicability Data
A. Airplane Applicability
This standard repair is for the following airplanes:
B300 / B300C Series: All Variants, All Serials.
B. Affected Part Numbers:
101-910118-11 /-12 Longeron with 101-9100184-25 / -35 Clevis Assembly (Repair Figure 201, Sheet 2)
130-910038-3 / -4 Longeron with 130-910036-1 Eyebolt (Repair Figure 201, Sheet 2)
117-910048-1 / -2 Longeron with 117-910054 Cowing Keeper (Repair Figure 201, Sheet 4)
117-910074-1 / -2 Longeron (Repair Figure 201, Sheet 6)
2. General Notes
A. This repair does not affect the published maintenance or inspection requirements. The accomplishment of this repair does not
reset the current component cycles.
B. This repair is authorized for the stated airplanes only.
C. This repair is applicable to an airplane which is unaltered by a third-party modification. It is the responsibility of the repair facility
to contact the originator or holder of any relevant non-Textron Aviation STC to determine compatibility.
D. It is the responsibility of the repair facility to ensure that the actual damage is within the damage limits defined by this repair.
E. All referenced documents shall be worked to the latest published revision.
F. Embodiment of this repair must be recorded in the airplane log book stating the repair number and revision.
G. Airplane weight and balance change is negligible.
H. All dimensions are in inches unless otherwise specified.
3. Limitations
A. Cracks shall not exceed 1 inch length.
4. Repair Procedure
A. Review the aircraft configuration and affected parts in Section 1.0 to determine appropriate Figure and Associated Views for
repair procedures.
B. Carefully derivet existing fasteners as required to complete the steps indicated below, taking care not to oversize existing holes
and not to damage any surrounding and/or underlying structures. All damaged rivets must be removed and replaced.
C. If the firewall attach angle (ref P/N 101-980030-11/-12) is damaged, remove and replace with a new or serviceable detail. If no
damage is found, omit this procedure.
D. Use fluorescent penetrant and 10x magnification or eddy current NDT technique in accordance with the King Air SIRM, 20-00-
00, to locate and mark the end of crack.
E. Carefully cut out the crack from the longeron as shown in Figures and Associated Views as determined in Step 4.A, removing
minimum material necessary and taking care not to damage any surrounding and/or underlying structures. Trim 0.10 beyond
end of crack. Maintain 2D minimum edge distance on existing fasteners and maintain minimum 0.25 corner radius. Ensure
surface finish is 63 RMS (or better).
F. Deburr all sharp edges.
G. Use fluorescent penetrant and 10x magnification or eddy current NDT technique in accordance with the King Air SIRM, 20-00-
00, to ensure no crack remains and no cracks have formed. Report any adverse findings to Textron Aviation Team Structures
for further assessment before continuing with this repair.
H. (Ref. Figure 201, Sheet 2) Repair, Omit step I.
(1) Make new repair angle SR-KA-ATA54-16293-1, as shown in (Ref. Figure 201, Sheet 2) and Associated Views. Form to
nest into existing longeron.
Make from 0.040 thick, 2024-O condition Al Clad sheet per SAE AMS-QQ-A-250/5. Minimum bend radius 0.06. Maintain
R 0.25 minimum corner radius. Heat treat to T42 condition per SAE AMS2770 after forming.
NOTE: Permissible to make from 0.040 thick, 2024-T3 Al Clad sheet per SAE AMS-QQ-A-250/5. Minimum
bend radius 0.12. Maintain R 0.25 minimum corner radius.
(2) Make repair filler SR-KA-ATA54-16293-2, as shown in (Ref. Figure 201, Sheet 2) and Associated Views. Profile to match

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 Structural Inspection and Repair Manual (Rev D6)
54-00-15 (Rev D3)

removed section of longeron.

Make from 0.040 thick, 2024-TO condition Al Clad sheet per SAE AMS-QQ-A-250/5. Minimum bend radius 0.06. Maintain
R 0.25 minimum corner radius. Heat treat to T42 condition per SAE AMS2770 after forming.
NOTE: Permissible to make from 0.040 thick, 2024-T3 Al Clad sheet per SAE AMS-QQ-A-250/5. Minimum
bend radius 0.12. Maintain R 0.25 minimum corner radius.
I. (Ref. Figure 201, Sheet 4 or Sheet 6) Repair, Omit step H.
(1) Make new repair angle SR-KA-ATA54-16293-3, as shown in (Ref. Figure 201, Sheet 4 or Sheet 6) and Associated
Views. Form to nest into existing longeron.
Make from 0.050 thick, 2024-O condition Al Clad sheet per SAE AMS-QQ-A-250/5. Minimum bend radius 0.06. Maintain
R 0.25 minimum corner radius. Heat treat to T42 condition per SAE AMS2770 after forming.
NOTE: Permissible to make from 0.050 thick, 2024-T3 Al Clad sheet per SAE AMS-QQ-A-250/5. Minimum
bend radius 0.12. Maintain R 0.25 minimum corner radius.
(2) Make repair filler SR-KA-ATA54-16293-4, as shown in (Ref. Figure 201, Sheet 4 or Sheet 6) and Associated Views.
Profile to match removed section of longeron.
Make from 0.050 thick, 2024-TO condition Al Clad sheet per SAE AMS-QQ-A-250/5. Minimum bend radius 0.06. Maintain
R 0.25 minimum corner radius. Heat treat to T42 condition per SAE AMS2770 after forming.
NOTE: Permissible to make from 0.050 thick, 2024-T3 Al Clad sheet per SAE AMS-QQ-A-250/5. Minimum
bend radius 0.12. Maintain R 0.25 minimum corner radius.
J. Check repair parts after forming using fluorescent penetrant and 10x magnification NDT technique in accordance with the King
Air SIRM, 20-00-00, to ensure no cracks have formed. Discard and make new if any discrepancies are discovered.
K. Drill/ream and countersink all holes to match new and existing fastener locations as shown in Figures and Associated Views as
determined in Step 4. A, in accordance with the King Air SIRM, 20-50-02. Maintain minimum 2D edge distance on all fasteners,
except where noted, and 4D to 6D spacing on all fasteners, except where noted.
(1) If tooling hole shown in Figures can maintain minimum 2D edge distance in repair angle, match drill to 0.191/0.196 dia and
install a MS20470AD6 fastener.
L. Deburr all sharp edges.
M. Protect all bare aluminum with chem film per MIL-DTL-5541, Class 1A (Alodine 600 or 1200S, or Bonderite M-CR 600 Aero or
M-CR 1200S Aero), in accordance with the King Air SIRM, 20-30-00, and prime with one coat of MIL-PRF-23377, Type 1,
Class C or PR143 primer.
N. Install repair parts in place using EA 9309.3NA or EA 9359.3 cold bond adhesive mixed and applied in accordance with the
manufacturer’s instructions.
O. Install fasteners as shown in Figures and Associated Views as determined in Step 4A, in accordance with the King Air SIRM,
20-50. Fastener length to be determined upon installation.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-15 (Rev D3)

Figure 201 : Sheet 1 : Crack, L/H / R/H Aft Lower Engine Cowling, Inboard/ Outboard Longerons, FS 147.00 At WL 84.75

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 Structural Inspection and Repair Manual (Rev D6)
54-00-15 (Rev D3)

Figure 201 : Sheet 2 : Crack, L/H / R/H Aft Lower Engine Cowling, Inboard/ Outboard Longerons, FS 147.00 At WL 84.75

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 Structural Inspection and Repair Manual (Rev D6)
54-00-15 (Rev D3)

Figure 201 : Sheet 3 : Crack, L/H / R/H Aft Lower Engine Cowling, Inboard/ Outboard Longerons, FS 147.00 At WL 84.75

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 Structural Inspection and Repair Manual (Rev D6)
54-00-15 (Rev D3)

Figure 201 : Sheet 4 : Crack, L/H / R/H Aft Lower Engine Cowling, Inboard/ Outboard Longerons, FS 147.00 At WL 84.75

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 Structural Inspection and Repair Manual (Rev D6)
54-00-15 (Rev D3)

Figure 201 : Sheet 5 : Crack, L/H / R/H Aft Lower Engine Cowling, Inboard/ Outboard Longerons, FS 147.00 At WL 84.75

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 Structural Inspection and Repair Manual (Rev D6)
54-00-15 (Rev D3)

Figure 201 : Sheet 6 : Crack, L/H / R/H Aft Lower Engine Cowling, Inboard/ Outboard Longerons, FS 147.00 At WL 84.75

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 Structural Inspection and Repair Manual (Rev D6)
54-00-15 (Rev D3)

Figure 201 : Sheet 7 : Crack, L/H / R/H Aft Lower Engine Cowling, Inboard/ Outboard Longerons, FS 147.00 At WL 84.75

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 Structural Inspection and Repair Manual (Rev D6)
54-00-16 (Rev D3)

STANDARD REPAIR - CHAFE, ENGINE COWL, INLET AIR DUCT ASSEMBLY, AT NACELLE CANT STA 39.40 FWD SURFACE
OF 109-910059 INLET DUCT PLATE
1. Airplane Applicability Data
A. Airplane Applicability
This standard repair is for the following airplanes
C90: All Variants, All Serials.
B. Affected Part Numbers:
109-910029-163 and 109-910029-173 Inlet Duct Assemblies using 109-910059-5, -17 or -23 Inlet Duct Plate
2. General Notes
A. This repair does not affect the published maintenance or inspection requirements. The accomplishment of this repair does not
reset the current component cycles.
B. This repair is authorized for the stated airplanes only.
C. This repair is applicable to an airplane which is unaltered by a third-party modification. It is the responsibility of the repair facility
to contact the originator or holder of any relevant non-Textron Aviation STC to determine compatibility.
D. It is the responsibility of the repair facility to ensure that the actual damage is within the damage limits defined by this repair.
E. All referenced documents shall be worked to the latest published revision.
F. Embodiment of this repair must be recorded in the airplane log book stating the repair number and revision.
G. Airplane weight and balance change is negligible.
H. All dimensions are in inches unless otherwise specified.
3. Repair Procedure
A. Carefully de-rivet existing fasteners and unbolt existing hardware as required to complete the steps indicated below, taking care
not to oversize existing holes and not to damage any surrounding and/or underlying structures.
B. Carefully trim inlet duct skin as shown in Figure 2, taking care not to damage any surrounding and/or underlying structures.
Maintain 2D minimum edge distance to existing fasteners (measured from center of hole to edge ofpart) and maintain minimum
radius as shown.
C. Deburr all sharp edges.
D. Use fluorescent penetrant and 10X magnification or eddy current NDT technique in accordance with the King Air SIRM, 20-00-
00, to ensure no cracks have formed. Report any adverse findings to Textron Aviation Team Structures for further assessment
before continuing with this repair.
E. Make a repair Angle (SR-KA-ATA71-22193-1), as shown in Ref. Figure 201, Sheet 2.
Make from 0.040 thick, 2024-T3 Al Clad sheet per SAE AMS-QQ-A-250/5. Bend radius 0.12.
F. Drill/ream and countersink/dimple as required, all holes in accordance with the King Air SIRM, 20-50-02. Maintain 2D minimum
edge distance (measured from center of hole to edge of part) and 4D to 6D pitch distance on all fasteners (measured from
center to center).
G. Deburr all sharp edges.
H. Install repair part in place using EA 9309.3NA or EA 9359.3 cold bond adhesive mixed and applied in accordance with the
manufacturer’s instructions.
I. Protect all bare aluminum with chem film per MIL-DTL-5541, Class 1A (Alodine 600 or 1200S, or Bonderite M-CR 600 Aero or
M-CR 1200S Aero), in accordance with the King Air SIRM, 20-30-00, and prime with one coat ofMIL-PRF-23377, Type 1,
Class C or PR143 primer.
J. Install fasteners as shown in Ref. Figure 201, Sheet 2, in accordance with the King Air SIRM 20-50. Install blind fasteners wet
using EA 9309NA or EA 9359 cold bond adhesive in accordance with the manufacturer’s instructions. Length of fasteners to be
determined upon installation.
K. Make sure no further chafe exists.

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 Structural Inspection and Repair Manual (Rev D6)
54-00-16 (Rev D3)

Figure 201 : Sheet 1 : Chafe, Engine Cowl, Inlet Duct Assembly, At Nacelle CANT

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 Structural Inspection and Repair Manual (Rev D6)
54-00-16 (Rev D3)

Figure 201 : Sheet 2 : Chafe, Engine Cowl, Inlet Duct Assembly, At Nacelle CANT

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