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THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 99-GT-14Three Park Avenue, New York, N.Y. 10016-5990

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Copyright © 1999 by ASME

All Rights Reserved

Printed in U.S.A.

COOLED COOLING AIR SYSTEMS FOR TURBINETHERMAL MANAGEMENT

Greg B. Bruening and Won S. ChangTurbine Engine Division

Air Force Research LaboratoryWright-Patterson AFB, OH

ABSTRACT

This paper evaluates the feasibility and potential impact onoverall engine performance when utilizing the heat sinksources available in a gas turbine engine for improvedturbine thermal management. A study was conducted toassess the application of a heat exchanger to cool thecompressor bleed air normally used air for cooling turbinemachinery. The design tradeoffs of this cooled cooling airapproach as well as the methodology used to make theperformance assessment will be addressed.

The results of this study show that the use of a cooledcooling air (CCA) system can make a positive impact onoverall engine performance. Minimizing the complexity andweight of the CCA system, while utilizing advanced, hightemperature materials currently under development providethe best overall solution in terms of design risk and engineperformance.

NOMENCLATURE

CCA Cooled Cooling AirTSFC Thrust Specific Fuel Consumption

(lbm/lbf-hr)FN Net Thrust (lbf)OPR Overall Pressure RatioT4 High Pressure Turbine Rotor Inlet

Temperature (°F)T3 Compressor Exit Temperature ( °F)CMC Ceramic Matrix CompositeACM Air Cycle Machines Cooling Effectiveness

Tga Turbine Rotor Inlet Temperature (°F)

Tmetal Average Bulk Metal Temperature (°F)T., Cooling Air Temperature (°F)OTa;i Delta Air Temperature Across Heat

ExchangerMn Mach NumberBPR Engine Bypass RatioOD Outer DiameterCapture Ratio Percent Fan Bypass Air That Flows Through

Heat Exchanger%Wa, s Percent Total Engine Airflow That Enters

High Pressure CompressorSLS Sea Level Static Inlet ConditionMax AB Maximum AfterburnerFN/Wa Specific Thrust (lbf/lbm/sec)T/W Engine Thrust-to-Weight Ratio

INTRODUCTION

The need for improved engine performance will drivefuture turbine engines toward higher and higher operatingtemperatures. To achieve this, increased material temperaturecapability and improved cooling techniques have been a majorfocus in the turbine industry. However, further improvementsin these areas may be limited due to the time and costassociated with developing a new material that meets thehigher temperature requirements while maintaining sufficientstrength and manufacturability characteristics.

Significant progress was made in the 1960's to allow theturbine to reliably operate at gas temperatures that exceed the

Presented at the International Gas Turbine & Aeroengine Congress & ExhibitionIndianapolis, Indiana — June 7-June 10, 1999

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melting temperature of the turbine materials. Figure 1illustrates the trend in turbine inlet temperatures that hasresulted in significant improvements in engine performanceand aircraft capability. Today, the challenge of designingturbines to operate at higher gas temperatures continues. Inaddition, the desire for better specific fuel consumption (SFC)has driven engine designs toward higher pressure ratios,resulting in increased compressor bleed air temperatures.These higher temperatures make it very difficult tosufficiently cool the turbine with compressor discharge airwithout significantly penalizing the engine cycle performance.Therefore, new and innovative approaches will be necessaryto achieve the next level of performance capability, similar tothe improvements achieved with the introduction of turbineairfoil cooling.

4000

3500 --------^v^ s°lidific n Advanced

=- 3000Arcs Turbine Turbmc

-^j- Development

2500

Uroduction ^K Tmna rarurc

j000

500^

S^"x ~

CcrsnK^ Cnscnl Vines

1000 F. _ _ _ _ _._ _ _ II_ _ Dircctionafh_ TurbmcConvective Soiidif,-d

S00 Cooing Turb ine

1940 1950 1960 1970 1980 1990 2000 2010 2020

Production Or Demonstration Date

Figure 1 — Turbine Inlet Temperature Trends

One approach being considered today in the turbineengine community is the concept of first cooling thecompressor bleed air before it is used to cool the turbine. Aheat exchanger is added in the bleed air flowpath to transferthe heat from the bleed air to another source. Two potentialheat sink sources are the fan bypass air and the engine fuel.This concept significantly reduces cooling flow and turbinematerial temperatures, resulting in improved engineperformance and life.

The notional engine cycle considered for this study is anadvanced, variable cycle fighter engine as shown in Figure 2.The cycle and configuration is based on a projection ofavailable technologies associated with a year 2010-15 initialoperational capability (IOC). The variable cycle turbofanconcept consists of a two stage front fan, a core driven fanstage mechanically linked to a 4 stage high pressurecompressor (HPC), a single stage variable area high pressureturbine (HPT), and a two stage low pressure turbine. Thebasic cycle characteristics consist of an overall pressure ratio

(OPR) capability of 50, a fan pressure ratio of 8.5, and amaximum turbine rotor inlet temperature (T, ,) of 3 800°F.The component effeciencies assumed are consistent withcurrent technology trends. Applied to a typical fighter withthe capability to operate up to Mach 2.4 in the tropopause, thiscycle results in a maximum compressor exit temperature (T 3 )of 1600°F. This is the temperature of the bleed air extractedfrom the compressor. The high temperature T 3 and T41

conditions both contribute significantly to the challenge ofadequately cooling the turbine. The advanced materialsselected and the associated temperatures are based on thesuccessful transition of technology efforts currently underwayin industry. However, even with these materials, the need forCCA is not eliminated for the high operating temperaturesexpected of future engines.

Engine Cycle HP Turbine MaterialsVariable Cycle Fighter Engine Ceramic Matrix Composite Vane

(2010-2015 IOC) (2400°F Avg Bulk)Throttle Ratio = 1.06 Single Crystal Nickel BladeBypass Ratio = 0.4 (1950°F Avg Bulk)Overall Pressure Ratio = 50 Single Crustal Nickel ShroudFan Pressure Ratio = 8.5 (1950°F Avg Bulk)

^T41 = 3800°F Max Multi-Property Disk (1500 t Rim)

Figure 2 — Notional Advanced Variable Cycle Engine

COOLED COOLING AIR CONCEPTS

This study considered both an air-to-air and a fuel-to-airheat exchanger for cooling the compressor bleed air. Eachapproach assumes a CCA system capable of reducing thecompressor bleed air temperature by as much as 400°F at themaximum T, and T41 operating condition.

Figure 3 is an illustration of a fuel-to-air heat exchangersystem for cooling the HPT rotor, which includes both thedisk and blades. The CCA system was analyzed assuming anexternal heat exchanger in order to enhance maintainability ofthe system. The bleed air is taken off at the compressor exitthrough a bleed manifold. The bleed air is then cooled as it ispassed through a fuel-to-air heat exchanger and is eventuallyintroduced back into the bore of the engine through diffuserstruts. The bleed air then follows the same path that itnormally takes to eventually cool the rotor. The temperature


and pressure conditions at the low pressure turbine (LPT)allow it to be cooled with compressor interstage bleed and,therefore, does not require CCA. A small amount of CCA isalso used to cool the last compressor stage disk. The fuel isassumed to enter the heat exchanger at 250°F, assuming aheat load requirement similar to modem fighter aircraft. Theheated fuel exiting the heat exchanger is then injected intothe combustor as it normally would. For safetyconsiderations, this system includes a fuel bypass capabilityin case a fuel leak is detected in the heat exchanger.

Air

------- 250°FFuel

o I dAir

rFan

—Fuel

ass Svstem

H Shroud Shroud0

High Pressure — HP LP TurbineCompressor(HPC) Turbine

p V B V B V

Combustor t fi

pressure bleed air is further compressed through thecentrifugal compressor to overcome the pressure losses of theheat exchanger. A CCA system obviously becomes morecomplex with the addition of an ACM because of the rotatingmachinery and necessary control system to properly balancethe bleed flow split. This also negatively impacts the size andweight of the heat exchanger as well as the fuel temperature,which will be discussed later in this paper.

Fan

Figure 3 — Fuel-to-Air Heat Exchanger Concept

For the case illustrated in Figure 3, the HPT vanes donot require CCA. The temperature of the bleed air directlyfrom the compressor exit is adequate to cool the vanesbecause of the high temperature capability of the ceramicmatrix composite (CMC) material. However, if cooledCMC's were unavailable the turbine vane material is limitedto the 1950°F-nickel alloy material assumed for the turbineblade. As a result of using a lower temperature capablematerial, the vane would require CCA to achieve full life.This presents an additional challenge to the design becausethe turbine vane cooling air must have adequate pressuremargin to enter back into the core flowpath through the vanecooling holes. The turbine blade does not have this problembecause the bleed air pressure is increased by the pumpingeffect of the rotating turbine after it is injected into thecooling slots in the base of the rotor. For the turbine vane,however, the bleed air must overcome the bleed air pressurelosses from the heat exchanger. This is accomplished with anair cycle machine (ACM) which is added to the bleed airflowpath to increase its pressure. Figure 4 illustrates thisdesign. The ACM consists of a centrifugal compressor and aradial turbine connected by a common shaft. A portion ofthe high pressure bleed air is expanded through the radialturbine to drive the ACM compressor. This system isdesigned to allow the expanded bleed air to cool the lowpressure turbine (LPT) with adequate pressure andtemperature. This eliminates the need for using compressorinterstage bleed air to cool the LPT. The remaining high

Figure 4 — Fuel-to-Air Heat Exchanger Concept with ACM

A similar approach for cooling is to use an air-to-air heatexchanger. The heat exchanger is located in the fan bypassduct to utilize the cooler fan air to cool the compressordischarge bleed air. This approach assumes no CCA for theturbine vane, as in the fuel-to-air heat exchanger case.

Both the amount of air the heat exchanger must cool andthe level of temperature reduction required for cooling theturbine influence the size and weight of a heat exchanger. Theamount of bleed air to cool the turbine rotor, for instance, isdetermined from the cooling effectiveness characteristic of theturbine blade. The type of cooling technology assumed in theblade design ultimately determines the shape of the coolingcurve and directly impacts the amount of cooling air requiredfor a given cooling effectiveness. Both the engine cyclecharacteristics as well as the temperature capability of theblade material determine the required cooling effectiveness.The "advanced technology" cooling curve in Figure 5, whichassumes an advanced cooling utilizing quasi-transpiration or acombined impingement enhanced convection with advancedfilm cooling, significantly reduces the required cooling flowrate compared to the "current technology" cooling curve.Cooling the cooling air temperature by as much as 400°Freduces both the required cooling effectiveness and theamount of cooling air. The turbine blade design is lesschallenging with the lower cooling effectiveness. The reducedcooling air has a positive impact on engine performance.Without a heat exchanger, the turbine blade requires a more

3


aggressive cooling effectiveness of 0.84. The cooling flow — Bypassrate then becomes much more sensitive to increases in gas Ducttemperature as the curve flattens out.

Engine0.9 Advanced Core Bleed

No HEX Technology Air0.8 200 F s5.Tair — — — — —

40dF ATau0 .7 _ _ _ _ Current HEX Module Bypass Air

TechnologyC

La

0.6Figure 6 - Air-to Air HEX Installed In Fan Bypass Duct

° 0.5 Tgs - Tme tai

5 0.4 E - Tsaz - T`°°' Large differences in total pressure between two combiningTgas = Turbine Rotor Inlet Temp

LI 0 •i Tmetal =Avg Bulk Metal Temp streams can result in a large total pressure loss. This is due to

A Constant Tgas, Tmetal, Tcool ° THPC Bleed Air - OTair I large Mach number differences between the two streams0.2 T1dPC Bleed Air resulting in shear effects. It is desirable for the fan bypass air0.1

0 5 10 15passing through the heat exchanger to sustain minimalpressure losses to minimize a further pressure loss associated

Cooling Flow Rate (% Wa25) with recombining with the bypass air not passing through theheat exchanger. For the bleed air side, significant losses

Figure 5 - Turbine Blade Cooling Flow Requirements through the heat exchanger and air delivery pipes will result inadditional work required of the turbine rotor to pump the air

DESIGN CONSIDERATONS up through the blades with sufficient backflow margin. Theactual allowable pressure losses would depend on the specific

There are several design tradeoffs of a CCA approach design of the turbine and mixer components. However, for athat must be examined for it to be considered a feasible preliminary heat exchanger analysis, assumptions for maxsolution to improving engine performance. allowable pressure losses based on reasonable design practices

are made in order to do design tradeoffs. Figure 7 defines anThe heat exchanger itself must be compact, lightweight, allowable design space for the air-to-air heat exchanger that

and capable of operating in the high pressure and meets the cooling requirements of the engine configuration intemperature conditions of an engine environment. The heat this study. The design intent is to minimize the total CCAexchangers for this study assume a shell-tube type, cross- system weight while avoiding significant pressure losses. Theflow design. The tubes, case, and manifolds are made of a percent of fan bypass air that flows through the heatnickel alloy material. The heat exchanger is designed for its exchanger is defined as the capture ratio. The remaining fanmaximum heat transfer condition (3800°F T 4 , /1600°F T3) at bypass air passes around the heat exchanger and is not used2.4 Mn/50,000 ft. for cooling the bleed air. Figure 7 also illustrates design

tradeoffs of capture ratio with weight and pressure losses.An air-to-air heat exchanger system must be integrated

well with the fan bypass duct to minimize the im act to 20 Sneo-Tube Design}^ h 400 F pT, Bleed Air Side

engine size. A fighter-type engine cycle usually consists of a i 1/8" Tube OD, 10 mil thicknessInconel 625 Material

relatively low bypass ratio (BPR) with limited area in the 72`f 15 – – '5% Pressure Loss Fan Duct Length

40%bypass duct for additional hardware. Besides being compact, ¢Bleed Air Side) Constraint (20 40%

a fan duct heat exchanger must be designed structurally to \ 45°i°withstand foreign object damage as well as pressure sur ges 10a

during transient operation. The tubes inside the heat 'o'°^°60°%°exchanger will be exposed to high temperature, HPC ° Capture Ratiodischarge air. For this analysis, the CCA system consists of 5 =` ISYQ - - -

six heat exchangers located circumferentially within the fan Increasing Number Of Tubes

bypass duct. Figure 6 is an illustration of this configuration. 0The bleed air is distributed evenly among the six heat 0 100 200 300 400 500exchangers. This multiple heat exchanger design increases Total Cooled Cooling Air System Weight, lbs

the amount of bleed air pipes, but reduces the risk of acatastrophic engine failure in case of a single heat exchanger Figure 7 - Air-to-Air Cooled Cooling Air Sizing Criterialeak during flight.

4


The amount of cooling flow impacts the engineperformance and the weight of the CCA system. Figure 8compares the sensitivity of CCA system weight with coolingflow rate for both an air-to-air and a fuel-to-air heatexchanger system. For a fuel-to-air heat exchanger, both200°F and 400°F temperature reductions in the cooling airstream are illustrated. The weight of all the CCAcomponents is included, i.e., the air delivery pipes, thesensors and controls, the fuel bypass system, and theadditional hardware necessary to mount the heat exchangerto the engine case. The CCA weight is very sensitive to thecooling flow rate for an air-to-air system, compared to a fuel-to-air system. A fuel-to-air system has much greater heatsink potential for increases in cooling flow rate. The amountof cooling temperature reduction across the heat exchanger,i.e., 200°F versus 400°F, also influences weight sensitivity.

600 Total Weight Includes:Heat Exchanger (No ACM) A/A HEXAi Delivery Pipes 400°F ATzir - - -

500 Sensors/ControlsFuel Bypass System (F/A HEX Only)

400 - Misc. (Flanges, Clamps, Mounts, etc) - - - ----------

iIiIIIIJiIiIIIIIIIIJo i

I I 1400°F ATav

) 100 - F/AHEX-

200°F OTairNo HEX

00 5 10 15 20

Cooling Flow Rate (% Wa25)

Figure 8 — Cooled Cooling Air Sizing Comparison

mission, the fuel would be delivered to the combustor in eithera liquid or supercritical phase as heat loads change. This willrequire unique fuel control designs such as a liquid fuel by-pass loop and/or dual-phase fuel injectors in the combustor.

The complexity of the fuel system depends on the heatload placed on the fuel. Figure 9 illustrates the sensitivity thatthe cooling flow rate has on the fuel temperature for differentlevels of cooling air temperature reduction. For the case thatis cooling both the turbine vane and rotor, which includes anACM, the fuel becomes supercritical. This is a result of bothan increased amount of cooling flow and the additional heatadded to the air from the pressure rise through the ACM.Using a high temperature CMC vane material, however,eliminates the need for CCA for the vane which keeps the fuelsubcritical. Hence, the complexity and weight of the CCAsystem is reduced. This can be an important consideration inthe engine design. The tradeoff is with the increased riskassociated with development of a CMC vane material capableof high temperature applications.

LimitGumming 1 - 206)FOT,*lieP gslts (without ACM)

JP8+100

I IA Constant Tgas, Tmetal, No Hinx

THPC Bleed Air200

1000

u. 800

5)

It

Coking-Deposits

Fuel I1Critical

400°F ATair(with ACM)

400°F ATa - - - - -/ (withoult ACM)

600E

u400

10 15 20

With the fuel-to-air heat exchanger system, the impacton the fuel temperature is an important consideration. Theheat absorbed by the fuel causes its temperature to increasewhich introduces additional challenges to the fuel systemdesign. Current hydrocarbon fuels have an operatingtemperature limit of about 325°F. JP8+100 has beendeveloped recently which extends the temperature limit up to425°F. Temperatures above this limit cause the fuel to reactwith plumbing and form "gumming" deposits. This cancause fuel control valves to stick and fouling of the fuelnozzles and heat exchanger. Fuel systems that operate in thisrange may require maintenance to prevent these depositsfrom clogging the fuel system and heat exchanger.

As the heat loads increase, the fuel will operate above itscritical temperature limit (-700°F), which results in theformation of pyrolytic deposits. This can cause furtherfouling and fuel reaction with metal components. Inaddition, there are significant differences in fuel density as ittransitions into a supercritical fluid. Throughout an aircraft

Cooling Flow Rate (% Wa25)

Figure 9 — Fuel Temperatures For Various Fuel-To-AirCooling Concepts

It is interesting to note that advanced hydrocarbon fuelsare currently being developed to allow fuels to operate athigher operating temperatures without thermal decomposition[1]. Endothermic reactors are under consideration, as well,since they would increase cooling capacity.

ENGINE CYCLE IMPACT

The key objective of this study is to evaluate potentialengine performance benefits of a CCA system. For this study,engine specific fuel consumption, specific thrust, and thrust-to-weight ratio have been used to compare the various CCAconcepts.


To conduct this assessment, an engine modelingprogram was used to predict the performance of each cycle.Any modifications to the engine cycle impact the engineflowpath which ultimately affect its weight. An enginedesign program was used to generate a flowpath and weightestimate of each engine component. The overall engineweight can then be determined, based on inputs from thecycle model as well as the characteristics of the materialsassumed for each component. Similarly, the heat exchangercharacteristics were determined based on the engine coolingrequirements.

Cooled cooling air increases the overall pressure ratiocapability by allowing T 3 to operate significantly higher thancurrent engines. Table I compares a baseline cycle to thevarious approaches examined. The baseline cycle is limitedto 1400°F maximum T3 at the 2.4 Mn/50,000 ft. flightcondition,which reduces the overall pressure ratio from 50 to 32. Itassumes a 1950°F capable nickel alloy material for theturbine vane and blade and a 3800°F max T 41 . In satisfyingthe same turbine cooling requirements, each approachintroduces unique design challenges while having varyingeffects on overall engine performance.

Baseline No HEX A/A HEX F/AHEX F/AHEX(Mat'Is Only) (w/o ACM) (w/o ACM) (w/ ACM)

Overall Pressure Ratio 32 50 50 50 50

HPT Cooling Flow (% Wa25) 16.7% 24.7% 20.6% 20.6% 12.8%

Engine Bypass Ratio 0.42 0.33 0.51 0.44 0.39

Engine Core Corr. Flow, Ibm/sec 75.2 73.1 64.7 67.8 70.1

Cooled Cooling Air 0 0 240 63 230System Weight, lbs

Fuel Temperature 250 250 250 524 937(2.4Mn/50K), °F sub- sub- sub- sub- super-

critical critical critical critical critical

SFC, Ibm/hr/lbf (0.8Mn/40K) 0.922 -l.2% -4.6% -3.3% -2.4%(1.5Mn/50K) 1.17 +1.7% -1.7% -0.9% -1.7%

Specific Thrust, Dry, SLS 109.3 +0.7% -3.1% -0.6% +0.5%

Relative T/W Ratio Base +7.2% +4.3% +11,0% +2.0%(SLS Max A/B)

Table 1 - Engine Performance Results

To achieve higher OPR's without a CCA system, an enginedesign must rely more on advanced materials and/oradvanced cooling technology. The "materials only" cycle inTable I achieves a significant improvement in engine thrust-to-weight ratio (T/W). However, the increase in cooling flowpenalizes the cycle, resulting in only marginal improvementsin subsonic SFC and specific thrust. In addition, a high bladecooling effectiveness is required and the turbine disk materialmust be structurally capable of operating up to 1700°F.This presents a very high risk to the design relative to current

technology capability for highly loaded turbine disks. Ahigher temperature blade material would improveperformance by reducing the cooling flow, but the problem ofthe disk material remains. Similarly, the last stage of the HPCwill likely require the disk to be cooled, which can only beachieved with CCA.

The cycle with an air-to-air heat exchanger reduces theengine core size and weight by reducing the amount of bleedair required for turbine cooling. This also increases the enginebypass ratio which improves SFC but reduces specific thrust.The significant weight of the CCA system and the additionalpressure loss in the bypass duct due to the heat exchangerlimits the overall performance improvements. Also, high fanpressure ratios increase the fan duct air temperature whichlimits its heat sink capacity.

The cycle utilizing the fuel-to-air heat exchanger takesadvantage of the greater heat load capacity of the fuel versusthe fan duct air. This results in a more reasonable CCAsystem weight and size. The relatively compact heatexchanger could potentially be integrated into the engine corewhich would further reduce the complexity of the CCAsystem. The engine must be designed to accommodate thehigher fuel temperatures but by limiting the fuel to asubcritical phase, a dual-phase fuel delivery system is notrequired. The lighter weight CCA system, along with the lowdensity CMC vane material, significantly improves the enginethrust-to-weight ratio. SFC improves, as well, at about thesame specific thrust as the baseline. This appraoch appears tobe the best overall solution in balancing improved engineperformance with risk.

The cycle with an ACM uses a more conventional,1950°F capable vane material. However, the penaltiesassociated with this approach are substantial. Besides theincreased complexity of the fuel delivery system and controlsystem, the increased weight of the CCA system limits theimprovement to engine thrust-to-weight ratio compared to thebaseline cycle.

CONCLUDING REMARKS

Heat exchangers have been used for a long time inmechanical systems to improve the thermal management ofthe system. Aircraft today use heat exchangers to coolavionics components and the environmental control system.The use of a heat exchanger for turbine cooling application,however, presents some unique design challenges because itbecomes so closely integrated with the engine and cansignificantly affect the engine cycle.

6


The results suggest that a fuel-to-air heat exchangersystem offers the greatest potential for improved engineperformance while reducing some of the dependence onadvanced materials. Compared to fuel, fan air has limitedpotential as a heat sink. Also, the weight of an air-to-airsystem is very sensitive to potential increases in coolingflow requirements. For the fuel-to-air system, the key is tominimize its complexity and weight, i.e., eliminating theACM device by taking advantage of cooled ceramics for thevane. The additional challenges associated with a CCAsystem such as safety and reliability, however, must beaddressed by the engine research and developmentcommunity before these concepts will fmd their way intooperational systems.

ACKNOWLEDGEMENTS

The authors thank and acknowledge Jeffrey Stricker andChristopher Norden of the Air Force Research Laboratory atWright-Patterson Air Force Base for their assistance in theresearch and analysis that went into this paper.

REFERENCES

1. Edwards, T.,1993, "USAF Supercritical HydrocarbonFuel Interests, "AIAA Paper 93-0807.

2. Kays, W. M., 1984,"Compact Heat Exchangers," 3` d ed.,New York: McGraw-Hill.

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