•••

A SERVICE PUBLICATION OFLOCKHEED AERONAUTICALSYSTEMS COMPANY

EditorCharles I. Gale

Art DirectorCathy E. Miannay

Vol 21, No.3, July-September 1994

CONTENTS

2

3

11

16

Focal PointBen MethvinRegional DirectorInternational Sales

Conserving Turbine LifeThis update offers practical adviceon extending the service life of501 /T56 engine turbines.

C-130 Electrical System UpgradeA new mod package bringsMIL-STD-704E quality AC power tothe C-130’s avionics buses.

In Search of Fin StallIt’s not what you think: Under-standing C-l 30 rudder behaviorduring sideslip maneuvers.

Cover: A new C-130H recently deliveredto the 914th Airlift Group of the NewYork Air Force Reserve overflies thespectacular cataract of Niagara Falls.

Back Panel: Celebrating 40 years in con-tinuous production, the world’s mostversatile airlifter looks forward to a high-tech future as the C-l 30J takes shape.

Cover and photographs on pages 3, 16,18. and 19 by John Rossino.

by Lockheed SystemsCompany, a of Lockheed

by Lockheed accurate and should assumed, however, that

has approval from any government agency or unless noted for

and purposes only, and not construed for changes on or or as

any or procedures The marks are and owned

Corporation: “Hercu-les.” and Written must be obtained fromLockheed Systems Company before any

Address all News. Department Lockheed Systems

Company, GA Telephone Fax 40445’41017 Lockheed

Ben Methvin

Forty to the Future!

There has never been a program like that of theHercules. Forty years of continuous production isunprecedented. The wonderful Gooney Bird was inproduction for less than 11 years-although 10,629C-47/DC-3s were constructed. The C-133 was inproduction 9 years (50 were built), the Belfast 4 years(IO), the C-160 Transall 13 years (213), and theAntonov 12 for 16 years (145.0).

There are good reasons for the success of theHercules. Foremost, it is a sound design which hasbeen continuously improved. It fits a broad oper-ational spectrum. Like the basic design of the pickuptruck, the Hercules fits a world need for just this sizeair vehicle. It carries varied loads just the right dis-tance and puts them in just the right places. It is a

safe machine with no bad habits. Aircrews around the world test these capabilitiesdaily, and it consistently brings them home safe and sound.

The closed loop of continuous improvement and resulting longevity ofproduction seem to imply that the Hercules program will continue to flourish for along time. Each succeeding model was changed. The “A” (231 built) was a “hot-rod” with its short takeoff and quick response, the “B” (230) had longer range, andthe “E” (488) had both payload and range increases. The “H” (950+) took theexperience of the preceding models and was given more power and many structuraland maintenance improvements. Range and payload had been found to be nearoptimum and were not significantly changed.

Changes in a successful product are not made lightly. The C-13OJ. nowentering production, was recently referred to as “the son of Hercules” by oneprestigious English magazine. The “J” takes the best from all four ancestors, plusthose state-of-the-art improvements which make it significantly more cost-effectivein doing what it does best.

Would-be competitors have designs to replace the Hercules. It’s been triedbefore. The planned FLA will carry “more and bigger,” as will the Antonov 77T. But“more and bigger” is not consistent with the niche presently filled by the Hercules.Furthermore, neither will be able to achieve the price-to-value ratio of the Hercules.

I have been associated with the Hercules since the first flight of the prototype,and take great pride in working with the people who build it, support it, fly it,maintain it, and utilize its capabilities. I no longer even try to predict how long thismarvelous airplane will be in production, but I do know that only a Hercules canreplace a Hercules.

Ben Methvin’Regional DirectorInternational Sales

J. L. GAFFNEY - DIRECTOR

FIELD SUPPLY TECHNICAL RM&S CUSTOMERSUPPORT SUPPORT PUBLICATIONS DESIGN TRAINING

C. A. McLeish J. L. Bailey G. M. Lowe H. D. Hall S. S. Clark

T he single most important cause of degradedp e r f o r m a n c eand premature failure in gas turbine

engines is exposure to excessive heat. It makes littledifference from the standpoint of the effect on the enginewhat the cause of the overtemperature condition mightbe. What is important is only the degree of theoverheating and how long it continues. Each and everyexposure to excessive operating temperatures takes itstoll on engine life, decreasing operational efficiency andincreasing maintenance costs.

The important role that is played by the practices ofindividual Hercules aircraft operators and their main-tenance organizations in determining engine service lifeis not always fully appreciated. The fuel system of theAllison 501/T56 engine that powers the Herculesaircraft is provided with effective, automatic controlsthat keep engine temperatures within safe limits over awide variety of operating conditions.

As good as this system is, however, it was notdesigned to work alone. It cannot take the place ofcareful and conscientious management of the overalloperational environment. This is something only thehuman side of the equation can provide.

Life Cycle Factors

The normal maximum time between overhauls fora particular model of the Allison 50l/T56 engine isestablished on the basis of engineering considerations,user experience, and a careful evaluation of any safetyfactors involved. Within that time period, however, anengine’s individual maintenance requirements arelargely determined by the effect of the operatingenvironment upon major engine components.

Lockheed SERVICE NEWS V21 N3 3

For example, in areas subject to blowing sand orheavy burdens of dust in the atmosphere, thecompressor section may be seriously eroded long beforeother parts of the engine require attention. Moretypically, the repair intervals will be determined by thecondition of the turbine, combustion liners, and inletguide vanes of the “hot” section.

The turbine section components are made of metalalloys designed to be as durable as possible in a high-temperature operating environment. The service life isnevertheless rather dramatically affected by thetemperatures to which they are subjected. Metals do not“forget” incidents of exposure to overtemperatureconditions. Repeated exposure to high temperatures willeventually result in changes in their physicalcharacteristics that can lead to material failure. Let uslook at some of the effects of high-temperature engineoperation on turbine vanes and blades in more detail.

Sulfidation

Turbine blades and vanes are subject to gradualdeterioration as engine operating hours accumulate. Theprocess is commonly referred to as sulfidation. Turbinesulfidation is caused by the accelerated oxidation ofmetals in the presence of sulfur ions, in particularsulfides and sulfates. The oxidized metal goes out thetailpipe, leaving eroded blade and vane surfaces behind.

To help prevent such damage to the metal surfacesand extend the service life of the turbine sectioncomponents most often affected by sulfidation, diffusedcoatings of aluminum or aluminum and chromium (oftenreferred to as Alpak and AEP, respectively) are appliedto new turbine blade and vane surfaces.

First stage turbine blades after a service test.Top row: uncoated alloy.Bottom row: the same alloy protected with Alpak.

These coatings are effective while in place, but theyare eventually eroded away by the hot gases andabrasive particles passing through the turbine section.This means that sooner or later the base metal of turbinecomponents will become exposed to the damagingeffects of sulfur ions.

Controlling Sulfidation

The sulfur ions responsible for sulfidation comefrom a variety of sources and are not entirely avoidable.They may come from the sulfates in sea water, sulfur injet fuel, the sodium sulfate in aircraft cleaning solutions,hydrogen sulfide and sulfur dioxide in the air, or sulfur-bearing particulates injected into the atmosphere byindustrial processes.

While sulfidation is by far the most common andmost pervasive of the destructive processes affectingturbine components, it is also in some ways the mostcontrollable. The key to reducing the damaging effectsof sulfidation lies in understanding the factors involvedin the process. The rate at which sulfidation occurs inthe engines is dependent on three factors:

l Hours of operation.

l Concentration of sulfur ions.

l Temperature of the metal components.

There are limits to what can be done about some ofthese influences. It is obviously impractical to reduceaircraft operating hours just to avoid the sulfidationproblem. It is also impossible to avoid encountering thesulfur ions in the atmosphere totally, although it is wiseto avoid areas of heavy industrial pollution andunnecessary low-altitude exposure to sea air.

For all practical purposes, the first two of thesefactors can be regarded as constants, more or less

beyond the operator’s control. The situation is quitedifferent, however, in the case of the third factor.

How rapidly protective coatings like Alpak and AEPare worn away, and how rapidly sulfidation proceedsthereafter, is strongly affected by the operatingtemperature. Sulfidation is fundamentally a chemicalreaction and, like most chemical reactions, it proceedsmore rapidly at higher temperatures. It is no coincidencethat this destructive process is often called simply “hotcorrosion. ”

800 900 1000 1100TURBINE INLET TEMPERATURE (TIT)-C

Figure 1. Effect of temperature on sulfidation.

Sulfidation rates tend to be exponentially related tothe temperature of the metal, as shown in the chart inFigure 1. It is important to remember, however, thatturbine temperature is not a constant. The temperatureat which the turbine of the 501/T56 engine operates is avariable which is largely within the control of theoperator.

Reducing Sulfidation Effects

There are three principal ways in which flight crewsand maintenance specialists can help in reducing thetemperatures that turbine section components areexposed to, and thereby retard the sulfidation process.

l Fly at reduced cruise TIT whenever possible.

l Pay careful attention to starting temperatures.

l Ensure that the TIT sensing and indicating sys-tem is properly ‘maintained.

Lockheed SERVICE NEWS V21 N3

CRUISE DISTANCE-NAUTICAL MILES

Figure 2. Cruise time versus distance.

Cruise Conservation

Consider a mission where all parameters are fixedexcept for cruise TIT. A reduction of only 30 or 40degrees metal temperature will reduce the sulfidationfactor by almost one-half. Even further increases inturbine life may be realized with further reductions incruise TIT. When operation of the aircraft can beaccomplished with a lower TIT, the life of the turbinesection is greatly increased.

Furthermore, the sacrifice in terms of airspeed thatwill result from a modest reduction in TIT is really quitesmall. Figures 2 and 3 provide a comparison of powersettings and their effect on cruise time or fuelconsumption over a given distance. These charts helppoint up the relatively minor time advantage that can beexpected from exposing the turbine sections of anaircraft’s engines to high cruise TIT.

For 501-D22A, T56-A-15, T56-A-16, and T56-A-423 engines, decreasing power from 1010°C to 971 ' Cincreases block time only 3.8 minutes per hour of blocktime, while a further reduction to 932°C increases blocktime by just 6 minutes per hour. Moreover, it is possibleto achieve fuel savings of approximately 3 percent witha TIT reduction from 1010°C to 971 “C. Approximately5 percent total fuel savings may be realized by reducingengine power from 1010°C to 932°C.

30

20l STANDARD D

. 24,000 FT. ALT.

10

00 1000 2000 3000 4000 500

CRUISE DISTANCE-NAUTICAL MILES

Figure 3. Cruise fuel versus distance.

This points up two of the reasons why Lockheedrecommends that climbout TIT for this series of enginesbe set no higher than 971 ' C whenever possible, andcruise temperature be set no higher than 932°Cwhenever possible. Relatively little extra speed will begained from the use of high TIT settings, and thesehigher settings are also wasteful of fuel. But mostimportant, lower power settings will contribute greatlyto a longer turbine life.

The same conservation philosophy can also beapplied to 501-D22, T56-A-9, and T56-A-7 engineswith good effect, although in these cases the temperaturesettings will be lower. A number of operators ofHercules aircraft equipped with these engines havefound that keeping continuous cruise TIT at 900°C orbelow yields excellent results.

The nature of aircraft operations requires thataircraft engines be subjected to high power outputs andhigh temperatures on many occasions during their timein service. However, experienced operators havediscovered that conservative engine temperature settingsin cruise and at other times when safety and operational

Lockheed SERVICE NEWS V21 N3 5

considerations allow it will yield worthwhile benefits interms ofturbine section reliability and reduced operatingcosts.

Stress Rupture

Another heat-related factor which must be taken intoconsideration in realizing maximum engine service lifeis stress rupture. Stress rupture is breakage of turbineblades as a result ofphysical stress. The characteristic istime-dependent and a result of the stress applied bycentrifugal force (turbine rpm), sonic resonance, andtemperature in the turbine’s operating environment.

Of these, only the temperature (TIT) can be readilycontrolled. Turbine blades are more flexible when theyare hot, and therefore they are more subject to theeffects of centrifugal force and sonic vibration whenthey are exposed to excessively high temperatures.Keeping turbine blades well within their normaloperating temperature range is the key to reducing thepossibility of stress rupture damage.

The relationship between temperature and turbineblade “stress rupture life,” a measure of resistance tobreakage of this type, is shown in Figure 4. Note how a30-degree C increase above the normal operatingtemperature will decrease the stress rupture life by two-thirds. On the other hand, 30-degree decrease in blademetal temperature can double stress rupture life.

Relatively few instances of actual turbine bladebreakage due to stress rupture have been recorded forthe Allison 501/T56 engine. In the majority of thesecases, however, inspection of the affected parts hasrevealed previous exposures to overtemperatureconditions.

A Good Start on Conservation

Starting the engines is always a minor moment oftruth for an aircraft mission; the start can also besomething of a moment of truth as far as the engineturbines are concerned. To obtain good starts withconsistency, be sure to comply with the checklistprocedures closely. Pay particular attention to thefollowing:

l The throttle must be in the GROUND IDLE(military aircraft) or GROUND START (com-mercial aircraft) position.

l The propeller should be at minimum torque bladeangle.

l Aircraft boost pumps should be ON for the start toensure a good fuel supply to the engine fuelsystems.

Be sure that the residual TIT in the engine is lessthan 200°C before initiating a start.

The GTC or APU should be operating properly andat maximum output. Always check for a minimumbleed air manifold pressure of at least 35 psi beforethe start is initiated, and a minimum of 22 psiduring the start itself.

Air conditioning must be off during engine start,and in those aircraft so equipped, the ATM shouldbe off unless absolutely needed. In cases where theATM must be on for some reason, be sure tomonitor the bleed air pressure very closely whilestarting the first engine.

On manually controlled starter systems, be sure thatthe engine reaches 60% rpm before releasing thestarter switch or button; then release it promptly toensure maximum starter life.

Observe engine acceleration rate and the start timespecified in the applicable technical manual andoperator handbook.

6 Lockheed SERVICE NEWS V21 N3

Do not use enrichment on normal ground startsunless the engine will not start without it; however,all in-flight starts should be accomplished usingenrichment.

Closely monitor engine instruments throughout thestart cycle. TIT is the most important parameterduring starting. Watch this gage continuously whilethe start is in progress.

Check for the first rpm indication, which indicatesthat the start cycle has begun. Lightoff ordinarilyoccurs between 16% and 24% rpm, andacceleration should be continuous until the engine ison speed.

If enrichment is used, check for fuel flow andimmediate cutback after 16% rpm. Disregard it forthe rest of the start.

The secondary fuel pump pressure light should besteadily illuminated by the time the engine reaches65 % rpm. It should go out above 65 % rpm.

While the start cycle usually occurs withoutdifficulty, each start must be closely monitored so thatit may be discontinued in case an unexpected problemarises. The start should be aborted for any of thefollowing reasons:

l

l

TIT exceeds temperature limits cited in theapplicable manuals.

Bleed air manifold pressure falls below theminimum acceptable value of 22 psi.

Hesitating or stagnating rpm.

Fuel is observed pouring from the nacelle drain.

Torching-visible burning of fuel in and aft of thetailpipe-is observed. Note that a momentary burstof flame, an “enrichment burst,” is normal ifenrichment is selected.

Excessive smoke is seen from the exhaust.

Compressor surging or stalling occurs.

Abnormal vibration is noted.

Permissible start time is exceeded.

The engine does not light off (no increase in TIT)by 35 % or maximum engine rpm attainable with thestarter, whichever occurs first.

Problem Starts

The most desirable engine start is of relatively shortduration (less than 60 seconds), with peak TIT between780°C and 810°C. Note that starting temperaturescooler than 780°C are not necessarily better ones.

Temperatures on the cool side may result inacceleration rates inadequate to complete the start withinthe prescribed time limit. Low starting temperaturesmay be evidence of a lean TD null orifice valve settingor a lean fuel control acceleration schedule, or improperfuel nozzle flow patterns.

Improper fuel nozzle flow patterns can result in suchproblems as downstream burning of fuel, excessiveblade/vane heat soak temperatures, and a high incidenceof aborted starts.

On the other hand, starts with turbine inlettemperatures on the high side apply a great deal ofthermal stress to the turbine section of the engine. Therapid expansion forced upon turbine vanes, blades, andother critical components by the sudden application ofextreme heat can cause the parts to crack.

Such “hot starts” can do a lot of damage in a shortperiod of time and must be avoided. Remember that theTD system can only reduce high starting temperaturesafter they have been detected; it cannot prevent theinitial high readings from occurring. High temperaturesat start can often be corrected by making null orificeadjustments. If this does not solve the problem, furthertroubleshooting will be required. The following are thestart overtemperature limitations and the required actionif exceeded.

Over 830 Degrees C

TIT over 83O'C, but less than 85O'C, andexcluding the peak normally occurring at 94% rpm,requires the flight crew to record the extent of theovertemperature condition and call for maintenance atthe next layover where maintenance is available.

The action by maintenance personnel will be tocheck the TD valve null setting and adjust it towarddecrease. If it is found on the subsequent engine startthat the previous adjustment did not reduce the startingtemperature to within limits, it will be necessary formaintenance to perform the starting overtemperatnrecheck described in the applicable maintenance manual.

Over 850 Degrees C

If the TIT exceeds 85O'C, excluding the peakoccurring at 94 % rpm, the flight crew should proceed asfollows:

Lockheed SERVICE NEWS V21 N3 7

1.

2.

3.

Shut the engine down by moving the condition leverto the GROUND STOP position (not FEATHER),and record the maximum TIT.

Allow the engine to cool to below 200°C TITbefore attempting to restart. The engine may bemotored with the starter to cool it more rapidly, ifnecessary. When the engine has been cooled in thismanner, wait one minute to make sure that thetemperature has stabilized before initiating anotherstart. Remember not to exceed the starter dutycycle.

If 850°C is exceeded on the second start, shut downthe engine and record the temperature, then callmaintenance. Making another attempt to start theengine is not recommended.

Maintenance action will be to perform the startingovertemperature checks listed in the appropriatemaintenance manual.

Over 965 Degrees C

If TIT exceeds 965 ' C during a start, shut down theengine, record the peak temperature and call formaintenance. Make no attempt to restart the engine.Maintenance action will be to perform an over-temperature inspection, consisting of a borescope visualinspection of the turbine section components.

If no damage is found in the course of theseinspection procedures, all thermocouples must beremoved, visually inspected, and electrically tested. Ifthis inspection reveals no damage, maintenance canperform the starting overtemperature checks listed in theappropriate maintenance manual as required to correctthe condition.

Thermocouples

The proper operation of the TIT indicating systemhas a critical bearing on the overall operation andservice life of the 501/T56 engine. Malfunctioning ordamaged thermocouples, or installed thermocouples ofthe wrong type, can lead to increased operatingtemperatures within the turbine and materially reduceengine life.

It is important not to become complacent about theaccuracy of the TIT indicating system. It is possible toexceed the maximum allowable takeoff TIT limitationwithout knowing it if the engine has several deterioratedthermocouples. Under such circumstances, the operatorwould experience significantly reduced engine servicelife while mistakenly believing that the engine has beenoperating properly or has even begun developing morepower at the same throttle settings.

8 Lockheed SERVICE NEWS V21 N3

This turbine wane burn-through was caused by the use ofthermocouples of the incorrect type.

The engine TIT is measured by 18 thermocouplesinstalled in the turbine inlet casing. Each thermocouplehas two separate junctions that act as sensing elements.One junction of each of the 18 thermocouples isconnected in parallel to provide an averaged signal tothe TD amplifier.

The TD amplifier uses this signal to determine theactual TIT of the engine. The other junction of eachthermocouple is also connected in parallel. Its signalbecomes part of the averaged signal representing TITthat is displayed on the TIT indicator mounted on theengine instrument panel. Although unusual, it is possiblefor only one of the junctions to fail, causing incorrecttemperature sensing if the side connected to the TDamplifier fails, or a low instrument panel reading if theindicator side is affected.

The Allison 501/T56 engine incorporates a systemof multiple combustion chambers, which causes thetemperatures measured at the turbine inlet to besomewhat nonuniform. A system of multiplethermocouples is used to compensate for thesedifferences in temperature by sampling both the hotterand cooler areas and supplying an average temperaturesignal.

Since the temperature signal represents a valueobtained by averaging the input from 18 differentlocations around the turbine inlet, it is very importantthat all of the thermocouples be functioning properly atall times. If any of the thermocouples fail, the signalbeing sent to the TD amplifier and the TIT indicator willbe affected.

Thermocouples fail for a variety of reasons, in-cluding high resistance, open circuits, and probe tipdamage. The failure of the thermocouple is an event thatcan initiate a whole series of further events, all of thembad.

The basic problem is too much heat. Where one ormore of the thermocouples around the turbine area aredamaged or inoperative, the temperature signals beingreceived by the TD amplifier and the TIT indicator willno longer represent the true TIT values. The reason hasto do with the design characteristics of the TIT indicatorand the TD amplifier.

These devices respond to variations in the minutevoltages generated by the thermocouples as the TIT risesand falls during changes in engine power. But thevoltages will reflect the actual temperatures involvedonly as long as normal operating conditions prevail inthe TD control system. The electrical resistance loads inthe system circuitry are designed to yield accurateresults with 18 fully functional thermocouples. The loss

of any of the thermocouples will be accompanied by areduction in current flow that affects the voltage valuesfrom which the temperature data are derived.

As a practical matter, this means that if the currentflow is reduced by the failure of one or morethermocouples in the circuit, the voltage “seen” by theTIT indicator and the TD amplifier will also be reduced.Since lower voltage will be interpreted by the enginetemperature control system as lower TIT, the controlsystem increases the fuel flow to compensate. Theadditional fuel burns to produce more heat, which setsthe stage for overtemperature problems.

An important consequence of this effect is that overtime, thermocouple problems tend to be progressive andcumulative. A single thermocouple failure can begin asnowballing chain of events in which the increased fuelflow-and resulting higher temperatures-cause asecond thermocouple to fail, and then another, and so onuntil the engine is severely overheating.

Eroded or defective thermocouples that are stillproducing current can also cause temperature controlerrors, although for a different reason. In fact, thetemperature error in these cases can be even moresevere than when a thermocouple has droppedcompletely off line. One thermocouple with the aft wallof the probe tip eroded, for example, can cause the trueoperating TIT to increase by as much as 7.5'C.

All of these thermocouples were damaged by exposure to overtemperature conditions.

Lockheed SERVICE NEWS V21 N3

A sequence of failures involving five thermocoupleswith probe tip aft wall erosion can therefore raise theactual operating temperature of the engine by over 35 ' Cat every power setting above crossover. Such exposureto excessive operating temperatures is very destructiveto engine components and will result in significantlyshortened engine life.

Thermocouple Maintenance

Since properly functioning thermocouples play sucha vital role in the engine control system, it is clear thata well-designed and conscientiously followedthermocouple maintenance program is of utmostimportance. All thermocouples should be regularlyremoved, inspected, and tested as described in T. 0. 1 C-130B-6 or SMP 515C, and the applicable Allisonmanual.

One of the most important things to remember aboutthermocouple inspections is to do it often enough.Finding more than three defective thermocouples in oneengine at inspection time is a good indication that theperiod between inspections needs to be shortened.

When thermocouples are removed for examination,be sure to tag them as they are removed and keep arecord of the positions where they were located. Thecondition of individual thermocouples often providesvaluable clues about underlying problems involvingother components. For example, thermocouples withbadly burned probe tips or unusual carbon depositssuggest malfunctioning fuel nozzles or defectivecombustion liners, as does repeated thermocouplefailures at the same location.

Inadequate maintenance of fuel nozzles canconsiderably shorten engine life. Allison recommendsthat all fuel nozzles be regularly removed, inspected,and functionally tested as part of an ongoing fuel nozzlemaintenance program. An inspection interval of 1200hours is useful as a starting point for operators initiatingsuch a program, but this should be modified asnecessary to satisfy prevailing conditions. Manyoperators using JP-8 fuel, for example, have reduced thefuel nozzle inspection interval to every 600 hours.

For operators not equipped to accomplish fuelnozzle inspection at the operating facility, the mostpractical approach will be to remove the nozzles andreplace them with a set known to be serviceable every600 - 1200 hours (depending on your conditions). Theremoved units can then be sent to a properly equippedrepair facility. Those that meet specifications afterinspection, testing, and any necessary repairs can bereturned to service in the next cycle.

Also, don’t overlook bad fuel as a possible source ofthermocouple problems. Fuel contamination by dirt or

microorganisms can lead to thermocouple trouble byclogging fuel nozzles and altering fuel spray patterns.

Preventing Overtemperature Damage

Beyond strict adherence to the thermocouplemaintenance and inspection procedures set forth in theauthorized manuals, the best insurance againstovertemperature damage brought on by malfunctioningthermocouples is vigilance on the part of the flight crew.Higher than normal fuel flow and torque abovecrossover for a given TIT are possible signs ofmalfunctioning thermocouples.

Another is higher than normal torque for a givenTIT. Since fuel flow and torque are closely interrelated,an increase in fuel flow to raise the average TIT signalwill also cause an engine’s torque output to increase.Careful attention to the fuel flow and torque indicatorscan pay substantial dividends in terms of timelydetection of overtemperature conditions.

We have already seen that the TIT indicator willoften not be a reliable guide to the actual TIT when ther-mocouple trouble is present. Usually the indicated TITwill be lower, sometimes much lower, than the truevalue. You can make use of this fact to perform a simpletest when problems in the thermocouple system aresuspected. With all engines running and abovecrossover, and bleed air turned off, position the throttlelevers so that torque and fuel flow for each are aboutequal. An engine with thermocouple damage will showa noticeably lower TIT than the other three.

Finally, an important thing to remember is thatturbine engines never get better with age. They can onlydeteriorate as the combined effects of heat, erosion,sulfidation, corrosion, and wear gradually reduce theirefficiency. A 501/T56 engine that seems to beimproving with age-producing more torque at the sameindicated TIT, for example-should immediately attractyour attention. Some things are too good to be true, andthis is one of them. It is very likely that there areproblems in the TIT indicating system in this engine andprompt action is needed to prevent overtemperaturedamage.

The author and Service News wish to express specialthanks Wayne Shiver for his valuable assistance with thepreparation of the thermocouple section of this article.

Dare1 Traylor can be reached at 404-431-6567 (voice)or 404-431-6556 (fax).

“Conserving Turbine Life” first appeared in ServiceNews, Vl4N1, and proved so popular that the issue isnow out of print. To help ensure that this importantinformation remains accessible to Hercules operators,we arepleased to present an updated version of the samearticle in this issue. - Ed.

10 Lockheed SERVICE NEWS V2lN3

ior Staff Engineer”

T he electrical power system of the C- 130 Herculesairlifter has served well over the years, but as with

other aircraft electrical systems of older design, it hassometimes been subject to AC power system dropouts,spikes, and generator switchover anomalies. In the past,the aircraft’s analog electrical equipment tolerated theseproblems reasonably well and serious power fluctuation-induced failures were uncommon.

standards of MILrequirement, Locelectrical system upgradthat has been accepted byin most of the older C-130 Hercules aircraft in theUSAF inventory.

However, most current and next-generation digitalavionics components are considerably more sensitive tospikes and dropouts than the older equipment. This canlead to system malfunctions which result in erroneousreadings and a variety of other problems. The advent ofmore advanced avionics equipment clearly mandates“cleaner” electrical power.

The main components of the ESU modificationpackage for existing C-130 aircraft are shown in Figure1. The new system offers the following performanceenhancements:

l 10 KVA avionics AC buses that meet the powerquality requirements of MIL-STD-704E.

l Uninterrnptable AC power for a period of up to70 milliseconds.

New Hercules airplanes coming off the assemblyline at Lockheed’s production facility in Marietta,Georgia, now contain updated electrical systemsdesigned to meet many of these requirements. C-130aircraft Lockheed serial number 53 10 and up built forthe U.S. military, and all Hercules airlifters serialnumber 5358 and up, are equipped with electricalsystems that feature bus switching units, solid-stateinverters, and new generator control units. Theseupdated systems offer improved AC power suitable foruse with most types of modern avionics equipment.

l Avionics AC buses with both a primary and asecondary source of power.

l New generator control units, solid-stateinverters, a BIT status panel, and a modifiedoverhead control panel.

This article describes Lockheed’s ESU approach, andprovides additional information on the new capabilitiesthese improvements provide.

Updating the Fleet Bus-Switching System Components

The factory-installed upgrades provide newproduction Hercules aircraft with cleaner power for theAC avionics buses, but the U.S. Air Force recognizedthe need for a solution to the AC electrical powerproblems in older C-130s as well. The result was acompetitive procurement for improvements to theelectrical systems of existing C-130s that would addressthese issues, and in particular meet the high quality

Lockheed’s solution to the problem of designing apractical and affordable C- 130 ESU modificationpackage for existing Hercules aircraft is somewhatdifferent from the approach used in new-productionaircraft, but it achieves many of the same results whileoffering MIL-STD-704E quality power. The design isbased on a proven frequency converter which ispresently used in a USAF Special Operations Forces

Lockheed SERVICE NEWS V21 N3 11

OECP

GLC Line ContactorBTC-

I _ _ _ _ _ _

EXISTING EQUIPMENT

MODIFIED EQUIPMENT

NEWEQUIPMENT BUS

MAIN AVIONICS AC SYSTEMSPROVIDE Hz Hz I

figure 1. Electrical system upgrade (ESU) block diagram - one phase shown.

program. That system employs a versatile bus switchingsystem (BSS) to provide clean power under a widevariety of operating conditions. The Lockheed ESUemploys a similar design philosophy and includes thefollowing subsystems in the BSS portion of its circuitry:

l Rectifier control units

l 10 KVA inverters

The primary boost converter is set to regulate theoutput at 400 VDC. The secondary boost converter isset to regulate at 390 VDC. Whenever the primarycircuit no longer provides the 400 VDC output required,the load is automatically picked up by the secondarycircuit. If a loss of both input power sources occurs, thecapacitive power storage built into the system canmaintain the specified DC power output for a minimumof 70 milliseconds.

Rectifier Control Units 10 KVA Inverters

Two rectifier control units (RCUs) provide the 400VDC input required by the associated 10 KVA inverters(Figure 2). Input to the two RCUs is aircraft 1 15/200VAC, 400 Hz, three-phase power. Each of the RCUshas two inputs, each from separate AC buses. Thedesign of the RCUs provides a nonswitching selection ofthe AC input source. The RCUs feature a dual-redundant design that uses separate circuitry to developthe required DC output. If the primary DC/DC boostconverter portion of an RCU fails to supply the 400VDC output, the unit switches to the secondary inputsource and the secondary boost converter DC output.

The RCUs provide 400 VDC input to the 10 KVAinverters. These solid-state 10 KVA inverters recon-struct the DC to clean, uninterruptable, regulated115/200 VAC, 400Hz power, which is then supplied tothe avionics AC buses (Figure 3). The power input tothe 10 KVA inverters-can drop as low as 220 VDC andthey will continue to operate and provide 115/200 VAC,400 Hz output power.

Should conditions arise in which one of the 10 KVAinverters can no longer provide MIL-STD-704E power,the power output of the inverter drops to zero volts. This

12 Lockheed SERVICE NEWSV2lN3

design feature ensures that avionics AC bus equipmentis not damaged due to voltage degradation.

When the 10 KVA inverter senses that the input hasdropped below 220 VDC, an orderly shutdown of theinverter is initiated and the circuits in the RCU are reset.When the DC voltage from the RCU to the affected 10KVA inverter increases to a minimum of 220 VDC, theinverter restarts and is again able to generate a stabilized115/200 VAC, 400 Hz output.

Fail-Safe Bypass Contactor

The Lockheed ESU designapproach incorporates animportant safety feature, the fail-safe bypass contactor(FBC). The BSS includes an FBC at the essentialavionics 10 KVA inverter output to help eliminatesingle-point failure and ensure safety of flight. The FBCalso provides automatic emergency bypass power to theessential avionics AC equipment if the essential BSSfails.

Additional Electrical System Enhancements

In addition to the BSS, the Lockheed ESUincorporates several other important new electricalsystem enhancements, including:

l New generator control units.

l Solid-state inverters.

Figure 2. Rectifier control unit (RCU).

l Modified overhead electrical control panels.

l BIT status panels.

Generator Control Units

The new generator controlunits (GCUs) are off-the-shelf units that have been modified to provide remotebuilt-in-test (BIT) status. Essentially the same modelGCU is now being installed on new Hercules aircraft.

The GCU upgrade offers improved control of allgenerator output functions, and works with eitherBendix or Leland (GE) generators. The new GCUsautomatically detect which generator is present andselect the appropriate internal voltage regulationfunctions. The GCUs also provide system AC busprotection by sensing when the electrical operatingparameters are out of limits or serious conditions suchas feeder faults are present.

For example, the GCUs automatically sense whenfrequency exceeds or falls below specified limits. If sucha situation arises, the affected line contactor is opened,the OUT light on the overhead electrical control panel(OECP) is illuminated, and the appropriate BIT code isdisplayed. If voltage is sensed to be outside theallowable limits, the system responds in the same way,except that the generator is also deenergized.

When frequency returns to within limits, the linecontactor automatically resets and the OUT light on the

PRIMARYDCTODCBOOSTCONVERTER

DCTODCBOOST

1

CAPACITORBANK

420JOULESENERGYSTORAGE

LockheedSERVICE NEWS V2lN3 13

OECP extinguishes. In cases where a voltage failure has Three positions are provided on the GCU switches,been detected, however, the crew must reset the an “on” position labeled to identify the associatedgenerator with the control switches on the OECP. Note generator (Figure 4), OFF/RESET, and TEST. Thethat if a feeder fault is present, the GCU will “latch” the “on” position enables generator operation. Selecting theaffected generator system out, and it cannot be reset OFF/RESET disables the generator and resets the GCU.with the engine running. More detailed information The TEST position enables the generator in order toabout the operation of the new GCUs will be found in allow monitoring of AC voltage and frequency butService News issues Vol. 19, No. 2 and Vol. 20, No.3. leaves the generator line contactor deenergized.

Solid-State Inverters BIT Status Panel

Two new solid-state inverters replace the C-130’stwo existing rotary inverters. One new inverter replacesthe AC instruments and engine fuel control inverter; theother replaces the copilot’s AC instrument inverter. Thesolid-state inverters use the same mounting and cablingas the existing equipment. The new inverters accept asinput the nominal 28 VDC off the DC buses and supplyMIL-STD-704E continuous 115/200 VAC output.

The OECP modification also includes adding a newBIT status panel. This panel provides unit failureannunciation for the ESU line replaceable units. TheBIT status panel provides rapid identification of a GCU,RCU, or 10 KVA inverter failure.

Other Kit Equipment

Overhead Electrical Control Panel

The modified OECP allows the crew to manage theelectrical power system during both normal andemergency operations. Minor wiring changes allow thenew GCU switches to be installed, providing completecontrol of the new GCUs. The panel overlay has beenreplaced to reflect the new controls, and other changeshave been included to afford Night-Vision ImagingSystem (NVIS)compatibility.

All electrical and mechanical equipment necessaryfor installation are included as part of the Lockheed ESUkits. The electrical components include cable assembliesand miscellaneous wiring bundles. The mechanicalcomponents include mounting angles, nameplates,brackets, shelves, and hardware for installing thevarious components. The ESU was designed tominimize impact on the overall aircraft as a whole, andas a result, the requirements for miscellaneous fittings,adapters, etc., are minimal.

Figure 3. 1 OKVA inverter.

14 Lockheed SERVICE NEWS V21 N3

Figure 4. Modified overhead electrical control panel (OECP) with APU - typical.

Equipment Changes

The ESU modification addresses three majorsystems: the BSS supplying uninterruptable power to thetwo avionics AC buses, the GCU upgrade, and the solid-state inverter upgrade.

Installing the BSS equipment requires no equipmentremovals. Presently, the aircraft essential and main DCbuses are each fed by two transformer-rectifier units,and these units are supplied power from the essentialand main AC buses. The addition of the BSS does notalter this bus arrangement. The existing bus-tie network,supplying the input power to the RCU, remainsunchanged.

The GCU upgrade requires the removal of theexisting voltage regulators and generator control panels,from the electrical control and supply racks, and thefrequency-sensitive relays from the lower main ACdistribution panel. The new GCUs that replace all ofthese units require half the space of the old equipmentand provide a weight savings of 52 pounds.

Replacement of the rotary inverters involvesremoving the 250 VA and 2500 VA inverters from theforward right-hand underdeck area. They are replacedby new 250 VA and 2500 VA solid-state inverters. Thenew solid-state inverters provide the same control andstatus as the old rotary inverters. In addition, the newsolid-state inverters use the existing mounting andconnectors.

A new color-coded legend is provided for the circuitbreaker panel. This permits ready identification of theequipment powered by the new, clean avionics ACbuses.

Summary

The new BSS with its RCUs and 10 KVA invertersoffers significantly improved C-130 avionics AC buspower. The BSS provides MIL-STD-704E quality powerwith +/- 1Hz frequency control and +/-3 VAC voltage

control, exceeding the specification requirements. TheBSS also ensures uninterruptable power for input ACpower losses of up to 70 milliseconds. The BSS requiresno AC bus tie network alterations to existing aircraft.The ESU includes new GCUs, solid-state inverters, andmodification of the OECP to incorporate all GCU andBSS functions.

The Lockheed ESU provides older C- 130 Herculesaircraft with a practical and affordable solution to therequirement for improved AC power to operate today’sdigital avionics equipment.

The author and Service News would like to extendspecial thanks to Steve Rice, chief systems engineer onthe ESU project, without whom this article would nothave been possible. Grateful acknowledgement is alsodue to Norm Stevens and Bob Parker for their helpfulcomments and suggestions during review.

Susan Berk can be reached at 909-395-2790 (voice), or909-395-6565 (fax).

Lockheed SERVICE NEWS V21 N3 15

by C. A. Mason, Engineer Specialist, andE. S. Barland, Engineer Specialist, Sr.

Flight Control Department

The term “fin stall,” as it applies to the Herculesaircraft, is often misunderstood. There is a small

group of C-130 aircrew members who are convincedthat almost any sideslip excursion brings on fin stall.Others believe that any fairly rapid yawing motionbeyond that expected of the typical student pilot cancause the laws of physics and aerodynamics to berepealed and control of flight to be lost-all under thebanner of fin stall.

The truth of the matter is that the Hercules aircraftis not susceptible to fin stall under these circumstances,and the term fin stall itself is inaccurate when applied tothe Hercules in such situations.

Flight Control System Design

So what is really going on? The large power effectscreated by the propeller slipstream that make theHercules such a great short-field tactical transport also

16

mandate a large vertical stabilizer (fin) for directionalstability, with a large rudder for low-speed control ofasymmetric power from the engines. These features alsohad to be considered when it came to engineering theflight control system.

At the time of the C-130’s design in the early 1950sa propeller-powered transport aircraft did not justify thecost and complexity of the fully powered, irreversibleflight controls that are used on many aircraft today. Amore conventional reversible system was selectedinstead. Each control surface was sized to provide thecontrol power required, and a hydraulic booster wasadded to assist the pilot in situations where the forcesneeded to move the controls were too heavy for manualpower alone.

One characteristic of this type of system is that thepilot enjoys excellent feedback from the flight controls.He is directly connected to the main control surfaces and

Lockheed SERVICE NEWS V21 N3

can feel each surface’s position and the force necessaryto hold it or move it. Another characteristic is that in theabsence of inputs from the pilot, the rudder will “float”to an angle where the aerodynamic loads on the rudderare zero.

Sideslip Basics

During straight flight, these same aerodynamicloads cause the rudder to float at the neutral position,and the pilot must apply pedal force to push the rudderout to a given deflection. All other things being equal(equal power on all engines, etc.), the airplane nose willmove in the same direction as the rudder when therudder pedal is depressed. A small sideslip angle willthen develop, which pulls the opposite wing up andresults in a turn.

If, instead, the pilot lowers the wing to bank in thedirection opposite the turn, a steady-heading sideslipwill result. The airflow now hits the vertical fin from theside opposite the rudder. This off-centerline flow acts toreduce the rudder force that the pilot had to overcomeinitially, and the difference in rudder pedal force canactually be felt by the pilot. Engineers call these forces“hinge moments” because they really are momentstrying to rotate the rudder surface about its hinge line.Pilots prefer just “forces” because that is what theyapply to the pedal.

If the rudder angle required to maintain a givensideslip angle is greater than the float angle for thatamount of sideslip, then pedal forces must be applied tokeep the rudder from returning toward neutral. Assideslip angles increase toward the higher angles, thereis a tendency for the rudder to float out more rapidly and

Figure 1. A graphical representation of the relationship between rudder angleand pedal force during sideslip manuevers in the Hercules aircraft.

MaximumRudder

BALANCEDWITH ZERO

PEDAL FORCE

SIDESLIP ANGLE (DEG)

SIDESLIP ANGLE (DEG)

RUDDER FORCE

less pedal force is required to-hold therudder deflection.

A point is finally reached where therudder angle required to hold the sideslipangle and the rudder float position areidentical. At this point the rudder stayswhere it is with no pilot force on thepedals. In effect, the rudder is balancedwith zero pedal force (Figure 1). Theterm “rudder lock” is sometimes appliedto this point, but this is misleadingbecause the rudder is not actually locked.It does require force on the oppositepedal to return to a normal flightattitude, but the rudder will readilyrespond to that restoring pedal force.

Rudder Force Reversal

Beyond this point, the pilot mustapply opposite pedal force to restrain therudder from floating further out. This isrudder force reversal, also called rudderoverbalance. It is not fin stall, becausethe airplane remains directionally stableand will return to straight flight if therudder is returned to neutral and thewings are leveled.

Getting into rudder force reversalrequires abnormally high sideslip angles,which are prohibited by the flightmanual. Rudder force reversal ispreceded by heavy, unmistakable buffeton the vertical tail (but not fin stall) thatusually begins between 16 and 24degrees of sideslip. A noticeablereduction in rudder pedal force starts atabout 17 degrees of sideslip.

Lockheed SERVICE NEWS V21 N3 17

Further increases in sideslip produce large yawingtransients and a reversal of pedal forces. If the pilot letsthe rudder float and does not reduce the sideslip, theaircraft will yaw out to a sideslip angle of 40 to 45degrees. The nose-up pitching tendency that occurssimultaneously can cause one wing to stall. The airplaneis now in very serious trouble! Both lateral anddirectional stability have been lost and recovery isproblematic at best.

Susceptibility to encountering rudder force reversalis greatest at low speed and high power (usually 75%power or greater) with flaps extended. It has never beenreported in a power-off condition. Power effects on thewing and fuselage of the C-130 reduce the directionalstability level sufficiently so that rudder force reversaloccurs before fin stall.

Recovery from a rudder force reversal conditionrequires returning the rudder to neutral and can beassisted by rolling wings-level, pushing the nose downto decrease angle of attack, and reducing power; butonly if altitude permits.

Understanding Fin Stall

So does fin stall really exist? Theoretically, yes.Any lifting surface will stall if given enough angle ofattack. The vertical fin is a lifting surface, and sideslipis just angle of attack viewed from overhead. Whathappens in true fin stall is complete separation of theairflow over the fin and rudder. Just like a wing stall,complete airflow separation over the fin removes itshorizontal “lift,” or side force. Directional stability isthen lost because the restoring tail moment, which istotally dependent on that side force, disappears.

However, getting fin stall to occur on a Herculesrequires sideslips far beyond the flight manual limits andmuch hard work on the part of the pilot, all in the wrongdirection, of course. Extensive wind-tunnel testing andmuch flight testing have shown that the Here’s verticaltail does not stall, even out to sideslips as highas 30 degrees.

But don�t go out and try to verify this! Fin stall isoutside the flight manual limits. You will get into rudderforce reversal well before getting that high a sideslipangle, and you risk loss of the airplane if improper flightcontrol inputs are made. Also, you have no real way ofmeasuring sidelip angle with the instruments providedon your instrument panel, and the indicated airspeedbecomes unreliable at large sideslip angles. Any linepilot who intentionally investigates this condition shouldseriously consider another line of employment.

Slip-Ball Savvy

It must be kept in mind that the turn and slipindicator, the “slip ball” on the instrument panel, is nota sideslip gauge. The slip ball, or slip-skid ball if youprefer, reacts only to one thing: lateral acceleration,whether that acceleration comes from aerodynamicforces, inertia forces, gravitational forces (due to bankangle), or any combination of these. In particular, theslip ball should not be considered a primary referenceduring engine-out flight.

When the Hercules is being flown correctly underengine-inoperative conditions, the ball may be slightlydisplaced from the center toward the good engines; butthe amount of displacement varies with airspeed, bankangle, altitude, air temperature, and aircraft grossweight. DO NOT use the ball as an attitude referenceduring asymmetric thrust situations.

“Fin-Stall” Scenarios

Now consider the following scenario: You areexiting the drop zone on a formation drop and hit thewake of the Hercules that has crossed in front of you.The airplane rolls and yaws pretty violently and you feelthe rudder try to go hard over all by itself. You push therudder back to center with strongpressure on the opposite pedal -and recover to

Was it fin stall? No, because you did not losedirectional control. Was it rudder force reversal?Perhaps, but only for the brief time you were passingthrough the wake. As soon as you cleared the wake, therudder would have returned to neutral on its own. Yes,it was scary, but should you have also pushed the noseover and reduced power? Not at that low altitude. Thesituation would have returned to normal once you exitedthe wake, given enough time, altitude, and speed. But itis better to know what was going on and reactaccordingly.

Here is another scenario: A training flight is beingconducted to give the student pilot experience flying theC-130 with an engine inoperative. Under the guidanceof an instructor pilot, the student is performing a seriesof approach and go-around maneuvers with the No. 1engine simulated inoperative. He is being careful tokeep his speed above the charted air minimum controlspeed.

Following the go-around, the flight crew begins anair traffic control-directed left turn at 1,700 feet aboveactual ground level. The student pilot relaxes the rightrudder in a misguided attempt to keep the mmcoordinated and the ball centered. The aircraft yawsrapidly to the left, pitches over, and begins a steepdecent. In just 25 seconds the aircraft hits the ground,before control can be regained and recovery from thedive completed.

Was this loss of control caused by fin stall or rudderforce reversal? Neither one! We have already seen thatthe fin does not stall at small yaw angles. In this engine-out scenario, the asymmetric power condition producesa left yaw which is being controlled by right rudder.Rudder force reversal does not occur when the ruddercontrol input is opposing the yaw.

Improper use of the rudder, combined withimproper bank angle, produced this loss of control.Maintaining air minimum control speed (Vmca) ensuresonly that the airplane can be controlled in steady flightat the prescribed conditions of maximum rudder (orpedal force), favorable bank angle, etc. It does notensure additional yaw control for maneuvering,recovering from further upsets, or relaxing favorablebank angle or rudder deflection.

With No. 1 engine inoperative, turns should bemade to the right to avoid the large increase in Vmcacaused by the adverse bank angle. Remember that you,not the air traffic controller, are flying the airplane. Thecontroller most likely does not know or comprehend theimportance of maintaining a favorable bank angle duringengine-out operation at low speeds. Make sure that youcommunicate your problem (simulated or real) to airtraffic control and tell them what you plan to do.

Lockheed SERVICE NEWS V.21 N3 19

Controllability Guidelines

To prevent departures from controlled flight,especially during engine-out situations, the followingguidelines are suggested:

Maintain airspeed control. Do not let the airspeeddrop below minimum control speed.

Use smooth and coordinated flight control inputs.Control inputs should be slow enough so that anyresulting roll rates or yaw rates, intentional orunintentional, can be stopped at any time.

Make asymmetric throttle changes at a rate no fasterthan can be controlled, or balanced, by opposingrudder and aileron.

Never turn into the inoperative engine(s),

While the probability of encountering rudder forcereversal in flight is very remote, it is important that youknow about this flight characteristic, know how to avoidit, and know how to recover from it. Even moreimportant is that you fully understand the aircraft’shandling qualities with an engine inoperative.

The low-speed limits specified by the air minimumcontrol speeds and stall speeds are intended to preventoperation at airspeeds where any sudden loss of enginepower or reduction in control capability would causeloss of control of the aircraft. The effects of bank angleon minimum control speed and on stall speed aresignificant and must be respected.

Charles Mason and Scott Barland can be reached at404-494-4881 (voice), or 404-494-30.55 (fax).

' l;iw. ~ .,

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