Reference cooled 1st GP Turbine Blade 1-100-362-06/08

Problem Statement

Air Technology Engines Inc. (ATE) has been communicating with Honeywell Engineering

since mid-2016 in reference to an observed deterioration/erosion of the cooled 1st GP

Blade Leading edge, outer, blade tip section. See attachment labeled GP Blade Tip

Erosion. This deterioration has been ATE observed in approximately 50% of the engines

received for a 2500-3000 hour inspection. The outer, leading edge, blade tip sections of

many/most blades are eroded through the parent metal exposing the internal blade

cooling core. The designed blade cooling is lost when the internal cooling core is

exposed to the gas path. All blades require replacement. The effect of this blade

internal cooling loss on other GP cooled turbine components is unknown at this time.

Background Information

The L13B model engines exhibited similar GP blade erosion in approximately 25% of the

engines returned for 2500 hour inspection and overhaul. ATE determined that by

controlling all combustor liner air cooling gaps, repairing/replacing the GP nozzle curl

and combustor liner inner wall detail, and practicing a 600 hour time between fuel

manifold flow testing, the noted GP blade erosion was reduced to under 10% of the

engine population serviced.

The L13B model engines use the 1-130-780-01 combustor liner design. The L-703 model

engines use the 1-130-780-03 combustor liner design.

The -03 combustor liner design incorporated a change in the juncture of the GP nozzle

curl and the inner wall of the combustor liner. This change altered the boundary cooling

air being supplied to the GP blade tip and cylinder sections. The liner inner wall was

extended 0.5 inches and shrouded the boundary cooling holes introduced in the GP

Nozzle curl detail (1-100-262-06). The change was implemented to reduce/eliminate

wear occurring between the GP nozzle curl and the combustor liner inner wall dimples

(detail 1-130-253-03).

Observations

The boundary air introduced in the inner wall of the 1-130-780-01 combustor liner

design supplies cooling air to the gas path metal surfaces and the GP Blade tip section

region. Since the GP blades are rotating components, they experience an average

temperature and profile based upon the net effects of the cooling air introduced

upstream of the turbine operating plane. The effectiveness of the boundary cooling air

can be seen by observing the air flow “wake” or “foot print” from the air delivered

through the 44 liner dimples and exiting over the GP Nozzle curl. See attachment

labeled L13B curl. The boundary air introduction appears excellent. The wear patterns

from the liner dimples are also evident.

T53-L-703 Background Information

The Certification of the T53-L703 model engine was performed with the 1-130-780-01

combustor liner design. More than 3000 hours of engine endurance were demonstrated

at turbine inlet temperatures above the L-703 specification with no evidence of turbine

blade erosion or deterioration. The GP Blade material at that time was INCO 713, not

the improved material (C101) currently approved.

Observations

The L-703 engine uses the 1-130-780-03 Combustor liner design. This design change was

introduced to reduce the fretting damage incurred on the GP Nozzle curl from the

contact surfaces of the dimples on the combustor liner inner wall (detail 1-130-253-03).

The change also lengthened the inner wall 0.5 inches for unknown reasons. It is the

opinion of John Moore, Ed Pease, and Jack Sweet (Combustor Design Expert), that this

inner wall extension adversely affected the introduction of the boundary cooling air

being introduced to cool the GP Turbine blade tip section. The effectiveness of the -03

boundary cooling air introduced by the modified combustor liner (1-130-780-03) can be

observed from attachment titled L-703 curl.

The boundary cooling air effectiveness is witnessed by the “wake” or “foot print” as

seen from the brownish stains on the GP Nozzle curl. There is a significant reduction in

“wake” or boundary air “foot print”. The fretting wear still exists between the liner

dimples (now 88 Vs 44) and the nozzle curl. As previously stated, it is the opinion of AT

et al, that the shrouding of the GP nozzle curl boundary cooling holes has dramatically

reduced the effectiveness of their air introduction and function.

Side Discussion Comment(s)

ATE has determined that the axial gap between the combustor curl/deflector (1-100-

500-0x) and the air diffuser and combustor housing outer flange interfaces, can

significantly influence the occurrence of “combustor rumble”. This “rumble”

phenomenon was experienced during the certification testing of the T53-L13, and

occurred in the idle speed range. Raising the engine idle speed resolved the rumble

mode concern. ATE has introduced a build requirement defining the axial gap

requirements to eliminate the occurrence of combustor rumble within the operating

envelop of the T53 engines. It is the opinion of ATE that combustor rumble contributes

to the degree of fretting occurring on the GP nozzle curl, and may in fact be the root

cause for the curl wear.

Summary and Recommendations

The background information and current observations presented herein support the

need to restore the design intent of the T53-L-703 certification combustor assembly bill

of material. The introduction of the post certification combustor liner design (1-130-

780-03) did not retain the design intent with respect to the boundary cooling air

introduced to cool the 1st GP Blade tip region. Furthermore, the -03 design change

introduced has demonstrated a significant reduction in durable service life of the 1st GP

Blade (1-100-362-06/08).

The design changes recommended by ATE, are supported by our extensive knowledge of

the T53’s operation and many years of design and field experience. See attachment

labeled “design changes recommended”.

Closing statements

The intent of this Summary is to communicate the most recent findings and

observations relevant to the subject being discussed. Air Technology Engines Inc. is

recognized as a leader in the T53 overhaul and repair business. We desire to continually

improve and expand our T53 technical knowledge and maintenance expertise through

shared communication and mutual understanding.

ATE is hereby requesting Honeywell approval to proceed with the design modifications

defined in the attached layout drawing provided above. Reference to the .38 inch

cutback of combustor liner inner forward wall and introduction of 44 holes/slots equally

spaced between existing dimples.

I look forward to your positive and timely reply.

John E. Moore

Air Technology Engines Inc.

Naples Florida

July 10th, 2017

Cc: Mike Turner, Ed Pease, Jack Sweet, Mike Jose, Paul Benz, Brenda Wise, Greg Carloni,

Doug Kult, Bob Dorsey, file. Honeywell service centers (NA), Robert Baito, Paul Elliott. Original Post May 5, 2018

v/r,

Clarke

Clarke Pleasants

Assistant Professor of Aerospace 

Middle Tennessee State University 

615-898-2054

clarke.pleasants@mtsu.edu

In each event the pilot stated that they did not stiff arm the collective. I am sure they probably did not move their arm in relation to their body. If their body was moving without moving their arm they were inducing biomechanical feedback into the system.

In one occasion the bounce was so severe that the pilots feet came off the pedals. This situation of biomechanical feedback is also mentioned in the NTSB report DCA16FA199 of the July 06, 2016 accident of the Bell 525 test aircraft as a major factor.

Possible Mechanical factors The connection where the input rod connects to the servo all the bolts to include those on the servo should turn freely by hand after keying the connecting hardware. ( in the book ) If this hardware is tightened too tight it will cause feedback in the control.

I have seen on many aircraft that the head of the Pilots collective stick additional switches / buttons have been added.

This adds to the weight at the end of a long arm, making it hard to keep the collective from dropping in flight. Making it hard or impossible to correct with friction, preset or Pilot applied.

One mechanical item that can effect this is not enough built in or ground adjustable friction. ( TM 55-1520-210-23-2 pages included ) & ( TM 55-1520-242-MTF page included )

This adjustment should be done with the use of a hydraulic power supply. UH-1 requirements are 6 GPM flow and 1,000 psi.

With the Pilots collective boot removed you will see at the base of the collective a boss with two set screws that can be tightened or loosened onto a large washer. The screws are used to adjust the preset friction and should be adjusted equally. If they are not they can cause binding in the shoe application.

This usually happens because the bottom one is hard to get to so people will only adjust the upper one.

I have included a photo of tools used to do the job. If you have the collective pulled up to about half travel it is relatively easy to make the adjustments. ( see enclosed photo ) This is what is used to set the basic friction of 8 to 10 lbs. up & down. Inspect the shoes and pads for serviceability.

Notice the CAUTION a force of less than 8 lbs. could result in damage to the aircraft “ collective bounce “ Tools

I have included a photo of the tools used to set the built in or ground adjustable friction.

A length of cord. The one pictured is parachute cord with loops on both ends long enough to go around the twist grip and attach to the scale hook end.

The scale pictured is a push/pull rod on one end and a hook on the other. JONARD TOOLS P/N GPP-15 This one from GRANGER about $65.00 You use the hook to the cord to pull up. Turn the scale around and push down with the rod end.

The other a 5/32 allen wrench with a reach of 16” is best. Both of the ones pictured are homemade.

Pete

Scenario:

You are doing a lift job with little margin. The Load calculation indicates the load is within 200 lbs of the aircraft’s maximum lifting capability for the altitude and temperature. As the lift is attempted, the pilot increases power to 50 psi on the torque meter, the load does not move, and the lift is aborted.

The flight manual supplement indicates that the maximum torque is 50 psi, which translates to 1100 horse power, and we should have been able to lift the load according to the performance charts. So why won’t the aircraft lift what the charts indicate it should be capable of lifting? The FMS states in paragraph 2-16 and 5-7 that the helicopter is torque limited by the transmission to 1100 HP, which corresponds to 50 psi on the torque gage at 6600 engine/324 rotor RPM. However in paragraph 7-13 there is a notation that states “the power output capability of the T53-L-703 engine can exceed the transmission structural limit (50 psi calibrated) under certain conditions.” The key word here is “calibrated.”

In the TCDS for our aircraft, R00005SE, under note 12, it states “torque pressure output by the torque sensing system varies with the individual engines. The calibration of this value is required on each engine and the value corresponding to take-off power, is stamped on the engine data plate.” This statement applies to both the L-13 and the L-703 engines.

In the installation documents of the 703 engine STC, there is reference to TM55-1520-210-23-1/2/3, which is the UH-1 maintenance manual. In book 2 Section I, paragraph 8-6 (d), there is a procedure for marking the maximum torque limit by using the information on the engine data plate and a conversion chart in that paragraph.

So 50 psi on the torque gauge does not necessarily equate to 1100 SHP. As stated above, the torque sensing system varies with every engine, and therefore every engine will produce more or less power at a specific torque setting, and thus the need for calibration of the torque gage

How do we know how much power our engine is making?

During the engine test cell runs, one of the test points is Torquemeter PSI at 1125 ft/lbs. That is the torque pressure number which is stamped on the engine data plate. Since the power rating is taken at 1125 ft/lbs, the conversion chart is used to mark the torque gage to indicate 1100 SHP, which will usually be slightly above or below 50 psi.

What do I do with this number?

As stated above, TM 55-1520-210-23 requires me to convert this number to an instrument marking on the torque pressure indicator which then becomes the maximum allowable torque, equal to 1100 SHP.

1 psi on the gauge is about 200 lbs of external load lifting capability.

If the gauge is not marked properly the aircraft may not be producing the power necessary to meet the performance charts and lift as much as it should, or conversely it may be producing over the maximum rated power.

If your engine produces more than 1100 shp, there is a possibility of causing structural damage to the transmission as well as the airframe.

Originally published 4 15 16

Have a great week!

Clarke

Clarke Pleasants

Assistant Professor of Aerospace 

Middle Tennessee State University 

615-898-2054

clarke.pleasants@mtsu.edu

Compressor washes should be a routine procedure for those who maintain and operate gas turbine engines.

Albeit some engineers/A&P mechanics might consider compressor washes to be just another mundane task we must do “because it is written into our operations procedures” - but let’s consider how compressor washes affect turbine engine performance and life cycles.

The T53 engine variant is a free power turbine engine that employs a five-stage axial and single-stage centrifugal compressor driven by a two-stage gas producer turbine. Dramatic engine performance degradation occurs with the ongoing accumulation of airborne contaminants within the compressor module thereby reducing the aerodynamic efficiency of the compressor blades that can result in deteriorating engine performance, unsatisfactory acceleration and higher than normal Exhaust Gas Temperature (EGT).

Contamination, especially in high salt operating environments, can also lead to abnormal corrosion of the engine components. In order to maintain engine performance and reduce the corrosive effects on the engine, the debris that builds up in the compressor needs to be removed.

We do this through routine compressor washes.

Tips for effective compressor washes

A typical compressor wash involves:

- chemical wash (pre-mix product solutions that are biodegradable and non toxic or spray solvents into an engine with low core temperature),

- pre-determined soaking period (as required) a thorough water rinse followed by;

- an engine ground/drying run activating bleed air functions to ensure relevant accessories are also dried

Operators may choose to establish a wash schedule, the frequency of wash events relating to the amount of contaminants being ingested into the engine subject to the operating environment, while abiding by recommended OEM procedures.

First and foremost, always refer to the OEM procedures to ensure compliance with the specific engine type. The OEM specifies which chemicals can be used to wash the compressor, often referencing between a military specification or readily available solutions with a set specific chemical parameter. It is recommended operators develop a compressor wash schedule that best meets the operating situation IAW Honeywell MM references;

-13B/17 engines; 330.2 Maint. Manual, section 72-00-00, page 702

-L703 engines; 290.2 Maint. Manual, section 72-00-00, page 702

-L13B engines: 350.2 Maint. Manual, section 72-00-00, page 702

Note: use injection hardware and equipment that is also approved by the applicable OEM or engineering group to ensure correct/ proper wash and not introducing FOD hazards. Removing the contaminants restores engine efficiency, resulting in better fuel economy. It may result in lower EGT, lower corrosion, and restored performance.

Compressor washing is the single most cost-effective maintenance procedure for turbine engines. Performance is affected by both gas turbine thermal efficiency and the mechanical health of the components. Since the compressor typically consumes approximately 60-70 percent of power generated, compressor health is critical to engine health management. Contamination in the compressor section leads to deteriorating thermal efficiency which causes reduced engine performance. Reduced engine power margins can induce safety of flight considerations especially as the working environments that contribute the most contaminants to these performance degradations are those environments we least want performance issues.

Not only is engine performance affected, but damage to the compressor blades caused by contamination can lead to engine failure. Performed correctly, compressor washes can dramatically improve engine performance, increase engine life cycles and lower operating costs through:

• First line of maintenance action in any engine performance recovery related issues;

• Reduce potential engine compressor stalling

• Lowering fuel consumption;

• Improving engine compressor efficiency;

• Reducing Exhaust Gas Temperatures (EGT);

• Increasing hot section rotating group longevity;

• Reduce corrosion effects on the engine; and

• Extend Time Between Overhauls (TBOs)

Colin L Kilmaster B.Avn.

Never pull the panels with the cells full as the fuel weight from the aft cells will swell them up and then you can’t put the floors down with out defueling.

I also suggest not trying to put a sealant between the floor panels and the structure. You not sealing any thing, all the tie-down rings are not sealed and any liquid passes into the cell area.

There are drain holes at the front & aft inner channel of that cell area they are about a half inch in diameter. I have seen panels damaged when trying to pull-up ones that were sealed down, they are made to be removed.

Back to the areas to be inspected. The areas that crack are covered inside and out by vertical structure as you will see in the photos, that is why you have to pull the floors.

Hope this helps and that you find no problems. ( keep looking ) More later

Pete Sacramento Fire

Peter Frinchaboy

Helicopter Maintenance Manager

Sacramento Metro Fire District

The T53 engine has gone through several variations of bearings from un-pinned to pinned and also vendor manufacturing issues and vendor changes. These changes have been addressed over the years by the issuance of several different service bulletins. The main changes have been with the #1/2/3/4 main line bearings by way of re-designs and vendor changes:

1. The #1 and #4 bearings are both Thrust Ball Bearings. They have gone through various changes but the final change hasbeen to eliminate the silica/bronze cage that was susceptible to metal generation and the presence of bronze colored particles in the oil and filters. This has been eliminated by the introduction of SB0179/0180 with the introduction of P/N 1-300-663-01 bearing.

2. The #2/3 bearings are both Radial Rollers Bearings. The original 1-300-176-xx bearings were un-pinned outer race bearings. With the upgrade of service bulletins these bearings were upgraded to the 1-300-584-xx bearing which had a pinned outer race. The 1-300-176-03 became the the 1-300-584-01 and the 1-300-176-04 became the 1-300-584-02. The pinned outer race was introduced at the same time the bearing housing for the bearing was modified with the introduction of the corresponding slot to prevent bearing race movement and subsequent metal generation. The 1-300-584-02 was later replaced by the 1-300-665-01 as the 1-300-584-02 had a tendency to premature failure as it was designed with one extra roller which weakened the roller cage causing metal generation. the 1-300-665-xx series remains the latest bearing as there has been a vendor change.

3. Another change was to the #21 (Power Shaft Roller Bearing) was the introduction of the 1-300-082-03 bearing, which was the original replacement to the 1-300-082-01/02. An alternate to the 1-300-082-03 is the 1-300-664-01 that was introduced when an alternate vendor was approved by Honeywell.

4. Another major bearing change was introduced into the Accessory Drive Carrier by way of the removing the separate support bearings and replacing them with the 1-300-658-01/02 and 1-300-672-02. This was under SB0093 and SB0110 in order to eliminate most of the areas of metal generation that could occur in the Accessory Drive Carrier. The spare drive gear was also eliminated at this time which was a problem for Bell 204 operators. This drive gear elimination caused them to come up with an alternate starter installation as this gear drove the cooling fan that was mounted to the spare pad.

 By Wayne Bond 2 April 2017

(Covers Part 27, Part 29, and Restricted Category rotorcraft accidents to U.S. registered aircraft. Does not include gyrocopters or experimental aircraft.)

Please access the information from this briefing directly anytime at the FAA’s publicly available, Rotorcraft Accident Dashboard.

Find “Dashboard Navigation” near the top of the default page and click either “Historical Briefing” or “Make Model Breakdown” for access to all 3 pages of visualized and interactive data.

Monthly Summary

July Totals:  17 accidents, 5 fatal accident, 8 fatalities

- July had the highest monthly accident total so far in FY24. The 17 accidents were also the highest monthly total since July 2023 (21 accidents).

- For 6 different FYs over the previous 10 years, July had the highest accident count of any month.

- The industry sectors of Aerial Application (6 accidents during the month) and Personal/Private (4 accidents during the month) together accounted for 59% of the accidents during the month.

- The 5 fatal accidents that occurred during July were 3 fatal accidents higher than any of the other single month in FY24. For further context, the months of Jan-Jun combined had 5 fatal accidents.

- Of the 5 fatal accidents during July, 2 of the 5 were U.S. registered but occurred outside of the U.S. (1 in Canada, 1 in Ireland).

FY24 Summary

FY24 Totals:  83 accidents, 14 fatal accidents, 27 fatalities

Accidents (includes both fatal & non-fatal accidents):

- The FY24 estimated accident rate for Oct-Jul was 3.44 per 100K hours (4% lower than the same period in FY23 and 13% lower than the 5 year average for the same period).

- The accident count of 83 for Oct-Jul remained the third lowest total for that time period for the 42 FYs on record, trailing only FY21 (80 accidents) and FY20 (81 accidents).

- The following 3 industry sectors accounted for 55% of the accidents for Oct-Jul: Personal/Private (25%), Aerial Application (17%), and Instructional/Training (13%).

Fatal Accidents:

- The FY24 estimated fatal accident rate for Oct-Jul was 0.58 per 100K hours (3% lower than the same period in FY23 and 27% lower than the 5 year average for the same period).

- The 5 fatal accidents in July increased the FY24 fatal accident count 56% from the FY24 Oct-Jun total.

- The Personal/Private industry sector accounted for 43% of the Oct-Jul fatal accidents, over 3 times higher than the percentage of the next highest industry sector.

- Through 10 months of FY24, the percentage of accidents with a fatality was 17%.

Fatalities:

- The FY24 estimated fatality rate for Oct-Jul was 1.12 per 100K hours (10% lower than the same period in FY23, 28% lower than the 5 year average for the same period).

- The 27 fatalities from Oct-Jul was the third lowest for that time period for the 42 FYs on record, trailing FY11 and FY17.

U.S. Helicopter Safety Team (USHST) Calendar Year Metrics

Includes only helicopter accidents that meet the following criteria:

1. U.S. registered aircraft

2. Operating in the U.S./U.S. territories (includes offshore)

- USHST Goal:  Reduce the 5 year average fatal accident rate to 0.55 per 100K hours by 2025.  The USHST uses the 5 year average fatal accident rate from CYs 2014-2018 (0.58 per 100K hours, revised from 0.62) as their baseline for measurement.

- The CY 2020-2024 5 year average fatal accident rate was 0.64 per 100K hours through Jul 2024.

Lee Roskop

Aviation Safety Coordinator (Rotorcraft)

Fleet Safety Section, AIR-723

Operational Safety Branch, Compliance & Airworthiness Division

817-222-5337

lee.roskop@faa.gov

No one expected the high gearbox pressure problem that was unwittingly introduced. The excess material that was added in the area of the lower power pinion bearing restricted the oil from getting out of the gearbox. The gears running in oil created foaming of the oil which resulted in the engine scavenge pump not being able to pump the oil and foam from the engine and the breather tube from the air/oil separator in the accessory gearbox would “spit oil and air) out of the gearbox vent. Oil scavenge temperatures would rise from approximately 250 degrees F as high as 300 degrees F. Local area in the vicinity of the air/oil separator gear would exceed 400 degrees F and the paint would discolor and blister. Instead of 0 gearbox pressure, the gearbox pressure would increase as high as 2 psi. The higher engine oil flows like 2900 #/hr would result in even higher gearbox pressure.

The higher gear box pressure causes many problems for the engine. The higher oil temperature causes oil to go acidic with fewer operating hours on the oil. Acidic oil reduces the pressure required to squeeze out the oil film between the teeth of the main reduction gears resulting in heavy gear wear. The intershaft oil seal is designed to release some air in the main reduction gear area raising the pressure in the RGB approximately one psi. It is that pressure that forces the oil from the RGB into the accessory gearbox. The excess material in the gearbox creates flooding in the main reduction (RGB) gear box and the gear losses resulting from the gears running in oil results in a 1% power loss in the engine (a loss of 12 to 14 horsepower). The heat imparted to the oil overloads the aircraft oil cooler and creates problems on hot days.

The high gearbox pressure restricts the oil scavenging from the main bearing packages and that results in the carbon seals running in oil and carbon buildup that restricts the seal elements from sealing properly. The bearings operate at a higher temperature which decreases bearing life. The seal elements in the main shaft seals run hotter, build up carbon and leak oil. The starter shaft seal becomes hard from the increased oil temperature and with the higher pressure leaks oil into the starter generator. The magnesium gearbox cover expands from the increased heat and the cover in line with the washers under the nuts yields. The yielding results in lower torque and oil leakage between the cover and gearbox housing.

Honeywell designed an improved starter shaft seal that requires gearbox pressure to seal. It drastically decreases the starter shaft oil seal leakage problem but does not fix any other engine problems resulting from the high gearbox pressure. The proper fix is to remove the excess material in the gearbox housing. Over the years, there has been at least four different manufacturers of the accessory gearbox housing. All of them have excess material that blocks the oil flow coming through the 6 o,clock strut in the inlet housing and into the gooseneck in the accessory gearbox. Each manufacturer has made the problem worse by taking liberties to beef up the gearbox making it stronger but actually decreasing the life of the engine.

ATE supplied Honeywell with ATE Engineering Orders (approved by the FAA) and Honeywell issued SB 0175 to repair housings having excess material.

There are two other reworks that are required to assure continued operation of the accessory gearbox. The upper flange of the gearbox housing that mates to the inlet housing contains the seat for the bearing retainer that holds the upper power pinion gear bearing. The groove in the housing just below the seat is for the 0’Ring seal. Without the O’Ring in place, the steel bearing retainer should be above the mating flange of the gearbox (to the inlet housing) a minimum of 0.0005 inch and a maximum of 0.0015 inch to assure proper clamping of the power pinion bearing retainer.

Apparently many gearboxes have had the upper flange resurfaced and nothing done to remove the interference when the gearbox is bolted to the inlet housing. The excess load can cause one of the two extensions for the two aft bolts to crack.

Another problem over time is corrosion of the upper seat for the power pinion bearing retainer. The corrosion does not occur equally and has caused cocking of the upper power pinion bearing. The depth and the flatness of the seat is critical. There is a repair that machines the seat flat and parallel to the upper flange. After machining the seat, the flange mating with the inlet housing has to be machined to maintain the bearing retainer seat (0,119 inch). The O’Ring groove has to be machined to it’s proper depth (0.070 to 0.080 inch) to assure that the clamping force of the bolts for holding the gearbox to the inlet housing seats the bearing retainer in it’s seat. An inadequate O’Ring groove depth could prevent the power pinion gear to be in it’s proper location.

Anyone wanting to see a partially reworked gearbox with one half of the gearbox reworked and one half unreworked can visit the Air Technology Engine Booth at the HAI Convention in Dallas.

The picture shows the accessory gearbox housing surface that mates with the inlet housing. The flange, O’Ring groove and the retainer groove for the power pinion bearing retainer. On the left is the reworks and the unreworked portion is on the right.

The 2nd picture shows the inside of the accessory gearbox looking forward. The polished area in the center of the gearbox is where the excessive material was removed. Opening the passageway eliminated the gearbox pressure, stopped the foaming and overheating of the oil, stopped the flooding of the main reduction gear cavity, lowered the main bearing temperatures, improved the main shaft and starter shaft seal lives and allows the oil cooler in the aircraft to operate within it’s design limits.

Have a great week!

Clarke

Clarke Pleasants

Assistant Professor of Aerospace 

Middle Tennessee State University 

clarke.pleasants@mtsu.edu

These nozzles have an extremely small orifice that can be plugged by small particles in the fuel or by (more common ) coking around the nozzle. The coking is caused by fuel burning on the face of the nozzle.

The T-53 is designed so that when the start fuel solenoid closes air from the burner housing forces air through the start nozzle forcing out the remaining fuel so it does not burn on the face of the nozzle.

This is not a foolproof system. I recommend inspecting start fuel nozzles every 300 hours and flow checking them or sending them out to your engine shop and having them flow check them. Some of the nozzles can be cleaned or repaired by the engine shop. I do not recommend trying to clean them yourself, the orifice can be damaged.

Start Fuel nozzles when installed are directed at the igniter plugs for good starting. If the only igniter and only start fuel nozzle are on opposite sides of the engine, a good cool start is not likely, slow or hot starts harm internal engine parts.

Peter Frinchaboy

Helicopter Maintenance Manager

Sacramento Metro Fire District 03/16/2017

Carbon ball blasting is caused by particles of fuel that have not been completely consumed. It usually the result of fuel particles being too large to be consumed. The size of the fuel particles after being atomized is referred as the Sauter Mean Diameter. As the fuel nozzle gathers operating hours, the fuel particle Sauter Mean Diameter can be larger in emitting from some fuel nozzles and not in others. Hot refueling often results in very fine sand particles getting into the fuel and the particles are too fine to be caught by filtration. Often the sand will become packed on the screen inside the fuel nozzle which affects the atomization of the fuel.. Often the sand causes erosion of the slots in the nozzle orifices and the end result is larger fuel particles. Larger fuel particles are not completely consumed and the particles are in the form of carbon. When the blade tip is above approximately 1650 degrees F, the rate of erosion from the carbon balls increases. When possible, an operator should consider sending the fuel manifold to a Service Center for flowing, cleaning and possible fuel nozzle replacement. Usually at approximately 600 hours, the cleaning is effective in restoring the fuel nozzles to their proper flows.

The inner wall section of the 1-130-780-01 combustor liner part number is 1-130-253-01. The 1-130-253-01 section of inner wall has dimples that are 0.030 inch in depth. They offset the inner wall 0.030 inch from the nozzle curl for cooling air that follows the wall of the turbine nozzle curl and cools the top 1/3 of the Gas Producer turbine blades. The nozzle curl where the liner dimples contact it, wears and as it wears, the cooling air is reduced. The nozzle curl can be welded and then ground to conform to the outer circumference but the 1-130-253-01 inner wall should be replaced. IT IS A MISTAKE TO BEND THE TABS OF THE Liner against the nozzle curl because the rate of carbon ball blasting increases and erodes the turbine blades. However if the tabs do not contact the curl, they may eventually break off. (From vibratory stress)

The 1-110-500-03 deflector is supposed to have a chamfer as a lead into the combustor liner outer flange. When installed in the engine, there should be a clearance between the outer flange on the combustor liner (where both parts contact in operation) and the 1-110-500-03 deflector to allow for thermal expansion of the liner. The clearance is approximately 0.050 to 0.070 inch. If the gap is excessive, the engine will vibrate and RUMBLE at idle. If there is little or no gap, the 1-110-500-03 deflector will crack around the inner flange where it is bolted to the air diffuser. When a deflector is repaired, caution has to be taken to control the dimension between the mounting flange of the deflector and the contact area where it meets with the combustor liner. Global now controls that dimension.

The advantage of an annular combustor is that a radial temperature profile from base to tip can be obtained quite easily and will benefit the life of the turbine blades. A more compact engine envelope is also obtained. The 44 chutes in the outer wall provide cooling air to the base of the 1st stage GP blades. In the T53-L-13, 13A and 13B, the base of the blade is actually 150 to 170 degrees cooler than the gas temperature of 1730 degrees F at Takeoff power. Two thirds of the way outward on the blades, the blade leading edge temperature is equal to the gas temperature of 1730 degrees F. IF the inner wall is offset by 0.030 inch dimples from the nozzle curl (by design), the blade leading edge temperature will drop a minimum of 75 degrees from the hottest point of 1730 degrees F (2/3 outward from the base) to the tip.

By having the hottest point of the blade 2/3 outward from the base predetermines that should a stress rupture occur, it will occur 2/3 out on the blade and not at the base. It is what I consider a "soft" failure because the engine should continue to run and make adequate power.

The stress rupture life of the blades decreases as the metal temperature increases. The stress rupture life (in the Design Report) was based on a radial temperature profile of 1580 degrees F at the base increasing to 1730 degrees F at the leading edge 2/3 of the blade length outward from the base and then decreasing to 1655 degrees F at the tip. The gas temperature at 1730 degrees F turbine Inlet Temperature drops across each set of turbines. A good performing engine will have approximately a 700 degree F temperature drop across both sets of turbines. Approximately 61% of the temperature drop is across the Gas Producer Turbines and 39% across the free power turbines. By set of turbines, the turbine nozzles are included.

It is not the same for all engines because the temperature drop varies as the nozzle areas vary. A good performing engine will have a temperature drop of approximately 420 degrees across the GP turbines. Approximately 2/3 of that occurs in the 1st stage GP Turbine. That means the gas temperature exiting from the 1st stage GP turbine has decreased from 1730 degrees F to approximately 1450 degrees F. That temperature gradient is approximately the same as the metal temperature gradient across the blade.

The T53-L-703 is subject to excessive carbon ball blasting. The reason is not clearly understood because the 1st GP turbine blades are cooled by compressor discharge air and should have a lower metal temperature than the blades in the T53-L-13B. A comparison of the total area allowed for cooling of the blade tips leaves doubt. Some of the holes for cooling air did not consider that the holes around the circumference are partially blocked when the 1-130-780-03 is inserted onto the nozzle curl. I believe that the T53-L-703 tip cooling area is approximately 60% of the area of the design requirement of the T53-L-13B.

Air Technology Engines, Inc. is considering the submittal of an Engineering Order to the FAA that would allow additional holes in the forward section of the 1-130-780-03 combustor liner. The intent is to increase the amount of cooling air to the tips of the turbine blades making the area in the T53-L-703 (T5317 models) the same as the T53-L-13B (T5313B).

The atomization of the fuel is dependent on fuel quality and cleanliness. Fine particles of sand, small enough to get through the aircraft fuel filter, get lodged against the primary and secondary screens in the fuel nozzles and affect the amount and quality of the fuel going through the 0.004 inch orifices into the swirl chambers. There are two different fuel sprays (primary and secondary) emitting at different angles and creating a common spray angle directly in front of the nose of each fuel nozzle. Proper atomization is dependent on the proper flows coming from the primary and secondary fuel orifices in the fuel nozzles into the swirl chambers. As the flow falls off due to contamination, the particles of fuel increase in diameter and become carbon balls as a result of incomplete burning. The quality of the fuel and proper refueling procedures can reduce the carbon ball erosion problem.

It is important know your engine and how it performs. Cooler running engines last longer. When you know that the engine requires a higher EGT or TGT to make power, it is prudent to find out what is wrong. Even though the engine makes adequate power, more than likely costly damage is being done to the hot section of the engine. Sometimes all the engine needs is a compressor cleaning.

My name Is Edward C Pease (Ed Pease). I began working at Lycoming in 1957 as a test engineer on the T53_L-1 engine. In 1965, I was in charge of the development of the T53-L-13. For a period of four years I managed the development programs of all of the turboprop and turboshaft T53 engine models. I took early retirement in 1987 and became a turbine engine consultant on many makes and models of turbine engines. After 60 years in turbine engines, I limit my consulting to two overhaul shops in Florida. My assistant (John E Moore) at Lycoming for many years, still works with me at Air Technology Engines, Inc. We do their engineering which includes the development of repairs.

Originally Posted 10/16/2016