N 94 - 2 36 4

NASA Technical Memorandum 83679A-84-40-66-7000

Contingency Power Concepts forHelicopter Turboshaft Engine

Robert Hirschkron, Richard H. Davis, David N. Goldstein,and John F. HaynesGeneral Electric Co.Lynn, Massachusetts

and

James W. GauntnerLewis Research CenterCleveland, Ohio

Prepared for theFortieth Annual Forum of the American Helicopter SocietyArlington, Virginia, May 16-18, 1984

NASA

https://ntrs.nasa.gov/search.jsp?R=19840015580 2019-07-08T00:37:51+00:00Z

CONTINGENCY POWER CONCEPTS FOR HELICOPTER TURBOSHAFT ENGINE

by

Robecc HirschkronManager. Preliminary Qesign

and 'Richard H. Davis. David N. GoldsCein and John F. Haynes

Design Engineers. Preliminary DesignGeneral Electric Co.Lynn. Massachusetts

James W. GauntnerProject Engineer

NASA Lewis Research CenterCleveland. Ohio

Abstract

Twin helicopter engines are oftensized by power requirement of safemission completion after the failure ofone of the two engines. This study wasundertaken for NASA Lewis by GeneralElectric Co. to evaluate the merits ofspecial design features to provide a 21/2 minute Contingency Power rating,permitting ah engine size reduction. Themerits of water injection, cooling flowmodulation, throttle push and anauxiliary power plant were evaluatedusing military life cycle cost (LCC) andcommercial helicopter direct operatingcost (DOC) merit factors in a "rubberengine/rubber aircraft" scenario.

CRP

DOC

GW

HPT

IRP

LCC

MC

NH

OEI

OWE

SLS

TOP

T41

Notation

Contingency rated power

Direct operating cost

Gross weight

High pressure turbine

Intermediate rated power

Life cycle cost

Maximum continuous

Compressor RPM

One engine inoperative

Operating weight empty

Sea level static

Takeoff power

Turbine inlet temperature

Presented at the 40th Annual Forum of theAmerican Helicopter Society. Arlington.Virginia. May 16-18. 1984

Introduction

Suppliers of military and civilhelicopter engines have been aware of theneed for higher contingency power ratingsfor future rotorcraft.. These suppliershave both civil and military enginescurrently qualified with CRP usingconventional approaches. For example,the GE T700-401 (LAMPS) engine isdesigned and qualified to meet a 2 1/2minute contingency power requirement. 10%greater than intermediate rated power.However, savings in both DOC and LCC canbe achieved by still higher contingencypower ratings.

An incentive to provide highercontingency power for a given set of OEIrequirements is downsizing the enginewhich allows reductions in rotorcraftsize reductions as well as savings inoperating costs. However, the designfeatures required to provide contingencypower, while maintaining constant life,involve penalties in powerplant SFC.weight and cost, which partially offsetthe downsizing benefit.

the results of a systematic study ofseveral contingency power conceptsperformed by General Electric Company'sAircraft Engine Business Group for NASALewis in 1983" are reported in thispaper. The scope of the study included adefinition of turboshaft engine designswith varying contingency power capabilityto establish a parametric approach to the"cost" of including this feature.The changes in the engine design wereestablished and included in rotorcraftdesign, performance and economicpenalties in terms of life cycle cost(LCC) and direct operating cost (DOC) forthe military and civil rotorcraft.respectively. Rotorcraft design supportwas supplied by Sikorsky Aircraft undersubcontract to General Electric Company.

« NAS3-23705

The basic approach taken was to designall engines foe equal life whether theyhave contingency power ratings or not.The design changes necessary toaccomplish this objective result inpenalties in propulsion system weight,cost and SFC. These were balancedagainst the advantages of using thesmaller engine to obtain a quantitativeevaluation of the net system gain.

Baseline Rotorcraft & Engines

Conventional helicopter designs werechosen for the baseline civil & militaryrotorcraft. These designs were judged torepresent a significant segment of therotorcraft market in the 90's. Thelevels of cruise design speeds arelogical for the two missions. In the lowspeed case, one engine inoperative (OEI)takeoff is clearly sizing the engineswhile in the high speed case cruise poweris close to becoming the sizing condition.

The civil rotorcraft is a 30passenger commuter with 250 statute miledesign range. -Cruise speed is 160 lets at3000 ft altitude and fuel reserve issufficient for 25 statute miles plus 30minutes at 160 kts. Power sizingcriteria for the engines is determined byhover in ground effect (HIGH) with OEI at1000 ft altitude. 82.4°F ambienttemperature (ISA + 15°C) equivalent toCat A design criteria. The rotorcraftdesign features are detailed inTable 1.

The military rotorcraft is a 24 troopmarine assault transport (HelicopterExperimental Marine. HXM) with 200nautical miles design radius; that is.200 nautical miles out with troops andgear plus 200 nautical miles returnwithout payload. Design speed is 160 ktsand the power sizing criteria for theengines is determined by hover in groundeffect with one engine inoperative at3000 ft altitude. 91.5°F (hot day) alsoequivalent to Cat A design criteria. Thedesign features are also listed inTable 1.

The technology level of the engineswas selected to be consistent withqualification in the late 80's. Thebaseline turboshaft engine for both thecivil and military vehicles consists ofa 20:1 pressure ratio classaxial-centrifugal compressor, a two stagecooled high pressure turbine, an annularcombustor, and a three stage uncooledfree ppwer turbine. The 100 F lowerturbine inlet temperature for the civilengine was chosen because of the need formore mission life than in the militaryengine. Contingency power of +15% over 'intermediate rated power (IRP) or takeoffPower (TOP) was available for thebaseline engine at the high ambienttemperature design point as beingrepresentative of what would be designedinto a new engine. Smaller powerincreases are available at cooler ambienttemperatures where corrected speeds

Table 1. Baseline rotorcraft Characteristics

CIVIL COMMUTER MILITARY TRANSPORT

Entry into Service

Type

Gross Weight, Ib

Cruise Speed, kts

Payload, Ib

Passengers

Rotor dia. ft.

No. of Engines

SLS IRP SHP Class

Fuel Weight. Ib

Empty Weight. Ib

1990 1990

Conventional Single Main Rotor Helicopter

26000 Class 35000 class

160 . 180

6000 " 5760

30 24 Troop & gear

66 66

2 2

3000 5000

2500 plus 5500 plus

17000 plus 21000 plus

rather Chan turbine temperature, may belimiting. A summary of the engine cycleis provided in Table 2.

fuel consumption increase resulting fromthe turbine cooling flow and turbineefficiency changes.

Table 2. Cycle for baseline engines

CIVIL MILITARY

Sizing Condition,Alt/Mo/To

SHP @ sizing Condition.TOP or IRP

CRP/TOP or CRP/IRP

Design Corrected Flow class,Ib/sec

Cycle Pressure Ratio

Turbine Rotor InletTemperature atTOP or IRP. °F

1000K/5/0/82.4 F

2690

1.15

15

20:1 Class

2400°F

3000/5/0/91.5 F

3940

1.15

22

20:1 Class

2500°F '

Engine Sizing Procedure

The direct approach to increasing thecontingency power ratio beyond 1.15 is by"throttle push" to greater physical rotorspeeds and turbine inlet temperatures.Increases in rotor structural weight andturbine cooling flow are required tomaintain engine life at the baselineengine level and these will have adetrimental impact on engine weight,power and fuel consumption. Beyondcontingency power ratios of approximately1.30, the compressor must be rematched tolower speeds at IRP or TOP in order toprovide sufficient flow capacity, stallmargin and compressor efficiency at CRP.This results in a larger, heavier engine(to maintain IRP) and the attendantreduced cycle pressure ratio at normalpower settings is reflected in higherfuel consumption.

Fig. 1 shows how two differentengines with CP.P/TOP ratios of 1.15(baseline), and 1.35 are matched in termsof compressor corrected airflow versuscorrected speed for the civilapplication. This curve reflects atypical compressor characteristic andshows how increasing the CRP/TOP ratioreduces takeoff power to a lowercorrected airflow.

Table 3 lists the increase incompressor flow size required to maintainconstant takeoff power and the specific

Similar information is provided forthe military engine in Table 4. Notethat the compressor flow increaserequired (A%W2R) is 13.6% for themilitary engine versus 10.3\ for thecivil engine at CRP/IRP or CRP/TOP of1.465. This is a result of the militaryengine operating at 100 F hotter turbineinlet temperature at IRP than the civilengine at TOP although both engines arelimited to approximately the same turbinerotor inlet temperature at CRP.

Therefore, the increase intemperature from IRP to CRP is less forthe military engine than TOP to CRP forthe civil engine and this results in a

16

15

2 14

ce.o

13

12

CRP/TOP

O 1.15

D 1-35

I I96 10698 100 102 104

CORRECTED SPEED, ".

Fig. 1. Compressor flow-speed characteristic

Table 3. Cycle design point summary for civil rotorcraft

at constant takeoff power.

CYCLE REMATCIIEO CYCLE REMATCIIEONOM. CRP/TOP - 1.15 NOM. CRP/TOP - 1.25 MOM. CRP/TOP - 1.35 NOM. CSP/TOP . 1.45

A \ SFC

A \ Design Flow

A T41. °F

A \ NH

TOP CUP

BASE

BASE

BASE

BASE +i:5

TOP CRP

+ .72

+ 1.7

+140

BASE +3.2

TOP ' CRP

+ 1.6

+ 5.3

+215

-0.8 +3.2

TOP cap

+2.8

+10.3

+ 285

-l.S +3.2

Table 4. Cycle design point summary for military rotorcraft

at constant intermediate rated power.

CYCLE REMATCHED CYCLE REMATCHEDNOM. CRP/IRP - 1.15 NOM. CRP/IRP - 1.25 NOM. CRP/IRP - 1.35 NOM. CSP/IBP . 1.45

A \ SFC

A X Design Flou

A T41. °F

A \ NH

.

IBP cap

• BASE

BASE .

BASE

BASE +1.1

IRP CRP

+ .36

+ 1.0

+125

BASE +2.6

IRP

+ 1.0

+ 4.7

—-0.9

CSP

—

—+205

+2.S

IRP CRP

+1.B '

+13. S —

+205

-2.9 +0.7

relatively greater increase in compressorflow size required for the militaryengine. The absolute turbine temperaturelimit was set in order to keep the lowpressure turbine blades uncooled.

As the contingency power ratioincreases, the engine weight, specificfuel consumption, engine acquisition costand maintenance cost all increase for afixed IRP SHP level. These penalties aresubtracted from the benefits of idealizedengine and rotorcraft downsizing due toincreased contingency power ratio. Thenet gain is measured in terms ofdecreased vehicle gross weight, missionfuel burned, direct operating cost (DOC)and life cycle cost (LCC).

Rotocccaft DOC and LCC Models

The DOC and LCC models were developedfor the .baseline rotorcraft by SikorskyAircraft. The DOC model assumptions forthe civil commuter are listed in Table 5and the resulting DOC distribution isprovided in Fig. 2. Since the engine

1% DOC = 200 MILLiOn $ FOR FLEET OF 200 AIRCSAFTPER 2000 MRS/YEAR OVE.3. 10 YEARS

Fig. 2. OCC distribution/commuter

maintenance, insurance and depreciationcollectively represent only 10% of thetotal DOC distribution, engine downsizingresults in a modest DOC reduction.

The military engine LCC modelassumptions and resulting distribution of.LCC are shown in Table 6 and Fig 3. Notethat the % LCC changes are given from abase of the active weapon system LCC. notengine.only.Mission trade factors on a "rubberengine/rubber aircraft" basis relatechanges in engine contingency powerratio, weight, SFC. cost and maintenanceto changes in vehicle gross weight. LCC.DOC, acquisition cost, fuel burned, andOWE. These trade factors were used tocompare the net benefit of the variouscontingency power concepts designed forequal life.

Table 6. LCC model assumptions -military transport.

1983 Dollars

Production 400 Units

Marine Assault Helicopter -- LCC Model

Fuel $1.00/Gallon. 2% Spillage

Utilization 360 Hours/Year NominalValue

<,

20 - Year Life Cycle

Table 5. DOC model assumptions - commuter.

1983 Dollars

Production 400 Units

Development Cost Amortization intoVehicle Price

Airframe Maintenance Burden,1.5 X Direct Labor

Engine Burden 2 X Direct Labor

Insurance is 4% of Airliner Price

Aircraft Spares 15%

Engine Spares 30%

Straight-Line Depreciation. 10 Yearsto 25% Residual

Fuel $1.00/Gallon

Maintenance Labor $15/Hour

Utilization 2.000 Hours/Year

Crew Costs — $35.000/Year/Pilotfor Two Pilots (80 Hours/Month)

Off-Design Mission. 65% Load Factor

OTHERCON- / PERSONNEL

SUMA3LES 172

15 LCC = 100 MILLION $ FOR FLEETOF 400 AIRCRAFT FOR 20 YEARS

Fig. 3. LCC distribution/military transport

Novel Concents - Descriptions - DesignChanges. Operation. System Logistics

Several novel systems were proposedas alternative approaches to throttlepush to satisfy the increased CRPdemand. Each system was defined insufficient detail to permit an evaluationof required engine system designchanges. Seven systems were selected forpreliminary screening and five of thesewere chosen for quantitative evaluationas shown in Table 7.

Table 7. Contingency power concepts

Initial Screening Selected foe Evaluation

Throttle Push X

Cooling Flow Modulation (Rotor Only) X

Cooling Flow Modulation (Rotoc .& Stator) X

Watec Injection at Compressor Inlet

Water Injection into Turbine Cooling System X

Water Injection into Compressor and Turbine XCooling System

Solid Propellant Powered Emergency Power Unit (EPU)

Throttle Push

The throttle push concept isbasically the same as the baseline enginedesign with the hot section redesignedand cooling flow system resized forconstant engine life. Other designchanges such as increased diskthicknesses for higher RPM capability arealso required. No special add-on systemor components are required except forcontrol changes. These engines wereevaluated and compared to the baselineengines to account for all theconsequences of cycle changes required toobtain the higher contingency powerratios illustrated in Tables 3 and 4.The screening results indicated that thisapproach was promising enough to merit aquantitative evaluation.

Cooling Flow Modulation (Rotor Only)

In this concept, an electricallyoperated modulating valve is installedinto the rotor cooling supply so thatstage 1 HPT rotor blade cooling flow canbe adjusted from its maximum atcontingency power level to the lowerlevel adequate during the remainder ofits mission. This eliminates the overcooling penalty in the throttle pushconcept where cooling flows are set byblade coating limits at contingency andthen remain the same at other powerlevels. This concept was selected fordetailed study since the screeningresults were favorable.

Cooling Flow Modulation (Rotor and Stator)

This concept has the added feature ofalso increasing the cooling flow to the .stq 1 turbine shroud but only when

contingency power is required. Besidesreducing the SFC at powers < takeoffpower by lower cooling flows, thisfeature also takes on the function of anactive clearance control, by shrinkingthe shroud together with the blade, thusimproving turbine efficiency atcontingency vs. the rotor only coolingflow modulation system. Screeningresults also indicated merit in thisconcept worthy of quantitative evaluation.

Water Iniection into the Compressor Inlet

This is a well-known concept forincreasing engine power, where awater/antifreeze mixture is sprayedthrough, nozzles into the engine inlet.increasing engine flow and reducingcompressor work. However, the water/airratio is limited by the adverse effect ofwater on compressor stall margin. Also.the buildup compressor clearances must beopened up so that rubs do not occur whenthe water is turned on at contingencypower. This causes an-SFC penalty duringdry operation. In addition, the weightof water, tank, and nozzles becomes asizeable penalty, further aggravated bythe complexity and logistics problem ofsupplying demineralized water. As aresult of these screening results theconcept was eliminated from furtherconsideration.

Water Iniection into the •Turbine Cooling Flow

In this concept. HPT bladetemperatures during contingency operationare controlled by reducing the coolanttemperature by water evaporation ratherthan by an increased flow as in thecooling flow modulation system. . The

weight of water required per contingency-event is much less than that of thecompressor water concept but all theother system components (pump. tank,plumbing, valves and controls) are stillrequired. In addition, some losses willbe experienced as' a result of the stage 1HPT blade shrinkage which would increasetip clearance and thereby cause somereduction in turbine efficiency.Nevertheless, screening results werefavorable and this system was carriedinto the quantitative phase of the study.

Water Injection into Compressor andTurbine Cooling System

This combination was proposed todetermine if using one common water tank,and water system to serve two functionscould result in a synergistic advantage.This combination concept was also carriedthrough the quantitative phase of thestudy.

Solid Propellent Poveced EmergencyPower Unit

During the screening study, aseparate prime mover consistingof a turbine, a 3:1 gear reduction, andoverrunning clutch and a solid propellantcartridge were defined for weight and sizeestimates. The weight estimates for a 21/2 min. duration system were so large,that there was no net benefit. Thesystem was not competitive for durationsgreater than 30 seconds. Consequently,the concept was eliminated from furtherconsideration.

Duty Cycles. Life Analysis and Limits

Qualification Tests and Duty Cycles

In order to calculate relative lifeusage for the various engine componentsstudied, it was necessary to establishthe power levels and usage in the enginequalification test as well as during themission for both the civil and militaryengines.

Design changes for each contingencypower concept were made as required toensure engine life equal to the baselineengine life. These changes (primarilycooling flow increases) resulted in thedesign penalties included in theevaluation of each system.

For the civil engine, the testconditions are based on the FAA enginecertification requirements of 25 cycles,each of 6 hours duration. Contingencypower is reached twice during each cyclegiving a total time at contingency of 125minutes. The civil mission assumed is acombination of commuter and training

segments, based on in-service experiencewith commercial operators. The totalmission life is 2.000 hr/year for 10years or 20.000 hours, of which 500 hoursis allocated to. training flights. Onlyone 2 1/2 minute contingency, poweroperation is considered in the revenuepart of the mission, while it is assumedthat contingency power is reached for 30seconds once in each of 1,050 trainingflights, giving a total time atcontingency power of approximately 9hours for the composite mission.

The test conditions for the military •engine were based on the T700 "LAMPS"engine test requirements, which werethought to be representative for anyfuture military application. The 300hour test consists of fifty cycles,including four where contingency power isreached, for a total contingency time of10 minutes. The military engine missionwas derived from the time weightedaverage of eleven typical missions forthe HXM Marine Assault helicopter.The total mission life is 7,200 hoursover 20 years. Approximately 45 minutesof this time is spent at contingencypower, consisting of 144 usages in thepilot proficiency part of the mission.

A-summary of time at contingencypower together with those at other powersis shown in bar chart form in Fig. 4 forthe civil and military baseline engines.

CIVIL

MISSION TEST

20,000 ISOHRS. HRS.

//"TOP

100

0

| 80UIV»

g 602Oa.

t- 40

UJ

£ 20

n

-

r— i —

<MC

rCRP

TOP

MC

<MC

MILITARY

MIS5IO.N

7,200HRS .

TEST

300KSS .

icap•IRP

i?.?

<MC

Fig. 4. Time at power level for baseline engine

AC contingency power ratios greater thanthe 1.15 of the baseline engine theapportionment of power usage is unchangedfor the tests, but. has to be revised forthe missions due to downsizing theengines which results from higher CRP/IRPratios. In this case, the time spent atcontingency power is unchanged, while theremaining power spectrum has more time atrelatively higher powers than thebaseline engines.

Components Studied For Effects OfContingency Power

Fig. 5 lists the components which areaffected by speed or temperature atcontingency power and are considered tobe important in this study. The coldparts, compressor disks and impeller, areonly affected by an increase in speedfrom that of the baseline engine.Therefore, the stress levels in thesecomponents at contingency power may bebrought back to and can be maintained attheir original values by the addition ofmaterial. Higher gas temperatures bothinside and outside the combustor linercan be accommodated by more elaboratecooling air patterns and distributionwithout adversely affecting thetemperature profile at the combustordischarge. In the high pressure turbine,the nozzle and bucket airfoils aremaintained at their temperaturelimitations by the addition of coolingair and redesign of the airfoil coolingpassages. The first stage nozzle innerand outer bands require an additionalcooling system for_all contingency poweroperation.

COMPRESSOR

IMPELLER-!I—DISKS

The first stage shrouds need additionalimpingement cooling at contingency powerto stay below or at limitingtemperatures. There is no problem withthe cooling plates as there is enoughcooling to keep the cooling plates at orbelow acceptable temperature limits whenadditional air is required for bucketcooling. The high pressure, turbine rotordisks require additional material athigher contingency ratios to allow forspeed increases above the baseline enginelevel. The low pressure turbine bladesare uncooled. Consequently, when highercontingency ratios are encountered, amaterial change is required to retain theuncooled blade feature. To allow forhigher flowpath gas temperatures atcontingency power, the interturbine framehas an additional air insulation shieldaround the oil passages in the struts toprevent oil coking.

Limits

The high pressure turbine rotor bladeairfoils are limited by either themaximum surface temperature or the stresslevel, which are not necessarily at thesame location. The integrity of theairfoil coating during the life of theblade determines the maximum allowabletemperature. Creep.rupture determinesthe life of the blade. In the case ofthe Stage 1 bucket, the coatingtemperature becomes nore limiting thanthe creep rupture life requirements withincreasing contingency power ratio, suchthat the creep rupture life used •decreases with increasing contingencypower ratio.

HP TURBINE

NOZZLES 4 BANDSBLADESDISKSCOOLING PLATESSTATOR SHROUDS

COMBUSTOR

LINER

LP TU'SINE

FIRST STAGE VANEFIRST STAGE BLADE

Fig. 5. Components stud1e(j.

8

In the case of the nozzle airfoils,the limits are determined by the airfoilcoating allowable temperature or by thecreep rupture stress. The nozzle bladebands are limited by the coatingtempera ture.

The limit on the Stage 1 rotor shroudis the temperature at which the shroudsegment bonding material starts tooxidize which would lead to segmentseparation.

The cooling plates attached to thehigh pressure turbine rotor disks directcooling air into the blades on theforward sides of the disks. The rimtemperature is limiting as the thermalgradient is approximately equal to therim temperature minus the cooling airtemperature and this gradient determinesthe rim hoop stress.

Creep rupture is limiting in theStage 2 disk dovetail posts due to thedifficulty in obtaining enoughtemperature differential for coolingpurposes between the metal and coolingair.

The limiting consideration for theaxial and centrifugal compressor rotorsand also the high pressure turbine rotoris the burse margin Whenever the speedexceeds that of the baseline engine.

Life Analysis^

Component temperatures were generallycalculated from a turbine inlet gastemperature, which included an allowancefor combustor temperature distortion, anda cooling effectiveness, or the relativetemperature of an uncooled component.

Creep rupture stress and temperaturelimiting locations were based on analysesperformed on the baseline engine, as wasthe derivation of cooling effectivenesscurves versus cooling air flow for thecomponents studied, s

For the baseline engine first stagebucket, it was determined that LOOV lifeis used during the mission in the civilrotorcraft but during the qualificationtest for the military uses, for which thepower spectra are shown in Fig. 4. Therelative life usage for the civil andmilitary missions and tests are shown inFig. 6.

Fig. 7 shows an example of the lifesituation for the Stage 1 bucket militaryqualification test. Power usage has beendivided into three categories;contingency,, intermediate and partpower. The pare power component is acombination of all life usages for powers

below intermediate allowing for differentflight conditions, and is expressed as apercentage of intermediate rated power.

Whereas 100% of the life is used atthe baseline contingency power ratio of1.15, only 22% of the bucket life is usedat a 1.35 contingency power ratio due tothe increased cooling air required tosatisfy the blade coating temperaturelimitation at contingency power,resulting in much lower life usage atintermediate and part powers. A smallpart of the reduction in life usage atintermediate power .is due to a lowermetal temperature which results from thelower cooling air temperature for therematched engine at a contingency powerratio of 1.35.

In the life calculations, a speedincrease at contingency power above thatof the baseline engine is accounted forby decreasing the allowable metaltemperature .

CIVIL

MISSION TEST

20,000HRS.

150HRS.

MILITARY

MISSION TEST

7,200HRS.

AVJU

80

UJ

S 60in

UJu_- 40

X

20

n

-

CRP

TOP

<MC

rC?.?

/ 1

TOP |

300HRS.

rCRP

•CSP

IRP

<MC

Fig. 6. Stage 1 bucket life usage/baseline engine

Results and Summary

The engine design penalt ies werecompared to the r o t o r c r a f t b e n e f i t s foreach of the concepts and the net LCC orDOC reduct ion was then plot ted vs. theCRP/ IRP or CRP/TOP ra t io for each enginecons idered .

BASELINE ENGINE

CRP/IRP =• 1.15

TOTAL LIFE USED - 100%

THROTTLE PUSH ENGINE

CRP/IRP • 1.35

TOTAL LIFE USED - 22".

1000

100

£ 10

OPERATINGCONDITION^

0.1

IRP

1PP CRP

1000

100

£ 10

0.1

OPERATCON OH

ING

™>/

IRP

///

PP

CO

ATIN

GL

IMIT

CRP

/

/

1600 1800 2000 2200 2400

TMETAL * F

1600 1800 2000 2200 2400

THETAL " F

Fig. 7. Stage 1 bucket life usage in military qualification test

The results for LCC are plotted onFig. 8 vs. CRP/IRP. The results forthrottle push, cooling flow modulationand water into the turbine cooling systemcluster in a relatively narrow band. Thenet LCC reduction achievable is in- therange of 2.2% to 2.4% at the maximumuseful CRP/IRP ratio of 1.40. Combinedwater injection into the compressor andturbine cooling system is poorer than theother concepts, primarily due toadditional water weight and clearanceeffects. The nee advantage of designingfor higher CPR/IRP ratios diminishesbeyond 1.30 for all concepts. This iscaused by the rematching of engines tolower speeds at IRP in order to providefor compressor flow capacity ac CRP. Thesmall gain in net benefit beyond thebreakpoint in the net benefit curvessuggests that a balanced engineeringdesign should not aim at a CRP/IRP ratiohigher than 1.30.

The engine design penalties due toSFC. weight, cost and maintenance erodethe idealized benefits attainable byincreasing contingency'power ratios asshown on Fig. 9 for the throttle pushengines.

Results for the civil cotorcraft areshown on Fig. 10 with similar .trends.Throttle push and cooling flow modulationwere competitive systems with water intothe turbine cooling system showing upsomewhat poorer, especially at lowerCRP/TOP ratios.

-3 r-

-2

-1

•.*•'*•.

-*-*•-..

MAXUSABLE

I I

1.15 1.20 1.25 1.30 1.35 1.40

CONTINGENCY RATED POWER/INTERMEDIATE RATED POWER

CONCEPTS

THROTTLE PUSH

COOLING FLOW MODULATION(ROTOS ONLY)

- - COOLING FLOW MODULATION(ROTOR + STATIC PARTS)

-— H20 IN TURBINE COOLING FLOW

—X—X — H7° IN COMPRESSOR INLET AND* IN TURBINE COOLING FLOW

Fig. 8. Net LCC benefits/military rotorcraft

10

-6

-5

-4

-3

-2

-1

PENALTY DUE TO:

MAINTENANCE

WEIGHT

COSTSFC

NET

MAXUSABLE

1.15 1.20 1.25 1.30 .1.35 1.40

CONTINGENCY RATED POWER/INTERMEDIATE RATED POWER

Fig. 9. LCC penalties - throttle'push engine- military rotorcraft

-3 r

uOO

X<J

-2

-1' MAXUSABLE

I01.20 1.25 1.30 1.35 1.40 1.45 1.50.

CONTINGENCY RATED POWER/TAKEOFF POWER

CONCEPTS

THROTTLE PUSH

COOLING FLOW MODULATION(ROTOR ONLY)

. COOLING FLOW MODULATION(ROTOR + STATIC PARTS)

——-— H20 IN TURBINE COOLING FLOW

—X—X- H?° IN COMPRESSOR INLET AND* IN TURBINE COOLING FLOW

Fig. 10. Net DOC benefits/civil rotorcraft

The reversed slope of the throttle pushconcept is due to the more rapidlyincreasing fixed cooling flow penaltiesset at contingency power. The combinedcompressor inlet and turbine coolingwater injection system was again poorerthan the others as in the militaryrotorcraft. The effect of rematching athigher CEP/TOP ratios is also pronouncedfor the civil helicopter. At the highestCRP/TOP ratio of 1.50. cooling flowmodulation of rotor and stator has thelargest DOC reduction (2.7%). However,the DOC improvement between CRP/TOP ratioof 1.35 and 1.50 is modest, even for thebest system, and may not justify settingthe design objective at the highest ratio.

Proposed FAA Certification

There is an AIA proposal being circulatedin the Aircraft Industry for a new 30sec. "hot shot" rating with the objectiveof securing FAA review and approval. Theconcept is to set a rating which can beused only once in service, with perhapssome minor damage to engine components,but no signs of imminent failure. Theengine would require inspection andrepair after each use. The qualificationtest proposed, shown in Table 8. will setthe design life inasmuch, under thesecircumstances, there would be no use ofcontingency power during trainingmissions. Design limits were increased •to the criteria of some damage allowedfor the KPT nozzle bands and shroud, bondcoating.

This permitted cooling flowreductions and reduced the enginepenalties. Fig. 11 compares the engineDOC benefits for engines designed to theproposed certification vs. the currentcertification system.

The net benefit in DOC increases byup to 0.8* at CRP/TOP . 1.50. Theseresults and conclusions apply directlyonly to design options available in a newengine yet to be sized. The 30 sec. "hotshot" rating may show still more payoffvs. an engine redesigned to the currentracings if applied to derivatives ofcurrent engines.

Research and Technology Directions

A review of the engine designbarriers to increased contingency powercapability with lower penalties led to alist of Research and Technology needs forNASA planning.

11

Table 8. Proposed FAA rules

Results of AIA & SAE Committee Studies.

Proposed Changes in OKI Helicopter Rules in Industry Review.

Petition to FAA by Fall '83.

New Rule Published '85-'86

30 Second OEI 125% of Normal T/0

2 Minute OEI 110% of Normal T/0

Continuous Enroute OEI 100% of Normal T/0

30 Minute Certification Test Covers OEI Ratings.

"No Sign of Imminent Failure" after Certification Test,but Degree of Allowable Damage not Established.

Inspection required after in Service use.

Military Services may Adapt Similar Ratings in Future.

-4

-3

-2

COOLING FLOW MODULATION(ROTOR + STATIC PARTS)

MAX

USABLE

-11.20 1.25 1.30 1.35 1.40 1.45 1.50

CONTINGENCY RATED POWER/TAKEOFF POWER

LEGEND

PRESENT FAA REGULATION

PROPOSED FAA REGULATIONS

Fig. 11. Effect of proposed FAA regulationson DOC benefits

There are technology needs inmaterials development of improved turbineblade coatings, shroud bonding and shorttime creep rupture definition formaterials such as Mono N4 used in the hotsection parts.

Another need exists for developmentof a turbine cooling supply system forwider range of efficient cooling flowmodulation. Research and Technology inthese areas will also find application inother advanced engines.

A study was also recommended to applythese contingency power concepts toderivative engines.

References

1. VanFossen. G.J.. "Feasibility ofWater Injection into the TurbineCoolant to Permit Gas TurbineContingency Power for HelicopterApplications," NASA TM 83043. March.1983 - ' .

2. Brooks. A.. "The Impact ofContingency Ratings on AdvancedTurboshaft Engine Design". AHS HPS-1.November. 1979

3. Edkins, D. and Dugas, R.. "Gas•Turbine Engine Power Augmentation andEmergency Rating". NASA Contract No.DAAJ02-67-C-OOZ. Final ReportTF68-12. April. 1968

4. Hirschkron, R.. Manning. R.F. andHaynes. J.F.. "Contingency PowerStudy". NASA Contract NumberNAS3-21579. General Electric.TISrR81AEG045. October. 1981

12

1. Report No. NA$A TM_33579

A-84-40-66-7000

2. Government Accession No. 3. Recipient's Catalog No.

4. Title and Subtitle 5. Report Date

Contingency Power Concepts for Helicopter TurboshaftEngine 6. Performing Organization Code

505-42-62C

7. Author's)

Robert Hirschkron, Richard H. Davis,David N. Goldstein, John F. Haynes, andJames W. Gauntner

8. Performing Organization Report No.

E-2128

10. Work Unit No.

9. Performing Organization Name and Address

National Aeronautics and Spaee AdministrationLewis Research CenterCleveland, Ohio 44135

11. Contract or Grant No.

12. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, D.C. 20546

13. Type of Report and Period Covered

Technical Memorandum

14. Sponsoring Agency Code

15. Supplementary Notes

Robert Hirschkron, Richard H. Davis, David N. Goldstein, and John F. Haynes,General Electric Co., Preliminary Design, Lynn, Massachusetts; James W. Gauntner,NASA Lewis Research Center. Prepared for the Fortieth Annual Forum of the AmericanHelicopter Society, Arlington, Virginia, May 16-18, 1984.

16. Abstract

Twin helicopter engines are often sized by power requirement: of safe missioncompletion after the failure of one of the two engines. This study was undertakenfor NASA Lewis by General Electric Co. to evaluate the merits of special designfeatures to provde a 2-1/2 minute Contingency Power rating, permitting an enginesize reduction. The merits of water injection, cooling flow modulation, throttlepush and an auxiliary power plant were evaluated using military life cycle cost(LCC) and commercial helicopter direct operating cost (DOC) merit factors in a"rubber engine/rubber aircraft" scenario.

17. Key Words (Suggested by Authors))

Contingency powerHelicopterPropulsionTurboshaft engine

19. Security Classif. (of this report)

Unclassified

18. Distribution Statement

UnclassifiedSTAR Category

20. Security Classif. (of this page)

Unclassified

- unlimited07

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Washington, O.C.20546

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