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Copyright © 1999 by ASME All Rights Reserved Printed in U.S.A. Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78606/V003T01A002/2412146/v003t01a002-99-gt-014.pdf by guest on 27 September 2021
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 Tga Turbine Rotor Inlet Temperature (°F) Tmetal Average Bulk Metal Temperature (°F) This paper evaluates the feasibility and potential impact on T., Cooling Air Temperature (°F) overall engine performance when utilizing the heat sink OTa;i Delta Air Temperature Across Heat sources available in a gas turbine engine for improved Exchanger turbine thermal management. A study was conducted to Mn Mach Number assess the application of a heat exchanger to cool the BPR Engine Bypass Ratio compressor bleed air normally used air for cooling turbine OD Outer Diameter machinery. The design tradeoffs of this cooled cooling air Capture Ratio Percent Fan Bypass Air That Flows Through approach as well as the methodology used to make the Heat Exchanger performance assessment will be addressed. %Wa,s Percent Total Engine Airflow That Enters High Pressure Compressor The results of this study show that the use of a cooled SLS Sea Level Static Inlet Condition cooling air (CCA) system can make a positive impact on Max AB Maximum Afterburner overall engine performance. Minimizing the complexity and FN/Wa Specific Thrust (lbf/lbm/sec) weight of the CCA system, while utilizing advanced, high T/W Engine Thrust-to-Weight Ratio temperature materials currently under development provide the best overall solution in terms of design risk and engine performance. INTRODUCTION NOMENCLATURE The need for improved engine performance will drive CCA Cooled Cooling Air future turbine engines toward higher and higher operating TSFC Thrust Specific Fuel Consumption temperatures. To achieve this, increased material temperature (lbm/lbf-hr) capability and improved cooling techniques have been a major FN Net Thrust (lbf) focus in the turbine industry. However, further improvements OPR Overall Pressure Ratio in these areas may be limited due to the time and cost T4 High Pressure Turbine Rotor Inlet associated with developing a new material that meets the Temperature (°F) higher temperature requirements while maintaining sufficient T3 Compressor Exit Temperature ( °F) strength and manufacturability characteristics. CMC Ceramic Matrix Composite ACM Air Cycle Machine Significant progress was made in the 1960's to allow the s Cooling Effectiveness 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 (OPR) capability of 50, a fan pressure ratio of 8.5, and a illustrates the trend in turbine inlet temperatures that has maximum turbine rotor inlet temperature (T, ,) of 3 800°F. resulted in significant improvements in engine performance The component effeciencies assumed are consistent with and aircraft capability. Today, the challenge of designing current technology trends. Applied to a typical fighter with turbines to operate at higher gas temperatures continues. In the capability to operate up to Mach 2.4 in the tropopause, this addition, the desire for better specific fuel consumption (SFC) cycle results in a maximum compressor exit temperature (T 3 ) has driven engine designs toward higher pressure ratios, of 1600°F. This is the temperature of the bleed air extracted
resulting in increased compressor bleed air temperatures. from the compressor. The high temperature T 3 and T41 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78606/V003T01A002/2412146/v003t01a002-99-gt-014.pdf by guest on 27 September 2021 These higher temperatures make it very difficult to conditions both contribute significantly to the challenge of sufficiently cool the turbine with compressor discharge air adequately cooling the turbine. The advanced materials without significantly penalizing the engine cycle performance. selected and the associated temperatures are based on the Therefore, new and innovative approaches will be necessary successful transition of technology efforts currently underway to achieve the next level of performance capability, similar to in industry. However, even with these materials, the need for the improvements achieved with the introduction of turbine CCA is not eliminated for the high operating temperatures airfoil cooling. expected of future engines.
4000
35 ------ v ^ s°lidific n Advanced A cs u i e u mc =- 3000 -^j- Development
5 roduction ^K Tmnararurc U
5 S Cc s K Engine Cycle HP Turbine Materials " C sc i es 1 _ _ _ _ _ _ _ _ II_ _ i cc io a _ u mc Variable Cycle Fighter Engine Ceramic Matrix Composite Vane Co ec i e Soii i - (2010-2015 IOC) (2400°F Avg Bulk) S Cooi g u i b Throttle Ratio = 1.06 Single Crystal Nickel Blade
40 0 60 0 80 0 2000 20 0 2020 Bypass Ratio = 0.4 (1950°F Avg Bulk) Production Or Demonstration Date Overall Pressure Ratio = 50 Single Crustal Nickel Shroud Fan Pressure Ratio = 8.5 (1950°F Avg Bulk) Figure 1 — Turbine Inlet Temperature Trends ^T41 = 3800°F Max Multi-Property Disk (1500 t Rim)
One approach being considered today in the turbine Figure 2 — Notional Advanced Variable Cycle Engine 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 COOLED COOLING AIR CONCEPTS the heat from the bleed air to another source. Two potential heat sink sources are the fan bypass air and the engine fuel. This study considered both an air-to-air and a fuel-to-air This concept significantly reduces cooling flow and turbine heat exchanger for cooling the compressor bleed air. Each material temperatures, resulting in improved engine approach assumes a CCA system capable of reducing the performance and life. compressor bleed air temperature by as much as 400°F at the maximum T, and T 41 operating condition. The notional engine cycle considered for this study is an advanced, variable cycle fighter engine as shown in Figure 2. Figure 3 is an illustration of a fuel-to-air heat exchanger The cycle and configuration is based on a projection of system for cooling the HPT rotor, which includes both the available technologies associated with a year 2010-15 initial disk and blades. The CCA system was analyzed assuming an operational capability (IOC). The variable cycle turbofan external heat exchanger in order to enhance maintainability of concept consists of a two stage front fan, a core driven fan the system. The bleed air is taken off at the compressor exit stage mechanically linked to a 4 stage high pressure through a bleed manifold. The bleed air is then cooled as it is compressor (HPC), a single stage variable area high pressure passed through a fuel-to-air heat exchanger and is eventually turbine (HPT), and a two stage low pressure turbine. The introduced back into the bore of the engine through diffuser basic cycle characteristics consist of an overall pressure ratio 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) pressure bleed air is further compressed through the allow it to be cooled with compressor interstage bleed and, centrifugal compressor to overcome the pressure losses of the therefore, does not require CCA. A small amount of CCA is heat exchanger. A CCA system obviously becomes more also used to cool the last compressor stage disk. The fuel is complex with the addition of an ACM because of the rotating assumed to enter the heat exchanger at 250°F, assuming a machinery and necessary control system to properly balance heat load requirement similar to modem fighter aircraft. The the bleed flow split. This also negatively impacts the size and heated fuel exiting the heat exchanger is then injected into weight of the heat exchanger as well as the fuel temperature, the combustor as it normally would. For safety which will be discussed later in this paper. Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78606/V003T01A002/2412146/v003t01a002-99-gt-014.pdf by guest on 27 September 2021 considerations, this system includes a fuel bypass capability in case a fuel leak is detected in the heat exchanger.
Ai
------ 5 ° —Fuel o I Ai
ass Svstem a a S ou S ou
ig essu e — u i e u i e Com esso ( C Com us o t i
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 A similar approach for cooling is to use an air-to-air heat not require CCA. The temperature of the bleed air directly exchanger. The heat exchanger is located in the fan bypass from the compressor exit is adequate to cool the vanes duct to utilize the cooler fan air to cool the compressor because of the high temperature capability of the ceramic discharge bleed air. This approach assumes no CCA for the matrix composite (CMC) material. However, if cooled turbine vane, as in the fuel-to-air heat exchanger case. CMC's were unavailable the turbine vane material is limited to the 1950°F-nickel alloy material assumed for the turbine Both the amount of air the heat exchanger must cool and blade. As a result of using a lower temperature capable the level of temperature reduction required for cooling the material, the vane would require CCA to achieve full life. turbine influence the size and weight of a heat exchanger. The This presents an additional challenge to the design because amount of bleed air to cool the turbine rotor, for instance, is the turbine vane cooling air must have adequate pressure determined from the cooling effectiveness characteristic of the margin to enter back into the core flowpath through the vane turbine blade. The type of cooling technology assumed in the cooling holes. The turbine blade does not have this problem blade design ultimately determines the shape of the cooling because the bleed air pressure is increased by the pumping curve and directly impacts the amount of cooling air required effect of the rotating turbine after it is injected into the for a given cooling effectiveness. Both the engine cycle cooling slots in the base of the rotor. For the turbine vane, characteristics as well as the temperature capability of the however, the bleed air must overcome the bleed air pressure blade material determine the required cooling effectiveness. losses from the heat exchanger. This is accomplished with an The "advanced technology" cooling curve in Figure 5, which air cycle machine (ACM) which is added to the bleed air assumes an advanced cooling utilizing quasi-transpiration or a flowpath to increase its pressure. Figure 4 illustrates this combined impingement enhanced convection with advanced design. The ACM consists of a centrifugal compressor and a film cooling, significantly reduces the required cooling flow radial turbine connected by a common shaft. A portion of rate compared to the "current technology" cooling curve. the high pressure bleed air is expanded through the radial Cooling the cooling air temperature by as much as 400°F turbine to drive the ACM compressor. This system is reduces both the required cooling effectiveness and the designed to allow the expanded bleed air to cool the low amount of cooling air. The turbine blade design is less pressure turbine (LPT) with adequate pressure and challenging with the lower cooling effectiveness. The reduced temperature. This eliminates the need for using compressor cooling air has a positive impact on engine performance. interstage bleed air to cool the LPT. The remaining high Without a heat exchanger, the turbine blade requires a more
3 aggressive cooling effectiveness of 0.84. The cooling flow — Bypass rate then becomes much more sensitive to increases in gas Duct temperature as the curve flattens out. Engine Bleed 0.9 Advanced Core No HEX Technology Air 0.8 200 F s5.Tair — — — — — 40dF_ _A _ _ Current 0 7 HEX Module Bypass Air Technology
C Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78606/V003T01A002/2412146/v003t01a002-99-gt-014.pdf by guest on 27 September 2021 0.6
a Figure 6 - Air-to Air HEX Installed In Fan Bypass Duct ° 0.5 gs - metai 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 I large 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 (% Wa 5 wi ecom i i g wi e y ass ai o assi g oug e ea e c a ge o e ee ai si e sig i ica osses igu e 5 - u i e a e Coo i g ow equi eme s oug e ea e c a ge a ai e i e y i es wi esu i a i io a wo k equi e o e u i e o o o um e ai ESIG CO SI E A O S u oug e a es wi su icie ack ow ma gi e ac ua a owa e essu e osses wou e e o e s eci ic e e a e se e a esig a eo s o a CCA a oac esig o e u i e a mi e com o e s owe e o a a mus e e ami e o i o e co si e e a easi e e imi a y ea e c a ge a a ysis assum io s o ma so u io o im o i g e gi e e o ma ce a owa e essu e osses ase o easo a e esig ac ices a e ma e i o e o o esig a eo s igu e 7 e i es a e ea e c a ge i se mus e com ac ig weig a owa e esig s ace o e ai - o-ai ea e c a ge a a ca a e o o e a i g i e ig essu e a mee s e coo i g equi eme s o e e gi e co igu a io i em e a u e co i io s o a e gi e e i o me e ea is s u y e esig i e is o mi imi e e o a CCA e c a ge s o is s u y assume a s e - u e y e c oss- sys em weig w i e a oi i g sig i ica essu e osses e ow esig e u es case a ma i o s a e ma e o a e ce o a y ass ai a ows oug e ea icke a oy ma e ia e ea e c a ge is esig e o i s e c a ge is e i e as e ca u e a io e emai i g a ma imum ea a s e co i io (3 ° /1 ° 3 a y ass ai asses a ou e ea e c a ge a is o use M /5 o coo i g e ee ai igu e 7 a so i us a es esig a eo s o ca u e a io wi weig a essu e osses A ai - o-ai ea e c a ge sys em mus e i eg a e S eo- u e esig we wi e a y ass} uc o mi imi e e im ac o ee Ai Si e i 1/8" Tube OD, 10 mil thickness e gi e si e A ig e - y e e gi e cyc e usua y co sis s o a Inconel 625 Material e a i e y ow y ass a io ( wi imi e a ea i e 7 `f 15 – – '5% Pressure Loss Fan Duct Length y ass uc o a i io a a wa e esi es ei g com ac ¢ l d A r S d Constraint (20 %% a a uc ea e c a ge mus e esig e s uc u a y o 5°i° wi s a o eig o ec amage as we as essu e su g es 1 u i g a sie o e a io e u es i si e e ea °%° o ° °
e c a ge wi e e ose o ig em e a u e C ISYQ° - - - Ca u e a io isc a ge ai o is a a ysis e CCA sys em co sis s o 5 = si ea e c a ge s oca e ci cum e e ia y wi i e a Increasing Number Of Tubes y ass uc igu e is a i us a io o is co igu a io e ee ai is is i u e e e y amo g e si ea 1 3 5 e c a ge s is mu i e ea e c a ge esig i c eases Total Cooled Cooling Air System Weight, lbs e amou o ee ai i es u e uces e isk o a ca as o ic e gi e ai u e i case o a si g e ea e c a ge igu e 7 - Ai - o-Ai Coo e Coo i g Ai Si i g C i e ia eak u i g ig
The amount of cooling flow impacts the engine mission, the fuel would be delivered to the combustor in either performance and the weight of the CCA system. Figure 8 a liquid or supercritical phase as heat loads change. This will compares the sensitivity of CCA system weight with cooling require unique fuel control designs such as a liquid fuel by- flow rate for both an air-to-air and a fuel-to-air heat pass loop and/or dual-phase fuel injectors in the combustor. exchanger system. For a fuel-to-air heat exchanger, both 200°F and 400°F temperature reductions in the cooling air The complexity of the fuel system depends on the heat stream are illustrated. The weight of all the CCA load placed on the fuel. Figure 9 illustrates the sensitivity that the cooling flow rate has on the fuel temperature for different
components is included, i.e., the air delivery pipes, the Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78606/V003T01A002/2412146/v003t01a002-99-gt-014.pdf by guest on 27 September 2021 sensors and controls, the fuel bypass system, and the levels of cooling air temperature reduction. For the case that additional hardware necessary to mount the heat exchanger is cooling both the turbine vane and rotor, which includes a to the engine case. The CCA weight is very sensitive to the ACM, the fuel becomes supercritical. This is a result of both cooling flow rate for an air-to-air system, compared to a fuel- an increased amount of cooling flow and the additional heat to-air system. A fuel-to-air system has much greater heat added to the air from the pressure rise through the ACM. sink potential for increases in cooling flow rate. The amount Using a high temperature CMC vane material, however, of cooling temperature reduction across the heat exchanger, eliminates the need for CCA for the vane which keeps the fuel i.e., 200°F versus 400°F, also influences weight sensitivity. subcritical. Hence, the complexity and weight of the CCA system is reduced. This can be a important consideration in the engine design. The tradeoff is with the increased risk o a Weig I c u es associated with development of a CMC vane material capable ea E c a ge ( o ACM A/A HEX A e i e y i es 400°F AT ir - - - of high temperature applications. 500 Se so s/Co o s ue y ass Sys em ( A E O y 1000 ------Misc ( a ges C am s Mou s e c I 400°F ATair Coking (wi ACM)
-Deposits u. 800 400°F ATa - - - - - Fuel I1 / (wi ou ACM) iIiIIII iIiIIIIIIII 5 Critical I I 1400°F ATav i Limit o Gumming 1 - 206 FOT,* 1 - - F/AHEX 600 (wi ou ACM 200°F OTair E lieP gslts No HEX JP8+100 00 5 1 15 u 400 Coo i g l a e (% Wa 5 I I A Co s a gas me a o Hinx THPC ee Air Figure 8 — Cooled Cooling Air Sizing Comparison 200 10 15 20 With the fuel-to-air heat exchanger system, the impact Cooling Flow Rate (% Wa25) on the fuel temperature is an important consideration. The heat absorbed by the fuel causes its temperature to increase Figure 9 — Fuel Temperatures For Various Fuel-To-Air which introduces additional challenges to the fuel system Cooling Concepts design. Current hydrocarbon fuels have an operating temperature limit of about 325°F. JP8+100 has been It is interesting to note that advanced hydrocarbon fuels developed recently which extends the temperature limit up to are currently being developed to allow fuels to operate at 425°F. Temperatures above this limit cause the fuel to react higher operating temperatures without thermal decomposition with plumbing and form "gumming" deposits. This can [1]. Endothermic reactors are under consideration, as well, cause fuel control valves to stick and fouling of the fuel since they would increase cooling capacity. nozzles and heat exchanger. Fuel systems that operate in this range may require maintenance to prevent these deposits ENGINE CYCLE IMPACT from clogging the fuel system and heat exchanger. The key objective of this study is to evaluate potential As the heat loads increase, the fuel will operate above its engine performance benefits of a CCA system. For this study, critical temperature limit (-700°F), which results in the engine specific fuel consumption, specific thrust, and thrust- formation of pyrolytic deposits. This can cause further to-weight ratio have been used to compare the various CCA fouling and fuel reaction with metal components. In concepts. addition, there are significant differences in fuel density as it transitions into a supercritical fluid. Throughout an aircraft To conduct this assessment, an engine modeling technology capability for highly loaded turbine disks. A program was used to predict the performance of each cycle. higher temperature blade material would improve Any modifications to the engine cycle impact the engine performance by reducing the cooling flow, but the problem of flowpath which ultimately affect its weight. An engine the disk material remains. Similarly, the last stage of the HPC design program was used to generate a flowpath and weight will likely require the disk to be cooled, which can only be estimate of each engine component. The overall engine achieved with CCA. weight can then be determined, based on inputs from the
cycle model as well as the characteristics of the materials The cycle with a air-to-air heat exchanger reduces the Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78606/V003T01A002/2412146/v003t01a002-99-gt-014.pdf by guest on 27 September 2021 assumed for each component. Similarly, the heat exchanger engine core size and weight by reducing the amount of bleed characteristics were determined based on the engine cooling air required for turbine cooling. This also increases the engine requirements. bypass ratio which improves SFC but reduces specific thrust. Cooled cooling air increases the overall pressure ratio The significant weight of the CCA system and the additional capability by allowing T 3 to operate significantly higher than pressure loss in the bypass duct due to the heat exchanger current engines. Table I compares a baseline cycle to the limits the overall performance improvements. Also, high fan various approaches examined. The baseline cycle is limited pressure ratios increase the fan duct air temperature which to 1400°F maximum T3 at the 2.4 Mn/50,000 ft. flight limits its heat sink capacity. condition, The cycle utilizing the fuel-to-air heat exchanger takes which reduces the overall pressure ratio from 50 to 32. It advantage of the greater heat load capacity of the fuel versus assumes a 1950°F capable nickel alloy material for the the fan duct air. This results in a more reasonable CCA turbine vane and blade and a 3800°F max T 1 . In satisfying system weight and size. The relatively compact heat the same turbine cooling requirements, each approach exchanger could potentially be integrated into the engine core introduces unique design challenges while having varying which would further reduce the complexity of the CCA effects on overall engine performance. system. The engine must be designed to accommodate the higher fuel temperatures but by limiting the fuel to a /A E ase i e o E A/A E /A E subcritical phase, a dual-phase fuel delivery system is not (Ma Is O y (w/o ACM (w/o ACM (w/ ACM required. The lighter weight CCA system, along with the low O e a essu e a io 3 5 5 5 5 density CMC vane material, significantly improves the engine Coo i g ow (% Wa 5 1 7% 7% % % 1 % thrust-to-weight ratio. SFC improves, as well, at about the
E gi e y ass a io 33 51 39 same specific thrust as the baseline. This appraoch appears to
E gi e Co e Co ow I m/sec 75 73 1 7 7 7 1 be the best overall solution in balancing improved engine performance with risk. Coo e Coo i g Ai 3 3 Sys em Weig s ACM uses a more conventional, ue em e a u e 5 5 5 5 937 The cycle with a ( M /5 K ° su - su - su - su - su e - 1950°F capable vane material. However, the penalties c i ica c i ica c i ica c i ica c i ica associated with this approach are substantial. Besides the
S C I m/ / ( M / K 9 - % - % -3 3% - % increased complexity of the fuel delivery system and control (1 5M /5 K 1 17 +1 7% -1 7% - 9% -1 7% system, the increased weight of the CCA system limits the
S eci ic us y S S 1 9 3 + 7% -3 1% - % + 5% improvement to engine thrust-to-weight ratio compared to the
e a i e /W a io ase +7 % + 3% +11 % + % baseline cycle. (S S Ma A/
Table 1 - Engine Performance Results CO C U I G EMA KS
To achieve higher OPR's without a CCA system, a engine Heat exchangers have been used for a long time in design must rely more on advanced materials and/or mechanical systems to improve the thermal management of advanced cooling technology. The "materials only" cycle in the system. Aircraft today use heat exchangers to cool Table I achieves a significant improvement in engine thrust- avionics components and the environmental control system. to-weight ratio (T/W). However, the increase in cooling flow The use of a heat exchanger for turbine cooling application, penalizes the cycle, resulting in only marginal improvements however, presents some unique design challenges because it in subsonic SFC and specific thrust. In addition, a high blade becomes so closely integrated with the engine and can cooling effectiveness is required and the turbine disk material significantly affect the engine cycle. must be structurally capable of operating up to 1700°F. This presents a very high risk to the design relative to current
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 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78606/V003T01A002/2412146/v003t01a002-99-gt-014.pdf by guest on 27 September 2021 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.
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