MAINTENANCE MANUAL 72-00-00 - ENGINE - INSPECTION/CHECK

Manual Part No.3017042 Rev. 52.0 - 15/APR/19

72-00-00
Engine Model(s): MODEL(S)PT6T-3/PT6T-3B/PT6T-3BE/PT6T-3BF/PT6T-3BG

ENGINE - INSPECTION/CHECK
1. General
A. Personnel involved in maintenance practices should refer to Chapter 70-00-00, STANDARD
PRACTICES to familiarize themselves with general procedures.
B. For additional information regarding Special Tools and Fixtures, Equipment, and Consumable
Materials referred to in this section, refer to section TOOLS/FIX/EQUIP or CONSUMABLE
MATERIALS respectively.
C. Hours or Engine Hours means Engine Flight Hours. Engine flight hour is defined as the engine
operating time between aircraft takeoff (weight-off-wheels/skid off) and landing (weight-on-
wheels/skid on).
D. Unless otherwise specified, "Scheduled/Periodic Inspections" based on calendar times do not
apply during long-term storage (29 days or more) of engines/accessories preserved (on or off
aircraft) as per engine maintenance manual instructions.
2. Consumable Materials
The consumable materials listed below are used in the following procedures.

Item No. Name

PWC09-003 Silicone Sealer

3. Special Tools
The special tools listed below are used in the following procedures.

Tool No. Name

PWC34910-101 DELETED
PWC34910-109 Borescope Assemby
PWC34910-200 Guide Tube
PWC34913 DELETED
PWC34941 Wrench

4. Fixtures, Equipment and Supplier Tools

The fixtures, equipment and supplier tools listed below are used in the following procedures.

Name Remarks
Test Set Barfield Model 2312G-8 or equivalent

5. Procedure
The inspections outlined in Paragraphs 6., 7. and 8. following are considered a normal function of
operating organizations and are intended as a guide for minimum inspection and maintenance
requirements. The intervals at which these inspections are performed depend on the nature and
condition of engine operation and are in addition to the routine daily checks detailed in the
applicable aircraft manual. For example, engines operated in sandy or dusty environments or in
smog or salt-laden atmospheres should be subjected to regular inspection for corrosion and
compressor first-stage blade erosion as detailed in Table 601 and Paragraph 9., following.
Specific inspections (Ref. Para. 9.) and special condition inspections (Ref. Para. 10.) are given for
engines that have exceeded normal operating parameters or have been subjected to abnormal

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Manual Part No.3017042 Rev. 52.0 - 15/APR/19

operating conditions.
6. Inspection Frequency Definition (Ref. Table 601)
Maintenance inspection/check requirements detailed in this chapter relate directly to the engine, for
additional information detailing airframe accessibility refer to applicable aircraft manual.
A. 25 hours.
Periodic inspection requirements detailed at this frequency shall not exceed 25 hours of flight
and may be performed in conjunction with applicable aircraft maintenance schedule.
B. 50/100/150/300/600/900/1200 and 2500 hours.
Periodic inspection requirements detailed at these frequencies are in addition to the
requirements detailed at 25 hours.
C. Periodic Inspection - Tolerances
NOTE: The following tolerance is established for maintenance scheduling convenience only and
must be approved by the governing civil aviation authority.
Unless otherwise stated, the tolerance for periodic inspections is ten percent (10%), or up to a
maximum of 100 hours operating time, whichever is less. The tolerance for scheduled inspection
is ten percent (10%) or 30 days, whichever is less.
Subsequent intervals will be adjusted to re-establish the original schedule. When an inspection
is done more than 10% early, subsequent inspections will be advanced as required to not
exceed the maximum tolerance. Concurrence and final approval of the inspection interval
tolerance by the governing civil aviation authority is the responsibility of the owner/operator.
7. Periodic Inspections (Ref. Table 601)
Table 601 details periodic inspection criteria and frequency. Where no inspection limits are given in
the table, or in any other relative chapter within this manual, any item with wear or damage as
described in this table shall be replaced, or returned to an approved overhaul facility for possible
repair, as applicable.
Table 601 Engine Periodic Inspections

COMPONENT NATURE OF INSPECTION FREQUENCY

CAUTION: DO NOT MIX DIFFERENT BRANDS OF OIL WHEN CHANGING, OR

ADDING OIL BETWEEN CHANGES. THE CHEMICAL STRUCTURE OF
DIFFERENT BRANDS OF OIL MAY DIFFER SUFFICIENTLY TO MAKE THEM
INCOMPATIBLE (REF. SB5001) FOR APPROVED LISTING).

Oil Contents

Check oil level and replenish if required (Ref. 72-00-00, 25 hours

ENGINE - SERVICING). If engine has been stationary for
more than 12 hours, run affected power section for two
minutes, shut down and check oil level

Exhaust Duct Cracks or distortion 50 hours or 6

months
whichever comes
first.

P3 Filter Inspect for system contamination and flush (Ref. Para 100 hours
Housing and 10.N.).
Tubing (Post-
SB5309
Post-SB5359
and
Pre-SB5367)

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COMPONENT NATURE OF INSPECTION FREQUENCY

P3 Air Filter Clean (electrosonic) and visually inspect filter element 100 hours
Element (Ref. 73-10-07).

NOTE: 1. Individual operators may elect to extend this

interval if operating experience shows less
frequent cleaning and inspection to be
adequate.

NOTE: 2. Cleaning and visual inspection applies to

both AFCU P3 filter and optional clutch
carbon seal P3 filters (Ref. SB5174).

P3 Air Filter Clean and Inspect (Ref. 73-10-07). 100 hours

Drain Valve
(Post-
SB5309,
SB5359
and/or
SB5367)

NOTE: 1. It is recommended that drain valves on

engines operating in dirty or desert
conditions be cleaned and inspected at 50
hour intervals.

NOTE: 2. For operators who perform frequent

compressor recovery washes. It is
recommended to inspect and clean drain
valve at intervals not exceeding 50 hours.

Fuel Surge Visually inspect the accumulator for signs of metal 100 hours
Accumulator distortion on the sides and ends. The distortion will show
(Pre-SB5398) up as bulging of the sides or doming of the end cap.

If distortion is present, it is possible that cracking has

begun in the weld that joins the end cap to the body of
the accumulator. Bulging or doming can be verified
using a metal straightedge on the sides and ends of the
accumulator.

Replace the accumulator, if any metal distortion is

found.

Perform engine deceleration check (Ref. 71-00-00).

Fuel Surge Visually inspect. Replace accumulator if pop-out 150 hours

Accumulator indicator is extended (Ref. 73-10-03).
(Post-
SB5398)

Fireseals Cracks and security of brackets. 150 hours

Wiring Check for security of all accessible connections, clamps 150 hours or 6
and brackets and for evidence of wear, chafing, cracks months
and corrosion. whichever comes
first.

Inlet Screen Cleanliness and condition of mesh. 150 hours

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COMPONENT NATURE OF INSPECTION FREQUENCY

Do a visual inspection of the compressor air inlet screen
for broken and or missing wire mesh segments and
signs of brittleness. If you suspect an area, check the
integrity of the mesh by manipulating and flexing the
screen. Any deviation is a cause for replacement.
If the inlet screen is removed, do an inspection of first-
stage compressor blades per Para. 9.B.
Do a general visual inspection (GVI) of wash-ring
assembly and safety wire for security and obvious
damage. If any deviation is found during GVI, do an
inspection of first-stage compressor blades per Para.
9.B.

Gas Cracks, distortion, corrosion and evidence of 150 hours or 6

Generator overheating. months
Case whichever comes
first.

NOTE: No cracks are allowed.

RGB Output Check for oil leaks. 150 hours or 12

Shaft Seal months
whichever comes
first.

Tubing Check resistance of lead and heating element of heated 150 hours or 6
pneumatic tubes (Ref. 73-10-08). months
whichever comes
first.

Check for security of all accessible connections, clamps

and brackets, evidence of wear, chafing, cracks and
corrosion and evidence of fuel or oil leaks. Examine
insulation on pneumatic tubes for cuts in outer rubber
sheaths.

NOTE: Refer to Overhaul Manual, 72-00-00, ENGINE -

REPAIR for tubing repair instructions.

Surface cuts up to three inches long on Pre-SB5409 P3

heated tubes and Post-SB5409 Pg/P3 heated tubes
insulation, can be repaired. If surface damage exceeds
specified limits above, but depth is superficial, sheathing
may be repaired by replacement of damaged portion.
Examine the metal braid and electrical lead on heated
tubes for cuts. Up to three broken wires per plait, or six
broken wires per linear foot, are acceptable. Replace the
tube assemblies where metal braid is chafed or worn
through to the lead conductors.

Oil Filter Clean (Electrosonic), then examine the oil filter element 150 hours
Element (Ref. 79-20-02, OIL FILTERS - MAINTENANCE
(10 micron) PRACTICES).

Chip Accessory Gearbox - 2 detectors 150 hours

Detectors Reduction Gearbox - 3 detectors 150 hours

Clean with lint-free cloth and measure the minimum

weight lift capacity. Use a suitable ohmmeter to check
the continuity of the electrical circuit when the two poles
are shorted together. Reject the component if there is

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Manual Part No.3017042 Rev. 52.0 - 15/APR/19

COMPONENT NATURE OF INSPECTION FREQUENCY

no continuity. Loose pins and broken potting compound
are also the cause to reject them.

Pre-SB5354: Inspect colored stripes painted on

connector and housing. If stripes are displaced, reject
chip detector. If no stripes are apparent, paint three
equispaced colored stripes on connector and housing
using suitable paint Connector threads must be free of
paint.

NOTE: Refer to Chapter 79-30-01 for further inspection

details.

Bleed Air Check the orientation of the bleed air case assembly 150 hours
Case (Ref. 72-20-00).
Assembly
(Post-
SB5445)

Ignition Security 300 hours

System Units

Spark Igniters Cleanliness and condition. Functional check (Ref. 74-00- 300 hours
00)

Compressor At the required frequency and whenever condition of 300 hours or 12

Inlet inlet screen warrants its removal, check compressor months
Area inlet area for corrosion, dirt deposits and erosion and whichever comes
check first-stage blades as detailed in Paragraph 9. first. (Ref. NOTE)
following.

NOTE: Engines operating in a highly erosive

environment and/or conditions where there is a
high potential for blade damage require more
frequent inspections (Ref. Para.9. )

P3 Air Tube Inspect for contaminants and clean. (Located on R.G.B. 300 hours
Metering Tees Output Housing)

MFCU Throttle With power sections shut down and fuelboost pump ON, 300 hours or 12
Shaft check both sides of throttle shaft for fuel leakage. If months
leakage is evident, replace manual fuel control unit (Ref. whichever comes
73-20-01). first.

AFCU Check AFCU for bearing wash-out indicated by blue dye 300 hours and
(grease and fuel mixed) at AFCU vent hole (Ref. ENGINE whenever area is
- FAULT ISOLATION). accessible.

Drivebody inspection/driveshaft bearing replacement; See NOTE.

Route to qualified shop for detailed inspection and
Driveshaft Bearing replacement.

NOTE: This inspection is recommended for operators

that experience premature AFCU removals due
to:
1. Operating in a severe environment, with respect to
temperature, humidity, sand/dust or other
airborne contaminants. And/or
2. Very low or irregular utilization over a period of time.
Definition of regular/irregular utilization
Regular utilization: A unit which accumulates a minimum

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COMPONENT NATURE OF INSPECTION FREQUENCY

of 30 hrs/month.
Irregular utilization: A unit which remains inactive for
extended periods of time.
3. The interval of 3000 hours (mid-TBO) is
recommended to get the maximum benefit of
this inspection.

Fuel Pump Check for security and fuel leaks. 300 hours

If fuel pump gear-set and coupling operating time since

new (TSN) is LESS THAN Engine Basic TBO (Ref.
SB5003):

(a) Check fuel pump coupling in-situ for fretting and Every 600 hours
corrosion (Ref. 73-10-02, FUEL PUMP - MAINTENANCE
PRACTICES).

(b) Remove fuel pump and inspect the drive coupling Every 2000 hours
and cover accessory gearbox side for signs of reddish-
brown (iron oxide) stains. If stains are observed, return
the fuel pump to an approved overhaul facility (Ref. 73-
10-02, FUEL PUMP - MAINTENANCE PRACTICES).

Use the applicable safety wire to make sure the oil mist
hole at the end of the fuel pump drive spline (AGB side)
is clear of blockage.

If fuel pump gear-set or coupling operating time since

new (TSN) is MORE THAN Engine Basic TBO (Ref.
SB5003) OR is UNKNOWN OR fuel pump was
overhauled before 2007:

(a) Remove fuel pump and inspect the drive coupling 300 hours
and cover accessory gearbox side for signs of reddish-
brown (iron oxide) stains. If stains are observed, return
the fuel pump to an approved overhaul facility (Ref. 73-
10-02, FUEL PUMP - MAINTENANCE PRACTICES).

Use the applicable safety wire to make sure the oil mist
hole at the end of the fuel pump drive spline (AGB side)
is clear of blockage.

NOTE: As of January 2010, the fuel pump gear-set and

drive coupling replacement is required at every
pump overhaul.

Fuel Pump Install new fuel pump filter (Ref. 73-10-02). 600 hours
Filter

Fuel Pump Inspect filter for contamination and damage (Ref. 73- 300 hours
Filter 10-02). Replace as necessary. With power sections shut
down and fuel boost pump ON, check for leaks.

NOTE: Operators with helicopters equipped with fuel

filter impending bypass indicating device may
elect to waive this inspection and to perform
the fuel pump filter replacement at 600 hours.

Oil to Fuel Immediately after shutdown: Check temperature of 300 hours

Heater heater at heater fuel outlet or fuel pump filter housing.
If greater than 140°F (60°C) replace oil to fuel heater.

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COMPONENT NATURE OF INSPECTION FREQUENCY

Alternatively, perform either of the following methods:
(1) Perform touch test at heater fuel outlet or fuel pump
filter housing. If too hot to touch, replace oil to fuel
heater.

NOTE: Temperatures less than 140°F (60°C) may be

considered to be comfortable to touch.

(2) Immediately after shutdown apply a temperature

recorder to the fuel pump filter housing and replace oil
to fuel heater if the center of the recorder turns to black.

NOTE: The temperature limit of 140°F (60°C) may be

verified using the Temp-Plate Temperature
Recorder (PWC05-329).

Fuel Nozzle Do a leak test and functional test of fuel manifold 600 hours
Assemblies adapter and nozzle assemblies (Ref. 73-10-05, FUEL
MANIFOLD ADAPTERS - MAINTENANCE PRACTICES).
NOTE: When you remove the fuel nozzle and from your
service experience, you can do an in-service borescope
inspection of the hot section components (Ref. Para. 12.
A.).

For improved hot-section durability, inspect and clean 600 hours

fuel nozzle assemblies.
At any fuel nozzle assembly removal, inspect gas
generator case for cracks at boss welds (Ref. Para. 13.
C.).

NOTE: As an alternative to the removal for inspection

and functional check, a 300 hour in-situ
cleaning procedure is offered (Ref. 72-00-00,
ENGINE - CLEANING).

If in-situ cleaning is used the fuel nozzles shall be 1200 hours

removed for inspection and functional check to coincide
with airframe maintenance schedule.

Oil Filter Clean (Ultrasonic Method) and inspect filter elements at 900 hours or 24
Element an overhaul facility, using the approved equipment (Ref. months
Engine Overhaul Manual), prior to further use. Following whichever comes
this cleaning at overhaul level, the filter may be utilized first.
for a further 900-hour or 24-month period maintaining
the same inspection and cleaning schedule.

P3 Air Filter Clean (ultrasonic) and visually inspect filter element 900 hours
(Ref. 73-10-07). Do pressure drop check of elements
(Ref. 73-10-07, ADJUSTMENT/TEST).

Oil Scavenge Remove retaining plate at rear fireseal and inspect tubes 900 hours
and for fretting wear (Ref. 79-20-06).
Pressure
Tubes

Flanges and All external joint faces shall be sealed with silicone 900 hours or 12
Joint grease (PWC09-003) or Corrosion-X corrosion inhibiter months
Faces (PWC15-011) (Ref. 72-00-00, ENGINE - CLEANING) at whichever comes
major inspection. All external bolts, studs and flanges first.
are also treated with a film of silicone grease or NOTE: Engines

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COMPONENT NATURE OF INSPECTION FREQUENCY

corrosion-X corrosion inhibiter. Particular attention operating
should be paid to accessory gearbox mount flanges. in a highly
corrosive
environment
require
more frequent
sealing.
Establish
schedule/frequency
as
required to ensure
proper
corrosion
protection.

T5 System Carry out functional check (Ref. Para. 9. C. following) at 1200 hours
major inspection or when Troubleshooting indicates
necessity.

Gas For Pre-SB5239 gas generator cases inspect seam welds 1200 hours
Generator as detailed in Para. 9. following. initially:
Case 600 hours
thereafter

RGB Inspect clutch assemblies (Ref. Overhaul Manual). 2500 hours

(PT6T-
3BE/-3BG
only)
See Notes 1
and 2.

NOTE: 1. A gearbox operating as a PT6T-3BE/BG unit for any time during the 2500 hours
maintenance interval requires the regular inspection at 2500 hours.
NOTE: 2. A gearbox with more than 2500 hours accumulated since new or overhaul being
converted to a PT6T-3/-3BE/-3BG must have the 2500 hours inspection carried
out during the conversion.

8. Unscheduled Inspections
A. Chip Detectors (Ref. 79-30-01)
(1) If a chip detector warning light illuminates inspect chip detectors for particles. In addition
the oil strainer must be cleaned and checked and the appropriate oil filter checked.
(2) Inspect chip detectors at each engine oil change (Ref.
SB5001).
B. Fuel Pump Filter (Ref. 73-10-02)
(1) On new aircraft, check filter after each day of operation until no contamination is
found.
(2) Whenever any component upstream of the filter has been replaced, check filter after first
engine run.
C. Fuel Pump (Ref. 73-10-02)
(1) If airframe boost pump (either electrically or engine driven) fails or is inadvertently left off
for a cumulative time in excess of 10 hours, the engine driven fuel pump must be removed
and replaced.
(2) The removed fuel pump should be forwarded to an approved overhaul facility for inspection
(Ref. Para. 10. A.).

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D. Silver Flakes in Oil

NOTE: Presence of silver flakes in the oil system is acceptable (Ref. 79-30-01).
(1) If silver flakes are found:
(a) Flush applicable oil system (Ref. 72-00-00, ENGINE CLEANING).
(b) Inspect oil filters (Ref. 79-20-02) after next 5, 10, 25 and 50 flight hours. If after 50
flight hours silver flakes are no longer present, return to scheduled interval (Ref Table
601). If silver particles are still evident after 50 flight hours, consult your local P&WC
field service representative.
NOTE: Procedures to identify metal particles are found in (Ref. 70-00-00, STANDARD
PRACTICES - MAINTENANCE PRACTICES).

E. Reduction Gearbox Replacement

(1) During ground test, check for air, and oil leaks. Rectify as
required.
(2) Following the post installation ground test and after the first 5-10 and 20-25 flight hours,
inspect chip detectors and oil filters for contamination (Ref. 79-30-01).
(a) If silver flakes are found, flush oil system (Ref. Para. 8. D.).

F. Power Section Replacement

(1) During ground test, check for air, oil and fuel leaks. Rectify as
required.
(2) Following the post installation ground test and after the first 5-10 and 20-25 flight hours,
inspect chip detectors and oil filters for contamination (Ref. 79-30-01).
(3) If silver flakes are found, flush oil system (Ref. Para. 8.
D.).
9. Specific Inspections
A. Gas Generator Case Longitudinal Seam Welds (Pre-SB5239)
(1) A dye penetrant in-situ spotcheck inspection of the gas generator case longitudinal seam
welds must be accomplished using the following procedure.
(a) Pre-clean inspection area, apply cleaner/remover (SKF-NF/ZC-TB) to lint free cloth and
then wipe surface clean with cloth.
NOTE: Do not spray cleaner/remover on surface.
(b) Apply penetrant (SKF-HF/S) and allow a 15-to-30 minute penetration period.
(c) Wipe to remove penetrant from surface using cleaner moistened cloth.
(d) Shake can containing developer (SKD-NF) vigorously for approximately 2 minutes, then
apply thin film and allow to dry 5-to-15 minutes or until dry to touch.
(e) Inspect the seam weld (Ref. Fig. 601). Cracks will show as bright red lines against white
developer background.

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Figure 601 Location of Seam Weld on Gas Generator Case

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CAUTION: IN CONTRAST, VOIDS OR SURFACE SCRATCHES WILL SHOW AS

WHITE LINES AND THEREFORE SHOULD NOT BE MISCONSTRUED AS
CRACKS.
(f) Should a crack be evident, the area should be cleaned, then the crack verified using a
5X or 10X magnifying glass.
(g) If crack is confirmed, power section must be removed and sent to overhaul facility for
gas generator case modification.

B. First-stage Compressor Blades

CAUTION: KEEP COMPRESSOR INLET AREA CLEAR OF ALL FOREIGN OBJECTS

THAT COULD BE INGESTED BY ENGINE AND CAUSE DAMAGE DURING
SUBSEQUENT OPERATION.
(1) General
NOTE: The following inspection should be done whenever the air inlet screen is removed
(Ref. 72-20-00). Due to limited access, the use of a mirror is recommended.
(a) The limits stated under 'Acceptable Damage Limits' and 'Acceptable Erosion Limits' are
considered acceptable without repair. The actual damage must be documented and
additional inspections scheduled to assess the rate of defect progression.
(b) Continued operation of a compressor with FOD, either accepted as-is or blended is
dependent on satisfactory engine performance.
NOTE: When assessing damage, it should be understood that continued operation with
a large number of blades damaged or eroded at or near maximum acceptable
limits may result in reduction of compressor efficiency and engine
performance.
(c) The primary considerations of acceptable damage limits are based on structural
integrity. Rotors with damage more than 'Acceptable Damage Limits' cannot stay in
service unless damage can be removed by blending (Ref. 72-30-05, COMPRESSOR
ROTOR - MAINTENANCE PRACTICES):
NOTE: Remove the minimum material required to remove the damage. Document the
damage (Ref. 72-30-05, COMPRESSOR ROTOR - MAINTENANCE PRACTICES).
(d) If damaged blade is within acceptable limits and is not reworked, inspect area at 150
hours or less as established by the operator. Subsequent inspections are at the
discretion of the operator based on the type of operating environment but must not
exceed 300 hours.
NOTE: The operating environment affects the rate of erosion and incidence of blade
damage. Engines operating in a highly erosive and/or conditions where there is
a high potential for blade damage; require inspections more frequently at an
interval established by the operator.
(e) If damage to the blades is acceptable, and/or after blending, engine performance is
within acceptable limits, inspection of the subsequent stages, (rotors, stators and
impeller) is not required. Under these conditions, the engine is considered acceptable
for service.
(f) If damage exceeds limits and cannot be removed by blending or engine performance is
unsatisfactory after repair, the engine or power section must be removed from service.

(2) Documenting Damage

(a) Defects/damage must be documented in the engine log book (i.e., engine module
TSN/TSO, cycles, component description, location and dimensions, if obtainable).
(b) If the blade is repaired document the amount of material removed during blending.

(3) Acceptable Damage Limits (Ref. Fig. 602)

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Figure 602 Compressor First-stage Blades Inspection

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(a) Multiple superficial leading edge nicks 0.030 in. maximum depth on any number of
blades.
NOTE: A nick is considered to be a sharp surface indentation caused by impact
resulting in parent metal loss.
(b) Leading or trailing edge dents 0.055 in. deep by 0.080 in. long. Four dents per blade on
either leading or trailing edge on any number of blades.
NOTE: A dent is considered to be an indentation caused by impact in which the parent
metal has been displaced.
(c) Leading or trailing edge tip curl without presence of tearing. Maximum acceptable
deformation is 0.100 in. on one edge; four blades only.
(d) Leading or trailing edge nicks 0.050 in. maximum depth. Four nicks maximum per
leading or trailing edge on any number of blades.
(e) No cracks are permitted.
(f) Blade tip rub is acceptable providing there is no rolled over material and an engine
performance check is acceptable.
(g) Area a and c;
1 Corrosion pitting, nicks and dents 0.005 in. deep are acceptable.

(h) Area b;
1 Corrosion pitting, nicks and dents 0.010 in. deep are acceptable.

(i) Area d;
1 Corrosion pitting 0.002 in. deep is acceptable.
2 No foreign object damage allowable in root fillet radius.

(4) Acceptable Erosion Limits

(a) Erosion at the root is acceptable providing that dimension from leading edge to apex of
eroded area does not exceed 0.250 inch and there is no feathered, (i.e., sharp)
edge(s).
(b) Feathered edges must be removed and original leading edge profile restored (Ref. 72-
30-05, COMPRESSOR ROTOR - MAINTENANCE PRACTICES) or the engine or power
section must be removed from the service.

C. Functional Check of T5 Temperature Sensing System

(1) Remove external leads from terminal block on gas generator case.
NOTE: 1. On Pre-SB5118 engines the T5 system comprises a wiring harness, a bus-bar
assembly and ten thermocouple probes. On Post-SB5118 and PT6T-3B engines
the system comprises an integral wiring harness which includes eight
thermocouple probes.
NOTE: 2. The following check will reveal circuit malfunctions such as intermittent short
circuits or an open circuit in the harness lead. If necessary, further checks may
be carried out on the system by disassembling the engine to gain access to
probes as detailed in Chapter 77-20-01, Page Blocks 401 and 601.
(2) Using Barfield Test Set Model 2312G-8 or equivalent, check insulation resistance of harness
when measured between either chromel or alumel terminal and ground. Minimum
acceptable resistance is 5000 ohms.
(3) Check continuity and loop resistance measured across chromel and alumel terminals.
Acceptable loop resistance is 0.58 to 0.74 ohms on Pre-SB5118 engines and 1.8 to 2.6
ohms on Post-SB5118, PT6T-3B and PT6T-3BE engines.
NOTE: Several broken or damaged probes would not necessarily result in an out-of-

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tolerance loop resistance. However erroneous temperature indications could occur

due to the resultant imbalance in the harness circuit.
10. Special Condition Inspections
A. General
The inspection data given here apply to engines that have been subjected to abnormal operating
conditions. Whenever the corrective action to be taken requires that the engine or component
be sent to an overhaul facility, precise details of the condition(s) must be provided to facilitate
overhaul inspection. The following paragraphs treat each special condition separately and
prescribe the corrective action to be taken.
B. Overtorque
For action to be taken in the event of overtorque conditions, refer to Figure 603.

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Figure 603 Overtorque Limits for Each Power Section

(SHEET 1 OF 2)

(PT6T-3 only)
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(SHEET 2 OF 2)

(PT6T-3B/-3BE/-3BF/-3BG only)
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C. Overtemperature

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(1) For action to be taken in the event of overtemperature conditions during starting, refer to
Figure 604.
Figure 604 Overtemperature Limits - Starting Conditions Only

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(2) For action to be taken in the event of overtemperature conditions other than during starting,
refer to Figure 605.
Figure 605 Overtemperature Limits - All Conditions Except Starting
(SHEET 1 OF 3)

(PT6T-3 only)
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(SHEET 2 OF 3)

(PT6T-3B)
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(SHEET 3 OF 3)

(PT6T-3BF/-3BG)
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D. Overspeed
(1) Gas Generator

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For action to be taken in the event of overspeed condition refer to Figure 606.
Figure 606 Overspeed Limits - Gas Generator
(SHEET 1 OF 2)

(PT6T-3)
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(SHEET 2 OF 2)

(PT6T-3B)
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(2) Power Turbine

Do the necessary maintenance procedure in the event of Nf overspeed condition (Ref. Fig.
607).
NOTE: 1. For No Load and Transient Limitations, refer to 71-00-00, ADJUSTMENT/TEST.
NOTE: 2. A power turbine overspeed that does not involve damage to No.3 and 4 bearings

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is not cause for rejection of reduction gearbox.

Figure 607 Overspeed Limits - Power Turbine
(SHEET 1 OF 2)

(PT6T-3)
icn-00198-g000013843-001-01

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(SHEET 2 OF 2)

(PT6T-3B/3BE/3BF/3BG)
icn-00198-g000013842-001-01

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E. Sudden Stoppage or Main Rotor Blade Strike

NOTE: Damage to the tail rotor or seizure of any part of the tail rotor transmission system
must not be considered to have caused engine sudden stoppage/blade strike unless
indications exist that internal damage exists in the main gearbox.
(1) Main Rotor Blade Strike:
(a) If helicopter main rotor blades have suffered dynamic major collision damage result in
replacement of one or more main rotor blades, remove the complete twin-pac engine
and sent to an overhaul facility for light overhaul (Sudden Stoppage).
(b) If damage to the main rotor blade(s) is not severe enough to cause an engine anomaly
or does not warrant their replacement, inspect as follows:
1 Examine the engine mounts and vibration insulators for distortion or damage.
2 Examine the exhaust duct for buckling or rippling.
3 Examine the oil filters and chip detectors for metal particles.
4 Examine the power turbines for sign of rub.
5 If any of the above conditions exist, engine must be shipped to overhaul facility for
light overhaul (Sudden Stoppage).
6 If none of the above conditions exist, and there is no bad conditions in reduction
gearbox. The engine can remain in service subject to a satisfactory ground test (Ref.
Applicable Aircraft Manual) and an oil filter inspection every 25 hours for the next
100 hours of operation.

(2) Sudden Stoppage:

(a) If any part of the transmission system between reduction gearbox output shaft and
main rotor has seized while the engine is operating, examine as given in Steps (1) (b).
(b) If a non-engaged clutch engages during the shutdown, then remove the reduction
gearbox and send to an overhaul facility for light overhaul (Sudden Stoppage).

F. Oil Systems - Unusual Conditions

(1) Unusual oil system conditions are:
(a) Oil temperature of the power section and/or RGB more than 120 °C (248 °F) (Ref. Note
1).
(b) Oil mixed with different oil types or with unapproved oil brands or with chemical
materials.
(c) Engine shows high oil pressure.
(d) Engine shows low or no oil pressure.

(2) If you find the unusual oil system conditions (Ref. step (1) ), then do the steps that follow:
(a) Examine the oil filter and the chip detector for metal particles and too much carbon
deposits.
(b) Examine the Airframe Oil Cooling System (Ref. Rotorcraft Manual).
(c) Drain the oil. Flush the oil system, then fill the oil tank (Ref. 72-00-00, SERVICING).
(d) Do an engine ground run for 20 minutes at moderate power (Ref. Rotorcraft Manual).
(e) Examine the oil filter and the chip detector again for metal particles and too much
carbon deposits.
(f) If there are no metal particles and the oil pressure is in limits, the engine can stay in
service, do an oil filter and chip detector inspection after 10 flight hours and after 25
flight hours. If you do not find debris, then return to regular inspection intervals.
(g) If there are metal particles or unusual oil pressure condition continues, send the
affected reduction gearbox and/or power section to an overhaul facility for light

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overhaul (Engine Oil System - Unusual Conditions).

NOTE: 1. Oil temperature to a maximum of 120 °C (248 °F) permitted at OAT's of 43
°C (110 °F) and above. Oil temperatures between 115 to 120 °C (239-248
°F) at OAT's below 43 °C (110 °F) could be an indication of an oil cooling
issue.
NOTE: 2. For engine removed for light overhaul (Engine Oil System - Unusual
Conditions), state location of affected filter elements and whether there was
oil contamination.

(3) DELETED.
(4) DELETED.
(5) If the Total Acidic Number (TAN) or water content of the oil is not within the specified limits,
then do as follows:
(a) Do an oil analysis of the drained oil sample for the below specified limits:
1 TAN: Maximum 2 mg KOH/g
2 Water content: 800 - 1000 ppm (parts per million) maximum or 0.08 - 0.1%

(b) If you do not find TAN on an oil brand specification and the TAN is above 2 mg KOH/g,
or if the water content is more than 800 ppm, either by weight or volume, then drain
and discard the oil from the main oil tank and reduction gearbox.
(c) Fill the oil system again and operate the engine (Ref. Rotorcraft Manual).

G. Immersion in Water
Send engine to an overhaul facility for light overhaul (Immersion in Water). State if engine was
stationary and cold, stationary and hot, or rotating at time of immersion.
H. Dropped Engine or Component
Send engine or component to an overhaul facility for light overhaul (Dropped Engine or Hard
Landing).
I. Hard Landing
(1) Inspect all mounting pads for cracks or
misalignment.
(2) Inspect gas generator case for warping or
buckling.
(3) Inspect all attachment hardware at engine flanges for shearing or other obvious
damage.
(4) Inspect fireseals for warping or
buckling.
(5) Inspect all external tubes for
damage.
(6) Inspect exhaust area for cracks, warping or
distortion.
(7) Inspect accessory gearbox case adjacent to the starter and fuel control pads for
cracks.
(8) Inspect intake case struts at intersection with rear wall for
cracks.
(9) Inspect fuel control components, lines and fittings for
damage.
(10)If any of the above conditions are found, engine must be sent to an overhaul facility for
light overhaul (Dropped Engine or Hard Landing). If none of the above conditions are found,
engine may remain in service subject to a satisfactory ground test (Ref. applicable Aircraft

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Manual).
J. Inlet Blockage
(1) Power sections developing surging conditions or sudden increase in T5 temperature, inspect
as follows:
(a) Remove and inspect inlet screen.
(b) Inspect first-stage compressor blades for damage.
(c) Determine if observed T5 temperature exceeds limits (Ref. Special Condition Inspection,
Overtemperature).

K. Lightning Strike
(1) Examine the oil filter and the chip detector for metal
particles.
(2) If there are no metal particles, the engine is serviceable. Examine the oil filter every five
hours or once a day for the next 25 flight hours.
(3) If there are metal particles, drain the oil and fill the oil tank (Ref. 72-00-00, ENGINE -
SERVICING).
(4) Do an engine ground run for 20 minutes at moderate power (Ref. Applicable Rotorcraft
Manual).
(5) Examine the oil filter and the chip detector again for metal
particles.
(6) If there are no metal particles, the engine is serviceable. Examine the oil filter every five
hours or once a day for the next 25 flight hours.
(7) If there are metal particles, send engine to an approved overhaul facility for light overhaul
in accordance with the overhaul manual.
L. Contamination by Fire Extinguishing Agents
(1) In the event of engine contamination by fire extinguishing agents when you run the engine,
do the cleaning procedure that follows:
(a) Do a dry motoring run to remove remaining deposits (Ref. 71-00-00, POWER PLANT -
ADJUSTMENT/TEST).
(b) Clean the engine externally with clean water only (Ref. 72-00-00, ENGINE -
CLEANING).
(c) Do an engine motoring performance recovery wash, ignore the dry motoring run (Ref.
72-00-00, ENGINE - CLEANING).
(d) Remove engine for light overhaul. Record as a contamination by fire extinguishing
agents.

(2) In the event of engine contamination by fire extinguishing agents when you do not run the
engine, do the cleaning procedure that follows:
(a) Clean the engine externally with clean water only (Ref. Chapter 72-00-00, ENGINE-
CLEANING).
(b) Do an engine motoring performance recovery wash.
(c) Monitor over the next 50 flight hours for corrosion.

(3) For an engine contaminated by carbon dioxide and/or Halon fire extinguishing agent
(aircraft onboard fire extinguishing system), no engine maintenance is necessary.
M. Aircraft Flown Through Volcanic Ash or Smoke
(1) Do the external engine wash (Ref. ENGINE- CLEANING).
(2) Wash compressor and turbine (Ref. 72-00-00, ENGINE - CLEANING).

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(3) Drain and refill oil system with new oil (Ref. 72-00-00, ENGINE - SERVICING).
(4) Clean or change oil filters (3 places) (Ref. 79-20-02, OIL FILTERS - MAINTENANCE
PRACTICES).
(5) Examine the compressor blade for damage or indication of erosion (Ref. ENGINE -
INSPECTION/CHECK).
(6) Do a borescope inspection of the compressor turbine blades and shrouds (Ref. ENGINE -
INSPECTION/CHECK).
(7) Do a borescope inspection of the power turbine blades (Ref. ENGINE -
INSPECTION/CHECK).
(8) Do a power assurance check (Ref. Applicable Rotorcraft Manual).
(9) Return engine to service if no damages are found.
(10) Drain and refill oil system with new oil (Ref. 72-00-00, ENGINE - SERVICING) 50 ± 10
flight hours after original oil change.
N. Contamination of the AFCU P3 (Pneumatic) System (Ref. Fig. 608)
Figure 608 AFCU P3 (Pneumatic) System - Inspection

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1. Compressor Delivery (P3) Tube Assembly

2. Compressor Delivery (P3) Tube Assembly
3. P3 Air Filter Housing Assembly
4. Spring
5. Preformed Packing
6. Element
7. Preformed Packing
8. Preformed Packing
9. Filter Cover
10. Nut
11. Loop Clamp
12. Bolt
13. Bracket
14. Loop Clamp
15. Loop Clamp Grommet
16. Bracket
17. Spacer

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Figure 609 AFCU P3 Inlet Ports - Inspection

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(1) Inspect and flush the P3 (Pneumatic) system for possible contamination as follows:
(a) Remove the P3 air filter housing cover and filter element (Ref. 73-10-07).
(b) Using a high intensity light, inspect for debris at the inner area of the filter housing
assembly. Check center tube for security. Replace filter housing assembly if center tube
is loose (Ref. SB5367).
(c) Remove the filter element (6) from the cover (9) and inspect inside cover for debris.
Clean if required.
(d) Place suitable container under the P3 filter housing assembly (3) to contain fluid during
flushing procedure.
(e) Disconnect the P3 tube assemblies (1 and 2) at the AFCU.
NOTE: Clamps and associated parts (10 thru 17) may be loosened or removed as
required to gain access to the fitting and prevent damage to the P3 line.
(f) Using a high intensity light and a mirror if required, inspect the P3 air inlet ports on the
AFCU (Ref. Fig. 609). Replace AFCU if internal debris is evident.
(g) Inject flushing fluid through the P3 tube(s), allowing the fluid to flow freely through the
P3 filter housing. Tube assembly (2) may have to be removed to be flushed adequately.
Inspect contained fluid for presence of debris (particles). Repeat flushing process until
fluid is clean.
NOTE: The compressor wash cart and solutions may be used to flush the P3 tube(s)
(Ref. 72-00-00, ENGINE - CLEANING) Ensure that the outlet pressure is
regulated to 10 psi. As an alternate method, manual injection (pouring) of
petroleum solvent (PWC11-027) in the tube(s) is acceptable.
(h) Replace AFCU if debris was found during flushing.
(i) Dry the interior of the P3 tubes using clean dry shop air or equivalent.
(j) Install P3 tube assemblies to the AFCU. Torque tube fittings 90 to 100 lb.in. and
lockwire.
(k) If clamps securing tube assembly (1) have been removed or loosened, proceed as
follows:
1 Install grommets (15) and loop clamp (14) and secure tube assemblies with bolt
(12), bracket (13) and nut (10).
2 Install spacer (17) between bracket (13) and loop clamps (11) and secure with bolt
(12) and nut (10).
3 Secure loop clamps (11) to bracket (16) with bolt (12) and nut (10).
4 Torque nuts (10) 32 to 36 lb. in.

(l) Install P3 filter element and cover assembly (Ref. 73-10-07, P3 AIR FILTER,
MAINTENANCE PRACTICES).

O. Non-preserved Engine Procedure

NOTE: 1. If the preservation of the inactive engines and/or modules are not done per the EMM
(Ref. Servicing), the procedures that follow are necessary. The day ranges listed
below show the number of days the engine or engine module remained inactive
without the benefit of preservation.
NOTE: 2. All preservation related maintenance activity, including the routine verification of
humidity indicators per the EMM schedule, must be supported by entries in the
engine logbook. If you do not record these data in the engine logbook, it is assumed
that the preservation procedures were not done.
(1) 0 to 7 days:
(a) If the engine operated in a salt laden environment, do a compressor desalination wash
and a turbine wash (Ref. Cleaning).

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(b) Make sure that all the engine openings are clear, no blockage and there is no signs of
corrosion. Examine the engine externals for general condition and signs of corrosion.
(c) If you find discrepancies, do the necessary steps per EMM procedures.
(d) If the engine is scheduled to remain inactive, do the preservation procedure for the
possible period of inactivity (Ref. Servicing). Record the maintenance work in the
logbook.

(2) 8 to 28 days:
(a) Do the 0 to 7 days requirement.
(b) Operate the helicopter on ground for a minimum of 20 minutes and pull sufficient
power to maintain "light on gear" condition.
(c) Collect oil sample, then send it for oil AT analysis.
(d) Examine the oil filter element for the presence of debris.
(e) Do a check of the engine for the signs of oil and fuel leakage.
(f) If you find discrepancies, do the necessary steps per EMM procedures.
(g) If the engine or module is scheduled to remain inactive, do the preservation procedure
for the possible period of inactivity (Ref. Servicing). Record the above maintenance
work in the logbook.
(h) For inactive engines that stored in harsh environmental conditions such as too much
temperature changes or high humidity (more than 40%) or pollution and/or salt laden
environment, in addition to the above procedures:
1 Do a total acid number (TAN) test of the oil in the engine oil tank. If the TAN is more
than the oil manufacturer limits, then drain and flush the oil system.
NOTE: 1. Engine oil with high TAN can possibly cause corrosion to the metal
surfaces.
NOTE: 2. The TAN limit is given by the engine oil manufacturer. Refer to the oil
brand specification sheet.
2 Do an inspection of the inner side of the AGB and RGB as much as possible with a
borescope for the signs of corrosion or flaking/separation of the internal coating. If
you find corrosion or flaking/separation of the internal coating or you cannot do
borescope inspection, remove the engine and do light overhaul for non-preserved
engine overhaul level inspection.
a If the engine or module is scheduled to remain inactive, do the preservation
procedure for the possible period of inactivity (Ref. Servicing). Record the above
maintenance work in logbook.

(3) 29 to 90 days:
(a) Do the 0 to 7 days and the 8 to 28 days requirements.
NOTE: Engine preservation related activity including the routine monitoring of the
humidity level must be recorded in the engine log book.
(b) If humidity level controlled and monitored at 40% or less, then do the steps that
follow:
1 Do a TAN test of the engine oil.
NOTE: 1. Engine oil with high TAN can cause corrosion to the metal surfaces.
NOTE: 2. The TAN limit is given by the engine oil manufacturer. Refer to the oil
brand specification sheet.
2 If the TAN is more than the oil manufacturer limits, then drain and flush the oil
system.

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3 Do an inspection of the inner side of the AGB and/or RGB as much as possible with a
borescope for corrosion and flaking of internal coating. If you find corrosion or signs
of flaking or you cannot do the borescope inspection, remove the engine/RGB and do
an overhaul level inspection.
4 Do a functional check of the mechanical fuel control arm. If fuel control lever do not
move easily, replace the fuel control components.

(c) If the engine is scheduled to remain inactive, do the preservation procedure for the
possible period of inactivity (Ref. Servicing). Record the above maintenance work in the
logbook.
(d) If the engine exposed to the conditions such as humidity levels more than 40%, harsh
environment and/or if the engine stored outside without monitoring the humidity level,
in addition to the procedures given before, do the steps that follow:
1 Replace the oil filter, then examine the chip detector.
2 Examine the fuel filter for the signs of fungus or slime. If you find the fungus or
slime, replace the fuel pump, Manual Fuel Control, Automatic Fuel Control and the
flow divider. If the engine operated with slime in the fuel system, do a functional
(flow) check of the fuel nozzles.
3 Disconnect the fuel tube at the inlet of the flow divider. Do a wet motoring cycle to
flush the fuel system.
4 Examine the engine externals for general condition and signs of corrosion.
5 Examine the compressor for corrosion and cleanliness. If dirty, do a performance
recovery. If salt laden, do a desalination wash.
6 Turn the rotor with your hand and examine for binding. If you find the binding or
resistance, find the cause. If you cannot find the cause, send the engine to an
overhaul shop for a light overhaul for non-preservation.
7 Operate the helicopter on ground for a minimum of 20 minutes at sufficient power to
maintain "light on gear" condition. Do a check for oil and fuel leakage.
8 Collect a ample of the engine oil, then send it for oil AT analysis. When you send the
engine back to the service, collect samples for oil AT analysis at an interval of 50
hours up to a maximum of 250 hours. After 250 hours if results show no signs of
debris generation, do the sampling at routine intervals.
9 If the engine is scheduled to remain inactive, do the preservation procedure for the
possible period of inactivity (Ref. Servicing). Record the above maintenance work in
the logbook.

(4) For period more than 90 days but less than one year:
(a) If the fuel system not preserved per Servicing for a period of six months or more,
remove the FMM/FCU, flow divider and and send to an overhaul shop for overhaul level
inspection.
(b) If the humidity is more than 40% or not monitored routinely per engine preservation
procedures, send the engine for a light overhaul for non-preservation.
(c) If the humidity recorded and being routinely monitored per the engine preservation
procedures requirement and it is not more than 40% during the period of inactivity,
then do the 0 to 7 days, 8 to 28 days and 29 to 90 days requirement.
(d) If the engine is scheduled to remain inactive, do the preservation procedure for the
possible period of inactivity (Ref. Servicing). Record the above maintenance work in the
logbook.

(5) For period of one year and more, regardless of humidity and environment condition, send
the engine to an approved overhaul facility for an overhaul level inspection for non-
preserved engine.
11. Borescope Equipment and Procedures

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A. General

CAUTION: HEAT CAN DAMAGE THE BORESCOPE. ENGINE TEMPERATURE MUST BE

LESS THAN 66°C (150°F) BEFORE YOU DO AN INSPECTION. THE
SATISFACTORY COOLING PERIOD FOR STANDARD AMBIENT CONDITIONS
IS 40 MINUTES (COLD SECTION) OR TWO HOURS (HOT SECTION) AFTER
ENGINE SHUTDOWN, BUT CAN CHANGE WITH AMBIENT CONDITIONS. IF
REQUIRED, DO DRY MOTORING RUNS TO ACCELERATE COOLING (REF. 71-
00-00, POWER PLANT - ADJUSTMENT/TEST).

The borescope is an optical device which enables an operator to perform visual specific
inspection of hot section areas of the engine as required to detect sulfidation due to local
environmental conditions, wear or damage, etc. Access is through ports or openings created by
removal of engine components. Personnel performing borescope inspection must be qualified to
do checks and analyse results.
B. Description
(1) A flexible videoscope or video borescope is an advanced type of borescope that contains a
very small chip embedded into the tip of the scope. It lets an operator examine the internal
areas of the engine without removal or disassembly of the engine. It sends the video image
from the distal tip and focusable lens assembly back to the display through internal wiring.
The diameter must not be more than 5 mm or it will bind in the guide tube.
(2) An operator can do the periodic inspection with the borescope (PWC34910-109) of the
following components:
(a) Compressor turbine (CT) blades and shroud segments.
(b) Leading and trailing edges of the CT vane ring.
(c) Inner and outer vane rings of the CT vane ring.
(d) Cooling rings and dome section of the combustion chamber liner.
NOTE: 1. Before you use the borescope, study the following procedures, the
borescope assembly and accessories (Ref. Fig. 610).
NOTE: 2. If you twist or pinch the borescope, it can be easily damaged and can cause
dangerous shocks. If you handle the borescope carefully, it will have a long
service life.

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Figure 610 Borescope and Accessories Installed (Typical)

icn-00198-g000022580-001-01

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(3) Basic Installation

(a) Remove one fuel manifold adapter. The maximum suggested working temperature is
66°C (150°F).
(b) Insert guide tube (PWC34910-200) through open manifold adapter port.

(c) DELETED.
(d) Put the borescope (PWC34910-109) into the guide tube (PWC34910-200).
(e) DELETED.
(f) DELETED.
(g) Remove starter-generator from accessory gearbox and attach wrench (PWC34941) to
splined shaft to be able to turn the compressor.

(h) DELETED.

(4) Various tip adapters are available for borescope. Refer to the borescope equipment
manufacturer for installation and removal instructions.
(a) DELETED.

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Figure 611 Side-viewing Adapter - Removal/Installation

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(b) DELETED.

(5) DELETED.
(6) DELETED.
(7) Guide Tube

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(a) Installation (Ref. Fig. 612)

Figure 612 Guide Tube Orientation

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1 Remove the applicable fuel manifold adapter(s) (Ref. 73-10-05, FUEL MANIFOLD
ADAPTERS - MAINTENANCE PRACTICES).
NOTE: The rigid, pattern-controlled guide tube must be engaged freely, no force
must be used.
2 Ease guide tube (PWC34910-200) through fuel manifold adapter port into
combustion chamber and exit duct zone turning guide tube counterclockwise to
achieve a three-quarter turn.
3 Installed, guide tube end locates between vanes of compressor turbine stator, while
supporting flange rests on adapter boss. Secure flange to boss.

(b) Removal (Ref. Fig. 612)

1 Loosen knurled screw to release guide tube supporting flange.
2 Withdraw guide tube, turning clockwise.

3 Install the fuel manifold adapter(s) (Ref. 73-10-05, FUEL MANIFOLD ADAPTERS -
MAINTENANCE PRACTICES).

(8) Troubleshooting
(a) The possible sources of, and solutions for, problems encountered when using borescope
are shown in Table 602.

C. Procedure
(1) Examine compressor turbine blades as follows:
(a) Install guide tube (Ref. Subpara. B.(7) ).

(b) DELETED.
(c) DELETED.
(d) DELETED.
(e) Refer to the borescope manufacturers documentation for possible sources of remedies
for problems found while you use the borescope.
Table 602 Borescope Troubleshooting

PROBLEM POSSIBLE SOURCE SOLUTION

DELETED.

CAUTION: MAKE SURE THAT THE ENGINE TEMPERATURE IS BELOW 60 °C (140

°F).
(f) Slowly put the borescope (PWC34910-109) into the guide tube. Stop engaging the
borescope (PWC34910-109) immediately distal tip touches the end of the guide tube.
NOTE: Make sure that the distal tip actuating lever is in the neutral position when you
install or remove the borescope.
(g) Remove starter-generator from engine accessory gearbox (Ref. Aircraft Maintenance
Manual).
(h) Attach wrench (PWC34941) to splined gearshaft.
(i) Loosen knurled knob on holding fixture.

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(j) Examine CT blade tip, root and air foil section (Ref. Views A and B, Fig. 613 and Para.
13. J.).
Figure 613 Borescope Views (Simulated)

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CAUTION: MAKE SURE THAT THE DISTAL TIP OF BORESCOPE IS NOT INSTALLED
BETWEEN COMPRESSOR TURBINE BLADES BEFORE YOU TURN THE
COMPRESSOR.
(k) Continue examination of the CT blades while a second operator, with the wrench, turns
the compressor.

(l) Retract the borescope (PWC34910-109) into the guide tube.

(m) Remove the wrench from splined gearshaft and install the starter-generator (Ref.
Aircraft Maintenance Manual).

(n) Remove the borescope (PWC34910-109) from the guide tube.

(o) Remove the borescope (PWC34910-109).
(p) Remove guide tube (Ref. Subpara.B. (7) ).

(2) Examine CT stator assembly as follows:

(a) Install guide tube (Ref. Subpara. B. (7) ).

(b) DELETED.
(c) DELETED.

CAUTION: MAKE SURE THAT THE ENGINE TEMPERATURE IS BELOW 60 °C (140

°F).
(d) Put the borescope (PWC34910-109) into the guide tube.
NOTE: The use of side viewing adapter can add to the inspection capability (Ref.
Subpara. B. (4) ).
(e) Examine vane leading and trailing edges, inner and outer rings of vane ring assembly
(Ref. Views C and D, Fig. 613 and 13. G.).
NOTE: To photograph area, refer to Subparagraph B. (6) .
(f) Remove borescope and holding fixture.
(g) Remove guide tube (Ref. Subpara. B.(7) ).

(3) Examine combustion chamber liner assembly as follows:

CAUTION: MAKE SURE THAT THE ENGINE TEMPERATURE IS BELOW 60 °C (140

°F).
(a) Remove the fuel manifold adapters as necessary (Ref. 73-10-05, FUEL MANIFOLD
ADAPTERS - MAINTENANCE PRACTICES).

(b) DELETED.
(c) Put the borescope (PWC34910-109) into a fuel manifold adapter port.
(d) Examine the combustion chamber liner (Ref. View E, Fig. 613 and Para. 13.D.).

(e) Remove the borescope (PWC34910-109).

(f) Install fuel manifold adapters (Ref. 73-10-05).

12. In-service Borescope Inspection

A. Hot Section Components

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In-service inspection is used to determine performance deterioration. Parts can be inspected

using a borescope (Ref. Para. 11.) or by removing the power section (Ref. Para. 13.) following.
Refer to FAULT ISOLATION for areas to be inspected.
Using a borescope is more practical for this inspection. Removal and installation of the power
section, components/assemblies may result in damage to parts and/or distortion of sealing
surfaces causing gas leaks and performance loss.
Acceptable damage must be documented and progression monitored. Record engine TSN/TSO,
fuel nozzle ports used, if found by borescope, component, description, location and dimensions
of defect.
If rotating components are found with unacceptable damage during inspection, an HSI must be
done before next flight. If damaged non-rotating components are found, except when holes are
burnt through the combustion chamber or the CT vane airfoil trailing edge defects are beyond
limits (Ref. Inspection), the HSI may be delayed, providing an engine performance/ground
power check is done.
An additional engine performance/ground power check and inspection of the affected area must
be done after 50 hours. Subsequent inspections and engine performance/ground power check
intervals will depend on the progression and level of deterioration. Keeping an engine in service
after components have deteriorated may substantially increase the cost of future
repairs/refurbishments.
Gas path components downstream of components having material missing must be inspected
for secondary damage. Pay special attention to rotating components. CT vanes burned through
at trailing edge (pressure and suction sides) may mean replacement of the complete set of CT
blades, depending on the area of the surface burned through.
The ability of an engine to produce the power required by the flight manual power assurance
check is the only engine airworthiness requirement. If, at an anticipated high ambient
temperature and/or altitude, T5 and Ng will approach or exceed the maximum limits, an HSI is
recommended before the actual conditions occur. When T5 and Ng approach, or are anticipated
to approach or exceed the maximum limits, troubleshoot the engine/installation before doing an
HSI.
Carbon accumulation inside fuel nozzle passages is the principle cause of spray pattern
degradation resulting in non-uniform combustion and local high temperature peaks. Exposure to
these peaks contributes to premature hot section deterioration. Carbon accumulation is
progressive and can affect all nozzles. Therefore, inspect all nozzles (Ref. 73-10-05,
Inspection/Check), to minimize premature deterioration occurring at other locations.
If component deterioration exceeds limits, either replace the individual component or do an HSI,
depending on the general condition of the hot section and engine performance.
The inspections are recommended concurrent with scheduled engine maintenance checks
applicable to individual aircraft installations.
B. Inspection of Combustion Chamber
NOTE: Combustion chamber includes small exit duct, inner and outer combustion chamber liner
assemblies.
(1) Combustion chamber liner cracks may be repairable, and an HSI is recommended before
component replacement is required. Converging cracks in the inner and outer combustion
chamber liner walls which do not meet are acceptable. Inspect affected area at 100 hours.
Subsequent inspections must not exceed 400 hours.
(2) Plasma top coating (ceramic) loss revealing undercoat (diffused aluminide) on outer and
inner liners is acceptable, providing base metal is not burnt or eroded.
(3) Small areas (approximately 1 sq.in.) bulging and/or hot spots on inner or outer liner walls
may be repairable and an HSI is recommended. Bulging and/or burning in the dome area,
associated or not with axial cracks, are acceptable provided the axial cracks (circumferential
cracks are unacceptable) do not exceed 1.0 inch in length or 0.030 inch in width. Engine
may remain in service, providing the associated fuel nozzle(s) is/are inspected (Ref. 73-10-
05, Inspection/Check) and replaced if not within limits. Subsequent inspections are at
operators discretion, but must not exceed 400 hours.

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(4) Holes in the inner and outer liner walls are unacceptable. An HSI must be done to replace or
repair the affected component, and the associated fuel nozzle(s) inspected (Ref. 73-10-05,
Inspection/Check) and replaced if not within limits.
(5) Cracked or distorted cooling rings on liners may be repairable. An HSI is recommended
before cooling rings are burned through and pieces enter the gas stream. Engine may
remain in service if it is understood that cooling air flow is changed, and the rate at which
the combustion chamber deteriorates may increase. Therefore, if an HSI is not done,
combustion chamber liner with distorted cooling rings or converging cracks must have the
associated fuel nozzle(s) inspected (Ref. 73-10-05, Inspection/Check) and replaced if not
within limits. Inspect damage at 100 hours. Subsequent inspections are at operators
discretion, but must not exceed 400 hours.
(6) Excessive carbon deposits inside combustion chamber could be result of poor fuel
atomization by the fuel nozzles (indicated by deposits around fuel nozzle bosses). Distortion
of combustion chamber liner cooling rings may produce carbon deposits in the dome area
downstream of the affected cooling ring. If excessive carbon deposits are found, flow check
fuel nozzles (Ref. 73-10-05, Inspection/Check). If CT blades are eroded, refer to Inspection
of CT Blades.
(a) If the nozzles are serviceable, the combustion chamber liner is the probable cause of
the deposits. Inspect the CT blades, as erosion by carbon particles (some carbon
particles remain in the gas stream and are not deposited in the combustion area) may
damage and cause CT blade replacement.
(b) If CT blades are not eroded, engine may remain in service and the CT blades and
combustion chamber liners inspected within 100 hours. Subsequent inspections are at
operators discretion, but must not exceed 400 hours.

(7) As the structural integrity of the small exit duct is not affected, cracks and open radial
cracks extending from the inner to the outer diameter are acceptable (length not limited),
and an unlimited amount of coating loss is acceptable.
(a) Holes less than 0.500 in. in diameter in the outer wall are acceptable.
(b) If holes or open cracks are found, an HSI is recommended.
NOTE: The engine may remain in service, and HSI delayed, providing it is understood
that the CT stator cooling air flow is affected, increasing the rate of
deterioration.
(c) When damage is found, inspect the CT vane ring (Ref. 72-50-01, Inspection/Check) and
associated fuel nozzle (Ref. 73-10-05, Inspection/Check).
(d) Inspect damage at 100 hours. Subsequent inspections are at the discretion of the
operator, but must not exceed 400 hours. If deterioration exceeds the above limits, an
HSI is recommended.

(8) Cracks along, or across, the dome to outer liner seam weld are acceptable provided the
cracks do not intersect the fuel manifold support bracket. Inspect damage at 100 hours.
Subsequent inspections are at the operators discretion but must not exceed 400 hours.
C. Inspection of Compressor Turbine (CT) Vane Ring Assembly

CAUTION: IF ANY CT STATOR VANE HAS TRAILING EDGE BURN THROUGH

EXCEEDING THE FOLLOWING LIMITS; THE CT BLADES MUST BE
DISCARDED.
(1) Measure trailing edge burn through (Ref. Fig. 614).
NOTE: Burned areas on vanes increase flow area which accelerates downstream component
deterioration, decreases Ng and increases T5.
(2) Cracks (Ref. Fig. 615) are repairable. Keep the engine in service and monitor deterioration
progression. Do an HSI before defects progress to such an extent that the CT stator
becomes unrepairable. This will increase the cost of the subsequent HSI. Inspect CT stator
having the defects shown within 100 hours. Subsequent inspections must not exceed 400
hours.

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(3) If defects have progressed beyond repair limits shown, engine may remain in service,
providing an engine ground power check (Ref. Aircraft Maintenance Manual) is done, and
the airfoil trailing edge damage does not exceed limits shown (Ref. Fig. 614). An HSI is
recommended if the vane trailing edge defects are beyond the limits shown or downstream
components are affected by the CT vane distress or engine performance is unacceptable.
NOTE: Refer to Subpara. A. for additional recommendations.

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Figure 614 CT Stator Damage - HSI Recommended

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Figure 615 CT Stator - Acceptable/Repairable Damage

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D. Inspection of CT Blades (Ref. Figs. 616, 617 and 618)

Figure 616 CT Blade Damage - Unacceptable Damage

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Figure 617 CT Blade Damage - Acceptable Damage

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(1) During compression turbine blade inspection, operators can incorrectly identify some surface
damage as sulfidation. These surface damages are result of the blade manufacturing
process and it will not effect turbine performance.
(2) Geometrical deviations must not be more than 0.005 in. deep or 0.005 in. high. There is no
width limitation (Ref. Fig. 618).

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Figure 618 CT Blade - Manufacturing Anomalies

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(3) The condition of blades and tips is critical to obtain rated power. Most significant blade tip
defects (rubs and oxidation) increase interturbine temperature (T5). Even if T5 is below
maximum, an HSI may be recommended if damage is beyond the limits shown (Ref. Fig.
617).
NOTE: Refer to Subpara. A. for additional HSI recommendations.
(a) Defects shown are acceptable, providing their condition is monitored by further
inspections and engine performance checks.
(b) Subsequent inspections are at the discretion of the operator, but must not exceed 400
hours.
(c) Blades with cracks 0.050 in. long in the upper 1/3 of the trailing edge may remain in
service for 100 hours.
(d) If defects are beyond those shown (Ref. Fig. 616), an HSI is recommended.

(4) Check blades for axial shift. Normally, with components within assembly tolerances, the
blade platforms are approximately in line. When a blade shifts, the blade moves axially and
can be seen as having moved in relation to the adjacent platform.
NOTE: 1. Checking each CT blade TE platform alignment with a borescope is quite difficult.
The recommended method is to look at the leading edge (LE) mismatch while
rotating the CT disk. When a mismatch is observed (one of the airfoil LE is
shifted), check blade trailing edge platform. Maximum shift is 0.020 inch.
NOTE: 2. Use 0.020 in. lockwire to examine the value of blade shift. The lockwire is wind
tightly around the tip of borescope, then attach with the tape. The position of the
free end must be within the field of view of the borescope. Make sure that the
lockwire is correctly attached. If the lockwire is loose, it will fall into the hot
section. When you calculate the value of shift, position the free end of the
lockwire adjacent to the platform being examined.
(5) If blade shift in excess of the limit is observed, an HSI is
recommended.
E. Inspection of CT Shroud Segments
(1) Heavy rubbing and oxidation are acceptable, providing T5 is within limits. Operating the
engine with the shroud segments burned may cause damage to the CT shroud housing
(cracking and burning of the attachment rim), and an HSI is recommended. A damaged
shroud housing must be replaced at the next HSI or refurbishment.
NOTE: Refer to Subpara. A. for additional HSI recommendation.
F. Inspection of Power Turbine (PT) Stator
NOTE: Inspection of the PT stator is recommended when upstream component damage does
not explain performance loss or when secondary damage is suspected.
(1) Damage on the vanes may produce an increase in flow area which will increase Ng and
T5.
(2) Cracks on the inner and outer rings and vanes are repairable. Keeping the engine in service,
the defects will progress until the stator becomes unrepairable. This will increase the cost of
the subsequent HSI. Inspect the damaged area within 100 hours. Subsequent inspections
must not exceed 400 hours.
(3) An HSI is recommended when the defects are still
repairable.
(4) If the defects are not repairable, replace PT stator at the next power section repair.
Providing Ng and T5 are within limits, there is no need for power section repair and PT
stator change, regardless of the amount of damage, unless structural integrity of the vanes
are affected (e.g. wide open cracks, excessive foreign object damage (FOD) and missing or
burnt material are unacceptable).
G. Inspection of PT Blades

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(1) Increased tip clearance of PT blades increases T5. If T5 is within limits, there is no need to
change the PT assembly, regardless of the amount of damage, providing structural integrity
of the components is not affected.
(a) Cracks, missing material, excessive FOD, heavy sulfidation or blade distortion are
unacceptable.

13. Hot Section Inspection

A. General
The inspection procedures given in the following paragraphs cover the entire scope of the engine
hot section inspection. The complete inspection procedure is performed whenever power
sections have been subjected to overtemperature conditions, or on a periodic basis (Ref.
SB5003) .
Basically, the procedure involves partial disassembly of the engine to gain access to the hot
section components, inspection of the components and repair or replacement as applicable,
followed by re-assembly.
Subparagraph B. lists the referenced procedures for disassembly to gain access. Table 603 lists
the hot section components to be inspected. The table also indicates if repair or replacement
procedures are given in this manual and the Chapter/Section/Units in which these procedures
are to be found.
Paragraphs C. through Q. give the component inspections and associated acceptance, repair and
replacement criteria.
B. Preparation
(1) Remove power section (Ref. 72-00-
01).
(2) Remove power turbine and exhaust section 72-50-
00.
(3) With assembly loaded to produce smallest gap, measure and record compressor turbine
blade tip clearance at the center and ends of each shroud segment using a wire or tapered
feeler gage, as follows:
(a) Pre-SB5211 (New Segments): Average clearance for the new segments must be
between 0.013 and 0.015 in. measured and loaded radially outward.
(b) Pre-SB5211 (Used Segments): The tip clearance at any one location must not be
more than 0.020 in. or less than 0.011 in.
(c) Post-SB5211 and PT6T-3BE (New Segments): Average clearance for the new
segments must be between 0.016 and 0.018 in. measured and loaded radially outward.
(d) Post-SB5211 and PT6T-3BE (Used Segments): The tip clearance at any one location
must not be more than 0.022 in. or less than 0.015 in.

(4) Remove compressor turbine disk (Ref. 72-50-

02).
(5) Remove fuel nozzles (Ref. 73-10-
04).
(6) Remove igniter plugs (Ref. 74-20-
00).
(7) Remove combustion chamber liner (Ref. 72-40-
01).
(8) Remove combustion chamber large exit duct (Ref. 72-40-01, COMBUSTION CHAMBER
LINER ASSEMBLY - MAINTENANCE PRACTICES).
C. Gas Generator Case Assembly
(1) Inspect case for cracks (Ref. Step (2) ), distortion, overheating or heavy corrosion. Repair
of case is not authorized at maintenance level. If condition of case warrants repair, power

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section must be shipped to an overhaul facility for inspection/repair in accordance with the
Overhaul Manual.
(2) Using a high intensity light, 10X magnifying glass and mirror if necessary, visually inspect
gas generator case for cracks. Inspect weldments at bosses and spot welds(Areas a) on
surface of case for cracks (Ref. Fig. 619).

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Figure 619 Gas Generator Case Inspection (Typical)

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CAUTION: VOIDS OR SURFACE SCRATCHES MAY SHOW AS CRACKS AND

THEREFORE SHOULD NOT BE MISCONSTRUED AS CRACKS.
(a) If a crack is confirmed, power section must be removed and sent to overhaul facility for
inspection/repair in accordance with the Overhaul Manual.

(3) Inspect case for protective coating loss. If more than 25% of surface area is affected, return
case to overhaul facility for strip and re-coat. If less than 25% of surface area is affected,
touch-up (Ref. 72-30-04).
Table 603 Hot Section Inspection - Component Repair/Replacement

Component HSI Para. Repair Replacement Chap/

Sec/
Unit

Gas Generator Case Assy 13. C. X N/A 72-30-04

Combustion Chamber Liner 13. D. X X 72-40-01

Combustion Chamber Large Exit 13. E. N/A X 72-40-01

Duct

Combustion Chamber Small Exit 13. F. N/A X 72-40-01

Duct

Compressor Turbine Vane Ring 13. G. X X 72-50-01

Compressor Turbine Shroud 13. H. X X 72-50-01

Segments

Compressor Turbine Shroud 13. I. X X 72-50-01

Housing

Compressor Turbine Disk Assy 13. J. N/A X 72-50-02

Power Turbine Disk Assy 13. K. N/A N/A 72-50-04

Power Turbine Vane Ring 13. L. X X 72-50-03

T5 Temperature Sensing System 13. M. X (Pre-SB5118) X 77-20-01

Exhaust Duct 13. N. N/A N/A -

Fuel Nozzles and Sheaths 13. O. N/A 73-10-05

Fuel Manifold Bolts 13. P. X X 72-30-04

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(4) Inspect diffuser pipes in gas generator case (Ref. Fig. 620).
Figure 620 Diffuser Pipe Inspection (Typical)

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(a) Cracks of any length on inner surface of any number of diffuser pipes are acceptable
provided that they will not converge and cause loss of metal (Ref. Details A, B and C)
and that they are stop-drilled.
(b) Cracks terminating in fishtail seam weld are acceptable and require no repair.
(c) Loss of material is acceptable on a maximum of four diffuser pipes provided that
detached material is removed from engine and dimensions of hole after blending are
within limits (Ref. Details A and B).
(d) Cracks in fishtail outer face are acceptable after stop-drilling.
(e) Fretting wear on outer face and edge adjacent to next fishtail is acceptable after blend
repair.
(f) Cracks in fretted areas are acceptable if stop-drilled.

(5) Reject engine if two adjacent diffuser pipes are missing or if more than two nonadjacent
diffuser pipes are missing.
(6) An engine missing a maximum of two nonadjacent diffuser pipes may continue in service
until the next scheduled removal for repair or overhaul provided all other diffuser pipes are
within allowable limits, performance of the engine remains acceptable and engine is surge
free. Remove debris and sharp edges, and borescope hot section and combustor 150 hours
after discovery. A maximum of 300 hours of operation is permitted with this condition, after
which the engine must be removed for repair.
NOTE: Operating with missing diffuser pipes may lead to accelerated hot section
deterioration. If engine is kept in service, perform borescope inspection of hot
section and combustion chamber at the recommended 150 hour interval.
(7) Cracking of diffuser pipes may be associated with vibrations. The source of vibration should
be identified and corrective action taken.
(8) Inspect threaded holes and shanknuts in center bore area for damage (Ref. 72-30-
04).
D. Combustion Chamber Liner
NOTE: Cracks in the liner surfaces are usually of a stress-relieving nature, and as such are not
serious in that the rate of growth decreases as the crack lengthens. Thermal stresses,
in effect, relieve original stress conditions. It is considered normal to observe a given
type of deterioration repeated from liner to liner. Typical liner distress consists of
buckling and cracking of cooling rings and buckling at the inner wall adjacent to the
dome.
(1) The following conditions are acceptable without repair (Ref. Fig. 621).

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Figure 621 Combustion Chamber Liner Cooling Ring - Inspection

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(a) Localized buckling and/or burning of all cooling rings (except louvered type)
accompanied by cracking, provided cracks do not extend into seam weld.
(b) Cracks in louvered cooling rings; provided that circumferential cracks in seam weld do
not exceed 0.300 in. in length and are stop-drilled.
(c) Straight-line cracks between two adjacent cooling holes.
(d) A maximum of seven cracks (each crack not exceeding one inch in length) in inner liner
adjacent to dome end.

CAUTION: CRACKS CONVERGING TOWARDS A COMMON POINT ARE NOT

ACCEPTABLE UNLESS THEY ARE SEPARATED BY AT LEAST THREE
INCHES OF SOUND MATERIAL AND ARE STOP-DRILLED.
(e) Localized areas that have been heated to an extent to buckle the liner, provided
buckling is shallow and is not associated with burning effecting a reduction in wall
thickness causing structural weakening.

(2) Acceptable conditions requiring repair are as follows:

(a) Buckled cooling rings, in which the gap has been eliminated, must be reworked. Use
suitable bar to restore uniform gap.
(b) Cracks not exceeding two inches in length must be welded.
(c) Circumferential cracks adjacent to seam welds must be repaired by welding.
NOTE: Cracks open in excess of 0.030 in. are not repairable.
(d) Cracks progressing from a free edge so that their meeting is imminent and could allow
a piece of metal to break loose must be repaired by welding.

(3) If any of the following conditions exist, the liner must be replaced.
(a) Severe buckling causing "kinked" metal.
(b) Multiple cracks (several small cracks propagating from one main crack).
(c) Cracks exceeding two inches in length.
(d) Circumferential cracks adjacent to seam welds which are opened in excess of 0.030
in.

E. Combustion Chamber Large Exit Duct

Inspect large (outer) exit duct for cracks, burning, distortion or fretting wear on the inner
diameter which contacts the combustion chamber liner. Any amount of coating loss in contact
area is acceptable. Coating loss in other areas acceptable provided parent metal erosion (if
present) is not in excess of 0.010 in. deep. No repairs authorized at maintenance level.
F. Combustion Chamber Small Exit Duct
(1) Inspect small exit duct for fretting wear at duct flange which contacts combustion chamber
liner. Wear up to 0.015 inch is acceptable. If wear is in excess of 0.015 inch, duct should be
sent to an overhaul facility (Ref. Para. 10. A.) for possible repair.
(2) Cracks up to 1.00 inch long in small exit duct outer wall are acceptable provided they are
stop-drilled using 1/16 (0.0625) inch drill.
(3) Coating loss of any amount is acceptable on both small and large exit ducts provided there
is no evidence of burning or erosion of parent metal.
G. Compressor Turbine Vane Ring (Ref. Fig. 622)
Figure 622 Compressor Turbine Vane Ring - Inspection

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(1) Inspect vane ring for cracks, coating loss, erosion of parent metal or impact damage. The
following acceptance limits apply. No repairs are permitted.
(a) Vanes
1 Multiple open cracks not exceeding 0.005 in. wide and up to 1.0 inch long in the
aerofoil is acceptable, provided there is no evidence of burning.
2 If two or more cracks converge to liberate a piece of metal, then the component
should be rejected.

(b) Outer Platform (Ring)

1 Multiple open cracks not exceeding 0.020 inch wide and up to 1.0 inch in length are
acceptable in the outer platform downstream of its mounting flange. No two or more
cracks may converge in such a manner that would allow a piece of metal to become
detached.
2 Outer ring open cracks which extend into the aerofoil are acceptable up to 1.0 inch in
length, provided no two or more cracks converge in such a manner that would allow
a piece of metal to become detached.

(c) Inner Platform (Ring)

1 Multiple open cracks through the wall of the inner ring less than 0.020 inch wide and
up to 1.0 inch in length are acceptable provided that 20% of the axial length of the
inner ring is intact.
2 Inner ring open cracks which extend into the aerofoil are acceptable up to 1.0 inch in
length, provided no two or more cracks converge in such a manner that would allow
a piece of metal to become detached.

(2) Coating Loss and Erosion

(a) Up to 0.250 square inch on any vane airfoil surface is acceptable provided that erosion
(if any) does not exceed 0.010 in. deep.
(b) Any amount of coating loss from both inner and outer rings is acceptable provided
erosion (if any) does not exceed 0.020 in. deep.
(c) Oxidation: up to 0.250 square inch by 0.005 in. deep on any vane LE . Maximum of
seven affected vanes is acceptable.

(3) Burning
(a) Any number of vanes which exhibit burn through on the trailing edge of 0.125 square
inch or greater (area of approx. 0.350 inch x 0.350 inch) and the void may be seen
when viewed from the leading edge in direction parallel to engine longitudinal axis,
requires the replacement of both the vane ring and the complete set of compressor
turbine blades.

H. Compressor Turbine Shroud Segments

(1) Inspect shroud segments for cracks, distortion, erosion or metal build up.
(a) Cracks are not acceptable. Replace segments.

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Figure 623 Compressor Turbine Shroud Segments - Inspection

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(b) If blade-to-shroud clearances for used segments are less than 0.011 in. because of the
shroud high-spot or if the tip rubs, then the segment can be grounded locally to a
length not more than one inch per segment. (Ref. Fig. 623).
(c) Post-SB5211) and PT6T-3BE: If blade-to-shroud clearances for used segments are
less than 0.015 in. because of the shroud high-spot or if the tip rubs, then the segment
can be grounded locally to a length not more than one inch per segment. (Ref. Fig.
623).

I. Compressor Turbine Shroud Housing

Figure 624 Compressor Turbine Shroud Housing - Inspection

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(1) Inspect seal ring groove for distortion and fretting wear. Minor wear is acceptable provided
high spots are removed by light stoning.
(2) Inspect compressor turbine shroud housing for cracks (Ref. Fig. 624).
(a) Pre-SB5211 - Cracks are acceptable in shroud segment retaining lip up to 0.300 in. in
length provided bulging or buckling is not present.
(b) Post-SB5211 - Cracks are not acceptable. Replace housing.

(3) Inspect interstage sealing ring for scoring wear on outer face and sides. Ring wear is
acceptable to the following limits:
Width: 0.225 in. min.
Thickness: 0.042 in (Pre-SB5159)
Thickness: 0.084 in (Post-SB5159 and PT6T-3B/-3BE)
Gap: 0.175 in. min.

J. Compressor Turbine Disk Assembly

(1) Inspect blades for tip rub, erosion, corrosion, impact damage, coating loss, cracks or shift.
Inspect condition of blade retaining rivets.
(a) Damage limits for blades are as follows (Ref. Fig. 625 and 626):

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Figure 625 Compressor Turbine Blades - Inspection

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Figure 626 Stages of Corrosion

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NOTE: If any of the following limits are exceeded or any cracks are found in blades,
disk assembly must be replaced. Send removed disk assembly for overhaul
inspection (Ref. Para. 10. A.).
1 Limits for erosion of leading edge tip are as shown on figure.
2 Inspect blade airfoil surfaces, particularly concave surfaces, for corrosion and loss of
coating using 10X power magnification. A loss of up to 25% of coating is permissible.
Assess stage of deterioration and accept or reject bladed disk as follows:
a Stage 1 - Initial Coating Deterioration (Sulfidation): Indicated by light color
change of the part of coating area. Can be rust colored or dark gray. Coating has
deteriorated but is possibly not damaged. Blades are acceptable for service again.
b Stage 2 - Initial Corrosion: Indicated by apparent rise of sulfidated coating above
the adjacent surface, with small, scattered blisters seen in the coating. Corrosion
of the base material has started. Accept or reject the blades, as per the decision
of the operator, based on previous experience. If blades stay in service, the
operator must increase turbine washes and schedule regular borescope inspection
of the blades every 200 hours.
c Stage 3 - Advanced Corrosion: Evidenced by clusters of ruptured blisters exposing
bare material. Craters so formed deepen progressively, and crater surfaces darken
with glazed appearance. Send disk assembly to an approved overhaul facility for
blade replacement in accordance with the Overhaul Manual.

CAUTION: THIS CONDITION MAY RESULT IN IMBALANCE OF THE

COMPRESSOR TURBINE ASSEMBLY AND SUBSEQUENT DAMAGE.
d Stage 4 - Severe Corrosion: Deep penetration with large ruptured blisters
exposing large areas of bare material. Send power section to an approved
overhaul facility for disk repair and inspection of No. 2 bearing area in accordance
with the Overhaul Manual.
NOTE: Corrosion may be identified as metal loss or pitting, but more usually it
appears as local swelling or build-up due to greater volume occupied by
the nickel oxides. These corrosion products vary in color from green to
black and in the advanced state there will be associated flaking. Care
should be taken to distinguish between corrosion build-up and possible
light brown or rust colored deposits which are essentially harmless
combustion byproducts. The latter are more widespread over hot section
components and while possibly affecting performance, will not directly
affect blade integrity.

3 Impact damage in Area A: Three nicks, dents or pits no more than 0.005 in. deep
per blade.
4 Impact damage in Area B: One nick, dent or pit no more than 0.005 in. deep per
blade.
5 Impact damage on leading or trailing edges: One nick, dent or pit no more than
0.020 in. deep per blade.

(b) Blade Shift (Ref. Fig. 627)

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Figure 627 Turbine Blade Protrusion

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1 Blade protrusion beyond disk rim must be equal within 0.010 in. either side of disk.

(c) Circumferential Movement

1 Blade circumferential movement must not exceed 0.030 in. at tip.

(d) Axial Movement

1 No blade axial movement is permitted.

K. Power Turbine Disk Assembly

(1) Inspect blades for corrosion, impact damage, coating loss, cracks or shift.

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(a) Damage limits for blades are as follows (Ref. Fig. 628):
Figure 628 Power Turbine Blades - Inspection
(SHEET 1 OF 2)

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(SHEET 2 OF 2)

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NOTE: If any of the following limits are exceeded or any cracks are found in blades,
disk assembly must be replaced. As replacement of disk involves balancing
procedures, the complete power turbine and exhaust section must be sent to
an overhaul facility.
1 Examine blade airfoil surfaces for signs of corrosion and pitting as follows:
a Slight corrosion and/or closely grouped pits on the airfoil surface up to 0.002 inch
deep maximum is acceptable provided:
- the total affected area is not more than 25% of the airfoil surface.
- the corrosion appears as minor roughening of the air foil surface.

b Mild corrosion and/or closely grouped pits up to 0.005 inch deep maximum is
acceptable provided:
- the total affected area is not more than 10% of the airfoil surface.
- the corrosion appears as moderate roughening of the airfoil surface.
NOTE: Severe corrosion more than 0.005 inch deep is not acceptable: severe
corrosion appears as significant roughening of the airfoil surface.

2 Post-SB5079 blades only - carry out visual inspection of uncoated PT blade airfoil for
evidence of corrosion and pitting as follows:
a Slight corrosion and/or closely grouped pits on the airfoil up to 0.003 inch deep
maximum is acceptable provided:
- the total affected area of each surface is not more than 50% of the surface.
- the corrosion appears as minor roughening of the surface.

b Moderate corrosion and/or closely grouped pits up to 0.005 inch deep maximum is
acceptable provided:
- the total affected area of each surface is not more than 25% of the surface.
- the corrosion appears as moderate roughening of the surface.
NOTE: Severe corrosion more than 0.005 inch deep is not acceptable: severe
corrosion appears as significant roughening of the airfoil surface.

3 Impact damage in Area A: Three nicks no more than 0.015 in. long by 0.005 in.
deep. Three dents or pits no more than 0.010 in. deep per blade.
4 Impact damage in Area B: One nick, dent or pit no more than 0.020 in. deep per
blade.
5 Impact damage on leading or trailing edges: One nick, dent or pit no more than
0.020 in. deep per blade.

(b) Blade Shift (Ref. Fig. 627)

1 Blade protrusion beyond disk rim must be equal within 0.016 in. either side of disk.

(c) Blade-to-Shroud Rubbing

1 PT disk assembly must rotate freely without rubbing the PT shroud. If assembly does
not rotate freely, send complete power section to overhaul facility to inspect any
deformation of the PT shroud in accordance with the Overhaul Manual.

(d) Circumferential Movement

1 Blade circumferential movement must not exceed 0.030 in. at tip.

(e) Axial Movement

1 No blade axial movement is permitted.

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L. Power Turbine Vane Ring (Ref. Fig. 629)

Figure 629 Power Turbine Vane Ring - Inspection

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(1) Inspect vane ring for cracks, coating loss, erosion of parent metal or impact damage. The
following acceptance limits apply. No repairs are permitted.
NOTE: Power turbine stator assembly is classified. If replacement is required, install like
class.
(a) Cracks
1 Cracks up to 0.400 in. long are acceptable in leading or trailing edges on any number
of vanes.
2 Cracks up to 0.600 in. long are acceptable in outer ring.
3 Cracks up to 0.400 in. long are acceptable in inner ring.

(b) Erosion or pitting on airfoil surfaces.

1 Up to 0.250 in. square by 0.010 in. deep on any one vane for any number of vanes
is acceptable.
NOTE: Vane rings deemed unacceptable for continued service may be shipped to an
approved overhaul facility for possible repair in accordance with overhaul
manual instructions.

M. T5 Temperature Sensing System

(1) Check harness and connections for
security.
(2) On Pre-SB5118 engines check condition of bus bar terminal straps and lugs and ceramic
insulation on probes and terminal block. Loose terminals and broken or missing insulation
may be repaired as detailed in (Ref. 77-20-01).
(3) Carry out resistance and probe heat response checks (Ref. 77-20-
01).
N. Exhaust Duct
(1) On Post-SB5364 engines, ensure that turbine exhaust sealing ring has been discarded and
that the associated groove in the exhaust flange is undamaged.
NOTE: New sealing ring is required at reassembly (Ref. 72-50-00)
(2) Inspect exhaust duct for cracks or distortion. A maximum of three cracks in exhaust port
flange, emanating from bolt holes, are acceptable provided cracks are not across the flange-
to-case butt weld or circumferentially around flange or weld (Ref. Fig. 630). Cracks may be

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stop drilled using 1/16 (0.0625) inch diameter drill.

Figure 630 Exhaust Duct - Crack Inspection (Typical)

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1. Exhaust Flange
2. Butt Weld
3. Typical Crack

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NOTE: No other repair or replacement of exhaust duct is authorized at maintenance level. If

repair or replacement is required, the complete power turbine and exhaust section
must be shipped to an overhaul facility.
O. Fuel Nozzles and Sheaths
NOTE: See removal/installation procedures (Ref. 73-10-05).
(1) Inspect nozzles for dissimilarity of carbon build up. Clean or replace as necessary and
perform functional check as detailed in Chapter 73-10-05.
(2) Inspect sheaths for fretting wear, erosion and carbon build
up.
NOTE: Fretting wear up to 0.010 in. deep is acceptable.
P. Fuel Manifold Bolt Removal
(1) Inspect manifold attachments for broken
bolts.
(2) Remove and replace broken bolts (Ref. 72-30-
04).
Q. Reassembly of Engine
(1) Install combustion chamber liner (Ref. 72-40-
01).
(2) Install igniter plugs (Ref. 74-20-
00).
(3) Install fuel nozzles (Ref. 73-10-
05).
(4) Install compressor turbine disk (Ref. 72-50-
02).
(5) Install power turbine and exhaust section (Ref. 72-50-
00).
(6) Install power section (Ref. 72-00-
01).

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