COOLED COOLING AIR SYSTEMS FOR TURBINE THERMAL MANAGEMENT

THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS Three Park Avenue, New York, N.Y. 10016-5990 99-GT-14 S The Society shall not be responsible for statements or opinions advanced in papers or discussion at meetings of the Society or of its Divisions or 0 Sections, or printed in its publications. Discussion is printed only if the paper is published in an ASME Journal. Authorization to photocopy for internal or personal use is granted to libraries and other users registered with the Copyright Clearance Center (CCC) provided $3/article is paid to CCC, 222 Rosewood Dr., Danvers, MA 01923. Requests for special permission or bulk reproduction should be addressed to the ASME Technical Publishing Department. Copyright 1999 by ASME All Rights Reserved Printed in U.S.A. COOLED COOLING AIR SYSTEMS FOR TURBINE THERMAL MANAGEMENT Greg B. Bruening and Won S. Chang Turbine Engine Division Air Force Research Laboratory Wright-Patterson AFB, OH ABSTRACT This paper evaluates the feasibility and potential impact on overall engine performance when utilizing the heat sink sources available in a gas turbine engine for improved turbine thermal management. A study was conducted to assess the application of a heat exchanger to cool the compressor bleed air normally used air for cooling turbine machinery. The design tradeoffs of this cooled cooling air approach as well as the methodology used to make the performance assessment will be addressed. The results of this study show that the use of a cooled cooling air (CCA) system can make a positive impact on overall engine performance. Minimizing the complexity and weight of the CCA system, while utilizing advanced, high temperature materials currently under development provide the best overall solution in terms of design risk and engine performance. NOMENCLATURE CCA TSFC FN OPR T4 T 3 CMC ACM s Cooled Cooling Air Thrust Specific Fuel Consumption (lbm/lbf-hr) Net Thrust (lbf) Overall Pressure Ratio High Pressure Turbine Rotor Inlet Temperature ( F) Compressor Exit Temperature ( F) Ceramic Matrix Composite Air Cycle Machine Cooling Effectiveness Tga Tmetal T., OTa;i Mn BPR OD Capture Ratio %Wa, s SLS Max AB FN/Wa T/W INTRODUCTION Turbine Rotor Inlet Temperature ( F) Average Bulk Metal Temperature ( F) Cooling Air Temperature ( F) Delta Air Temperature Across Heat Exchanger Mach Number Engine Bypass Ratio Outer Diameter Percent Fan Bypass Air That Flows Through Heat Exchanger Percent Total Engine Airflow That Enters High Pressure Compressor Sea Level Static Inlet Condition Maximum Afterburner Specific Thrust (lbf/lbm/sec) Engine Thrust-to-Weight Ratio The need for improved engine performance will drive future turbine engines toward higher and higher operating temperatures. To achieve this, increased material temperature capability and improved cooling techniques have been a major focus in the turbine industry. However, further improvements in these areas may be limited due to the time and cost associated with developing a new material that meets the higher temperature requirements while maintaining sufficient strength and manufacturability characteristics. Significant progress was made in the 1960's to allow the turbine to reliably operate at gas temperatures that exceed the Presented at the International Gas Turbine & Aeroengine Congress & Exhibition Indianapolis, Indiana June 7-June 10, 1999

melting temperature of the turbine materials. Figure 1 illustrates the trend in turbine inlet temperatures that has resulted in significant improvements in engine performance and aircraft capability. Today, the challenge of designing turbines to operate at higher gas temperatures continues. In addition, 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 to sufficiently cool the turbine with compressor discharge air without significantly penalizing the engine cycle performance. Therefore, new and innovative approaches will be necessary to achieve the next level of performance capability, similar to the improvements achieved with the introduction of turbine airfoil cooling. (OPR) capability of 50, a fan pressure ratio of 8.5, and a maximum turbine rotor inlet temperature (T,,) of 3 800 F. The component effeciencies assumed are consistent with current technology trends. Applied to a typical fighter with the capability to operate up to Mach 2.4 in the tropopause, this cycle results in a maximum compressor exit temperature (T 3 ) of 1600 F. This is the temperature of the bleed air extracted from the compressor. The high temperature T 3 and T41 conditions both contribute significantly to the challenge of adequately cooling the turbine. The advanced materials selected and the associated temperatures are based on the successful transition of technology efforts currently underway in industry. However, even with these materials, the need for CCA is not eliminated for the high operating temperatures expected of future engines. 4000 3500 --------^ v^ s lidific n Advanced Arcs Turbine Turbmc =- 3000 -^j- Development 2500 U j000 roduction ^K Tmnararurc 500 S ^ ^"x ~ ^ Cnscnl 1000 F. _. _ II Dircctionafh_ Turbmc Convective Soiidif,-d S00 Cooing Turbin e CcrsnK Vines 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 turbine engine community is the concept of first cooling the compressor bleed air before it is used to cool the turbine. A heat exchanger is added in the bleed air flowpath to transfer the heat from the bleed air to another source. Two potential heat sink sources are the fan bypass air and the engine fuel. This concept significantly reduces cooling flow and turbine material temperatures, resulting in improved engine performance and life. The notional engine cycle considered for this study is an advanced, variable cycle fighter engine as shown in Figure 2. The cycle and configuration is based on a projection of available technologies associated with a year 2010-15 initial operational capability (IOC). The variable cycle turbofan concept consists of a two stage front fan, a core driven fan stage mechanically linked to a 4 stage high pressure compressor (HPC), a single stage variable area high pressure turbine (HPT), and a two stage low pressure turbine. The basic cycle characteristics consist of an overall pressure ratio Engine Cycle HP Turbine Materials Variable Cycle Fighter Engine Ceramic Matrix Composite Vane (2010-2015 IOC) (2400 F Avg Bulk) Throttle Ratio = 1.06 Single Crystal Nickel Blade Bypass Ratio = 0.4 (1950 F Avg Bulk) Overall Pressure Ratio = 50 Single Crustal Nickel Shroud Fan 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-air heat exchanger for cooling the compressor bleed air. Each approach assumes a CCA system capable of reducing the compressor bleed air temperature by as much as 400 F at the maximum T, and T 41 operating condition. Figure 3 is an illustration of a fuel-to-air heat exchanger system for cooling the HPT rotor, which includes both the disk and blades. The CCA system was analyzed assuming an external heat exchanger in order to enhance maintainability of the system. The bleed air is taken off at the compressor exit through a bleed manifold. The bleed air is then cooled as it is passed through a fuel-to-air heat exchanger and is eventually introduced back into the bore of the engine through diffuser struts. The bleed air then follows the same path that it normally 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 is also used to cool the last compressor stage disk. The fuel is assumed to enter the heat exchanger at 250 F, assuming a heat load requirement similar to modem fighter aircraft. The heated fuel exiting the heat exchanger is then injected into the combustor as it normally would. For safety considerations, this system includes a fuel bypass capability in case a fuel leak is detected in the heat exchanger. pressure bleed air is further compressed through the centrifugal compressor to overcome the pressure losses of the heat exchanger. A CCA system obviously becomes more complex with the addition of an ACM because of the rotating machinery and necessary control system to properly balance the bleed flow split. This also negatively impacts the size and weight of the heat exchanger as well as the fuel temperature, which will be discussed later in this paper. Air HrFan ------- 250 F Fuel o I dair ass Svstem Shroud Shroud 0 High Pressure HP LP Turbine Turbine Compressor(HPC) p V B V B V Combustor t fi Fan Figure 3 Fuel-to-Air Heat Exchanger Concept Figure 4 Fuel-to-Air Heat Exchanger Concept with ACM For the case illustrated in Figure 3, the HPT vanes do not require CCA. The temperature of the bleed air directly from the compressor exit is adequate to cool the vanes because of the high temperature capability of the ceramic matrix composite (CMC) material. However, if cooled CMC's were unavailable the turbine vane material is limited to the 1950 F-nickel alloy material assumed for the turbine blade. As a result of using a lower temperature capable material, the vane would require CCA to achieve full life. This presents an additional challenge to the design because the turbine vane cooling air must have adequate pressure margin to enter back into the core flowpath through the vane cooling holes. The turbine blade does not have this problem because the bleed air pressure is increased by the pumping effect of the rotating turbine after it is injected into the cooling slots in the base of the rotor. For the turbine vane, however, the bleed air must overcome the bleed air pressure losses from the heat exchanger. This is accomplished with an air cycle machine (ACM) which is added to the bleed air flowpath to increase its pressure. Figure 4 illustrates this design. The ACM consists of a centrifugal compressor and a radial turbine connected by a common shaft. A portion of the high pressure bleed air is expanded through the radial turbine to drive the ACM compressor. This system is designed to allow the expanded bleed air to cool the low pressure turbine (LPT) with adequate pressure and temperature. This eliminates the need for using compressor interstage bleed air to cool the LPT. The remaining high A similar approach for cooling is to use an air-to-air heat exchanger. The heat exchanger is located in the fan bypass duct to utilize the cooler fan air to cool the compressor discharge bleed air. This approach assumes no CCA for the turbine vane, as in the fuel-to-air heat exchanger case. Both the amount of air the heat exchanger must cool and the level of temperature reduction required for cooling the turbine influence the size and weight of a heat exchanger. The amount of bleed air to cool the turbine rotor, for instance, is determined from the cooling effectiveness characteristic of the turbine blade. The type of cooling technology assumed in the blade design ultimately determines the shape of the cooling curve and directly impacts the amount of cooling air required for a given cooling effectiveness. Both the engine cycle characteristics as well as the temperature capability of the blade material determine the required cooling effectiveness. The "advanced technology" cooling curve in Figure 5, which assumes an advanced cooling utilizing quasi-transpiration or a combined impingement enhanced convection with advanced film cooling, significantly reduces the required cooling flow rate compared to the "current technology" cooling curve. Cooling the cooling air temperature by as much as 400 F reduces both the required cooling effectiveness and the amount of cooling air. The turbine blade design is less challenging with the lower cooling effectiveness. The reduced cooling 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 rate then becomes much more sensitive to increases in gas temperature as the curve flattens out. Engine Bypass Duct 0.9 Advanced Core No HEX Technology Air 0.8 200 F s5.tair 40dF ATau 0 Current.7 HEX Module Bypass Air Technology C 0.6 La 0.5 Figure 6 - Air-to Air HEX Installed In Fan Bypass Duct Tgs - Tmetai 5 0.4 E - Tsaz - T` ' Large differences in total pressure between two combining Tgas = Turbine Rotor Inlet Temp streams can result in a large total pressure loss. This is due to LI 0 i Tmetal =Avg Bulk Metal Temp A Constant Tgas, Tmetal, Tcool THPC Bleed Air - OTair Ilarge Mach number differences between the two streams 0.2 T1dPC Bleed Air resulting in shear effects. It is desirable for the fan bypass air 0.1 passing through the heat exchanger to sustain minimal 0 5 10 15 pressure losses to minimize a further pressure loss associated Cooling Flow Rate (% Wa25) with recombining with the bypass air not passing through the heat 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 in additional work required of the turbine rotor to pump the air DESIGN CONSIDERATONS up through the blades with sufficient backflow margin. The actual 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 a that must be examined for it to be considered a feasible preliminary heat exchanger analysis, assumptions for max solution to improving engine performance. allowable pressure losses based on reasonable design practices are made in order to do design tradeoffs. Figure 7 defines an The 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 in temperature conditions of an engine environment. The heat this study. The design intent is to minimize the total CCA exchangers for this study assume a shell-tube type, cross- system weight while avoiding significant pressure losses. The flow design. The tubes, case, and manifolds are made of a percent of fan bypass air that flows through the heat nickel alloy material. The heat exchanger is designed for its exchanger is defined as the capture ratio. The remaining fan maximum heat transfer condition (3800 F T 4,/1600 F T3) at bypass air passes around the heat exchanger and is not used 2.4 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 Bleed i 1/8" Tube OD, 10 mil thickness Inconel 625 Material relatively low bypass ratio (BPR) with limited area in the 72 `f 15 '5% Pressure Loss Fan Duct Length 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 a 10 during transient operation. The tubes inside the heat 'o' ^ 60 % exchanger will be exposed to high temperature, HPC Capture Ratio discharge 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. 0 The bleed air is distributed evenly among the six heat 0 100 200 300 400 500 exchangers. This multiple heat exchanger design increases Total Cooled Cooling Air System Weight, lbs the amount of bleed air pipes, but reduces the risk of a catastrophic engine failure in case of a single heat exchanger Figure 7 - Air-to-Air Cooled Cooling Air Sizing Criteria leak during flight. 4

i The amount of cooling flow impacts the engine performance and the weight of the CCA system. Figure 8 compares the sensitivity of CCA system weight with cooling flow rate for both an air-to-air and a fuel-to-air heat exchanger system. For a fuel-to-air heat exchanger, both 200 F and 400 F temperature reductions in the cooling air stream are illustrated. The weight of all the CCA components is included, i.e., the air delivery pipes, the sensors and controls, the fuel bypass system, and the additional hardware necessary to mount the heat exchanger to the engine case. The CCA weight is very sensitive to the cooling flow rate for an air-to-air system, compared to a fuelto-air system. A fuel-to-air system has much greater heat sink potential for increases in cooling flow rate. The amount of cooling temperature reduction across the heat exchanger, i.e., 200 F versus 400 F, also influences weight sensitivity. o 600 Total Weight Includes: Heat Exchanger (No ACM) A/A HEX Ai Delivery Pipes 400 F ATzir - - - 500 Sensors/Controls Fuel Bypass System (F/A HEX Only) 400 - Misc. (Flanges, Clamps, Mounts, etc) - - - ---------- iiiiiiijiiiiiiiiiiij I I 1400 F ATav ) 100- - F/AHEX 200 F OTair No HEX 00 5 10 15 20 Cooling Flow Rate (% Wa25) Figure 8 Cooled Cooling Air Sizing Comparison With the fuel-to-air heat exchanger system, the impact on the fuel temperature is an important consideration. The heat absorbed by the fuel causes its temperature to increase which introduces additional challenges to the fuel system design. Current hydrocarbon fuels have an operating temperature limit of about 325 F. JP8+100 has been developed recently which extends the temperature limit up to 425 F. Temperatures above this limit cause the fuel to react with plumbing and form "gumming" deposits. This can cause fuel control valves to stick and fouling of the fuel nozzles and heat exchanger. Fuel systems that operate in this range may require maintenance to prevent these deposits from clogging the fuel system and heat exchanger. As the heat loads increase, the fuel will operate above its critical temperature limit (-700 F), which results in the formation of pyrolytic deposits. This can cause further fouling and fuel reaction with metal components. In addition, there are significant differences in fuel density as it transitions into a supercritical fluid. Throughout an aircraft mission, the fuel would be delivered to the combustor in either a liquid or supercritical phase as heat loads change. This will require unique fuel control designs such as a liquid fuel bypass loop and/or dual-phase fuel injectors in the combustor. The complexity of the fuel system depends on the heat load placed on the fuel. Figure 9 illustrates the sensitivity that the cooling flow rate has on the fuel temperature for different levels of cooling air temperature reduction. For the case that is cooling both the turbine vane and rotor, which includes an ACM, the fuel becomes supercritical. This is a result of both an increased amount of cooling flow and the additional heat added 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 fuel subcritical. Hence, the complexity and weight of the CCA system is reduced. This can be an important consideration in the engine design. The tradeoff is with the increased risk associated with development of a CMC vane material capable of high temperature applications. 1000 u. 800 5) E u 600 400 200 JP8+100 It Coking -Deposits Fuel I1 Critical Limit 400 F ATair (with ACM) Gumming 1-206)FOT,* liep gslts (without ACM) I I A Constant Tgas, Tmetal, No Hinx THPC Bleed Air 400 F ATa - - - - - / (withoult ACM) 10 15 20 Cooling Flow Rate (% Wa25) Figure 9 Fuel Temperatures For Various Fuel-To-Air Cooling Concepts It is interesting to note that advanced hydrocarbon fuels are currently being developed to allow fuels to operate at higher 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 potential engine performance benefits of a CCA system. For this study, engine specific fuel consumption, specific thrust, and thrustto-weight ratio have been used to compare the various CCA concepts.

To conduct this assessment, an engine modeling program was used to predict the performance of each cycle. Any modifications to the engine cycle impact the engine flowpath which ultimately affect its weight. An engine design program was used to generate a flowpath and weight estimate of each engine component. The overall engine weight can then be determined, based on inputs from the cycle model as well as the characteristics of the materials assumed for each component. Similarly, the heat exchanger characteristics were determined based on the engine cooling requirements. Cooled cooling air increases the overall pressure ratio capability by allowing T 3 to operate significantly higher than current engines. Table I compares a baseline cycle to the various approaches examined. The baseline cycle is limited to 1400 F maximum T 3 at the 2.4 Mn/50,000 ft. flight condition, which reduces the overall pressure ratio from 50 to 32. It assumes a 1950 F capable nickel alloy material for the turbine vane and blade and a 3800 F max T 41. In satisfying the same turbine cooling requirements, each approach introduces unique design challenges while having varying effects 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 63 230 System Weight, lbs Fuel Temperature 250 250 250 524 937 (2.4Mn/50K), F sub- sub- sub- sub- supercritical 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 engine design must rely more on advanced materials and/or advanced cooling technology. The "materials only" cycle in Table I achieves a significant improvement in engine thrustto-weight ratio (T/W). However, the increase in cooling flow penalizes the cycle, resulting in only marginal improvements in subsonic SFC and specific thrust. In addition, a high blade cooling effectiveness is required and the turbine disk material must 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. A higher temperature blade material would improve performance by reducing the cooling flow, but the problem of the disk material remains. Similarly, the last stage of the HPC will likely require the disk to be cooled, which can only be achieved with CCA. The cycle with an air-to-air heat exchanger reduces the engine core size and weight by reducing the amount of bleed air required for turbine cooling. This also increases the engine bypass ratio which improves SFC but reduces specific thrust. The significant weight of the CCA system and the additional pressure loss in the bypass duct due to the heat exchanger limits the overall performance improvements. Also, high fan pressure ratios increase the fan duct air temperature which limits its heat sink capacity. The cycle utilizing the fuel-to-air heat exchanger takes advantage of the greater heat load capacity of the fuel versus the fan duct air. This results in a more reasonable CCA system weight and size. The relatively compact heat exchanger could potentially be integrated into the engine core which would further reduce the complexity of the CCA system. The engine must be designed to accommodate the higher fuel temperatures but by limiting the fuel to a subcritical phase, a dual-phase fuel delivery system is not required. The lighter weight CCA system, along with the low density CMC vane material, significantly improves the engine thrust-to-weight ratio. SFC improves, as well, at about the same specific thrust as the baseline. This appraoch appears to be the best overall solution in balancing improved engine performance with risk. The cycle with an ACM uses a more conventional, 1950 F capable vane material. However, the penalties associated with this approach are substantial. Besides the increased complexity of the fuel delivery system and control system, the increased weight of the CCA system limits the improvement to engine thrust-to-weight ratio compared to the baseline cycle. CONCLUDING REMARKS Heat exchangers have been used for a long time in mechanical systems to improve the thermal management of the system. Aircraft today use heat exchangers to cool avionics components and the environmental control system. The use of a heat exchanger for turbine cooling application, however, presents some unique design challenges because it becomes so closely integrated with the engine and can significantly affect the engine cycle. 6

The results suggest that a fuel-to-air heat exchanger system offers the greatest potential for improved engine performance while reducing some of the dependence on advanced materials. Compared to fuel, fan air has limited potential as a heat sink. Also, the weight of an air-to-air system is very sensitive to potential increases in cooling flow requirements. For the fuel-to-air system, the key is to minimize its complexity and weight, i.e., eliminating the ACM device by taking advantage of cooled ceramics for the vane. The additional challenges associated with a CCA system such as safety and reliability, however, must be addressed by the engine research and development community before these concepts will fmd their way into operational systems. ACKNOWLEDGEMENTS The authors thank and acknowledge Jeffrey Stricker and Christopher Norden of the Air Force Research Laboratory at Wright-Patterson Air Force Base for their assistance in the research and analysis that went into this paper. REFERENCES 1. Edwards, T.,1993, "USAF Supercritical Hydrocarbon Fuel Interests, "AIAA Paper 93-0807. 2. Kays, W. M., 1984,"Compact Heat Exchangers," 3` d ed., New York: McGraw-Hill. VA