Gas Turbine Engine with Air/Fuel Heat Exchanger

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(12) (io) Patent No.:

United States Patent US 9,771,867 B2

Karam et al. (45)

Date of Patent: Sep. 26,2017

(54) GAS TURBINE ENGINE WITH AIR/FUEL (58) Field of Classification Search HEAT EXCHANGER CPC ........ F02C 7/224; F02C 7/143; F28D 7/0058; F28D 7/1653; F28D 21/0001; (71) Applicant: Rolls-Royce North American (Continued) Technologies, Inc., Indianapolis, IN (US) (56) References Cited (72) Inventors: Michael Abraham Karam, Plainfield, U.S. PATENT DOCUMENTS IN (US); Eric Sean Donovan, Fishers, IN (US); Michael Stephen Krautheim, 2,294,254 A * 8/1942 Throckmorton ......... C10G 9/20 Fountaintown, IN (US); Daniel Kent 122/356 Vetters, Indianapolis, IN (US); Donald 2,407,165 A * 9/1946 Kreitner .................. F02C 7/143 G. Chouinard, Westfield, IN (US) 60/39.53 (Continued) (73) Assignees: Rolls-Royce Corporation, Indianapolis, IN (US); Rolls-Royce North American FOREIGN PATENT DOCUMENTS Technologies, Inc., Indianapolis, IN (US) EP 2128419 Al 12/2009 FR 2482196 Al 11/1981 (*) Notice: Subject to any disclaimer, the term ofthis patent is extended or adjusted under 35 OTHER PUBLICATIONS U.S.C. 154(b) by 523 days.

International Search Report and Written Opinion, PCT/US2012/ (21) Appl. No.: 14/319,680 072117, Aug. 27, 2013.

(Continued) (22) Filed: Jun. 30, 2014 (65) Prior Publication Data Primary Examiner Pascal M Bui Pho Assistant Examiner Eric Linderman US 2014/0338334 Al Nov. 20, 2014 (74) Attorney, Agent, or Firm Fishman Stewart PLLC Related U.S. Application Data (63) Continuation of application No.

(57) ABSTRACT PCT/US2012/072117, filed on Dec. 28, 2012.

One embodiment ofthe present invention is a unique aircraft (Continued) propulsion gas turbine engine. Another embodiment is a unique gas turbine engine. Another embodiment is a unique (51) Int. Cl.

gas turbine engine. Other embodiments include apparatuses, F02C 7/224 (2006.01) systems, devices, hardware, methods, and combinations for F02C 7/143 (2006.01) gas turbine engines with heat exchange systems. Further (Continued) embodiments, forms, features, aspects, benefits, and advan- (52) U.S. Cl.

tages of the present application will become apparent from CPC .............. F02C 7/224 (2013.01); F02C 7/141 the description and figures provided herewith.

(2013.01); F02C 7/143 (2013.01); F28D 7/0058 (2013.01); 19 Claims, 3 Drawing Sheets (Continued) 30-, X

US 9,771,867 B2

Page 2 5,392,595 A * 2/1995 Glickstein ............... F02C 7/224 Related U.S. Application Data 60/39.12 (60) Provisional application No. 61/581,850, filed on Dec. 5,553,448 A * 9/1996 Farrell ...................... F02C 3/10 29/401.1 30, 2011.

5,678,408 A 10/1997 Janes 5,724,806 A * 3/1998 Horner ...................... F02C 7/12 (51) Int. Cl.

60/728 F28F 9/26 (2006.01) 5,782,076 A * 7/1998 Huber ....................... F02C 7/08 F28D 21/00 (2006.01) 415/115 F02C 7/141 (2006.01) 5,791,148 A * 8/1998 Burrus ...................... FOID 5/18 F28F 9/02 (2006.01) 60/749 5,992,139 A * 11/1999 Kesseli ..................... F02C 6/12 F28D 7/00 (2006.01) 60/39.183 F28D 7/16 (2006.01) 6,253,554 B1 * 7/2001 Kobayashi ................ F02C 7/12 (52) U.S. Cl.

60/736 CPC ....... F28D 7/1653 (2013.01); F28D 21/0001 6,357,113 B1 * 3/2002 Williams ................ B23P 15/26 (2013.01); F28F 9/026 (2013.01); F28F 9/26 29/890.034 6,415,595 B1 * 7/2002 Wilmot, Jr. .......... B64D 13/006 (2013.01); F28F 20091029 (2013.01); Y02T 60/266 501671 (2013.01); Y02T 50/675 (2013.01) 6,422,020 B1 7/2002 Rice (58) Field of Classification Search 6,430,931 B1 * 8/2002 Horner .................... F02C 7/143 CPC .......... F28D 21/0008; F28D 2021/0026; F28F 60/785 9/026; F28F 9/26 6,672,072 B1 * 1/2004 Giffin, III ................. FOID 5/08 See application file for complete search history. 60/728 6,691,519 B2 * 2/2004 Little ........................ F02C 1/00 415/912 (56) References Cited 7,334,407 B2 2/2008 Spadaccini et al.

7,459,081 B2 12/2008 Koenig et al.

U.S. PATENT DOCUMENTS 7,568,336 B2 8/2009 Brault et al.

7,870,743 B2 * 1/2011 Lee ......................... FO1D 5/189 2,479,071 A * 8/1949 Henstridge ........... F28D 7/0058 416/96 R 165/155 7,874,139 B2 * 1/2011 Briesch ................. FO1K 23/068 2,575,683 A * 11/1951 Price ....................... F02C 3/107 60/39.281 123/41.19 7,926,289 B2 * 4/2011 Lee ......................... FO1D 5/081 2,608,054 A * 8/1952 Price ..................... B64C 21/025 415/145 290A C 7,926,292 B2 * 4/2011 Rabovitser .............. F02C 3/205 2,718,753 A * 9/1955 Bridgeman ............. F02C 7/143 60/730 60/39.465 7,954,324 B2 * 6/2011 Holland .................... F02C 6/18 3,064,947 A * 11/1962 Wynne .................. F28D 1/0471 60/736 165/149 8,721,265 B1 * 5/2014 Brostmeyer ............ F04D 19/02 3,228,464 A * 1/1966 Stein ..................... F28D 9/0012 415/1 165/166 8,943,827 B2 * 2/2015 Prociw .................... FO1D 25/12 3,235,001 A * 2/1966 Giannotti .................. FO IN 5/00 60/730 165/135 9,394,828 B2 * 7/2016 Eleftheriou ............... F02C 7/08 3,398,538 A * 8/1968 Hall .......................... F02C 7/08 2002/0144664 Al* 10/2002 Haldeman, III ........ C10L 1/328 60/262 123/25 B 3,735,588 A * 5/1973 Moskowitz ............... F28D 7/08 2004/0055740 Al* 3/2004 Meshenky .......... F02B 29/0462 165/135 165/125 3,747,339 A * 7/1973 Wolf ......................... F02C 7/08 2005/0235626 Al* 10/2005 Hull .......................... F02C 7/08 60/206 60/39.511 3,779,007 A * 12/1973 Lavash ..................... F02C 7/14 2005/0262848 Al* 12/2005 Joshi ....................... F02C 3/305 60/241 60/772 3,831,674 A * 8/1974 Stein ......................... F02C 7/08 2006/0064987 Al* 3/2006 Veninger ................. F02C 7/224 165/166 60/777 4,254,618 A * 3/1981 Elovic ..................... F02C 7/185 2006/0080967 Al* 4/2006 Colket, III .............. F23C 6/045 60/226.1 60/777 4,506,502 A * 3/1985 Shapiro ................... F02C 3/073 2006/0090472 Al* 5/2006 Ritland ................... F02C 7/047 60/39.43 60/772 4,949,544 A * 8/1990 Hines .................... FOIK 21/047 2006/0174627 Al* 8/2006 McQuiggan .............. F02C 6/18 60/728 60/772 4,993,223 A * 2/1991 Kretzinger ............ F28D 9/0018 2007/0089423 Al* 4/2007 Norman .................. F02C 7/143 165/166 60/772 5,004,044 A * 4/1991 Horgan ................. F28D 9/0012 2007/0101731 Al* 5/2007 Bayt ...................... B64D 13/06 165/145 62/7 5,050,668 A * 9/1991 Peterson ................... F02C 7/08 2008/0083226 Al* 4/2008 Joshi ......................... F02C 3/34 165/166 60/772 5,161,365 A * 11/1992 Wright ...................... F02C 7/16 2008/0083608 Al* 4/2008 Cipollini ............ BO1D 19/0084 60/39.461 204/157.5 5,165,224 A * 11/1992 Spadaccini ............. F02B 51/02 2008/0142189 Al* 6/2008 Norris ....................... F02C 7/14 60/723 165/11.1 5,185,997 A * 2/1993 Nishijima ............... F02C 7/185 2008/0202094 Al* 8/2008 Brault ..................... F02C 7/143 60/740 5,207,053 A * 5/1993 Spadaccini ............. F02B 51/02 60/226.1 2008/0310955 Al* 12/2008 Norris ....................... F02C 7/14 60/723 415/178 5,317,877 A * 6/1994 Stuart ..................... F02C 7/185 2009/0133400 Al* 5/2009 Callas ..................... F02C 7/224 60/736 5,335,501 A 8/1994 Taylor 60/730

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Page 3 2014/0260178 At * 9/2014 Eleftheriou ............... F02C 7/08 (56) References Cited 60/39.511 2015/0047367 At* 2/2015 Benignos ................ F02C 7/143 U.S. PATENT DOCUMENTS 60/782 2015/0298024 At* 10/2015 Watkins ................. 13O11)3/146 2010/0058801 At * 3/2010 Masani ................. FO1D 15/005 203/24 62/611 2010/0319359 At* 12/2010 Holt .......................... F02C 6/08 2015/0337730 At* 11/2015 Kupiszewski ............ F02C 3/22 60/39.465 60/782 2016/0177829 At* 6/2016 Loebig ...................... F02C 7/18 2012/0128467 At* 5/2012 Ruthemeyer ........... FO1D 5/146 415/115 415/1 2016/0237901 At* 8/2016 Zelesky .................. F02C 7/143 2012/0199335 At* 8/2012 Maurer ................. F28D 7/0058 2016/0237904 At* 8/2016 Scarboro ................. F02C 7/143 165/185 2016/0305324 At* 10/2016 Magowan ................. F02C 7/18 2012/0216543 At * 8/2012 Eleftheriou ............... F02C 7/08 2016/0312702 At* 10/2016 Thomas .................... F02C 7/18 60/772 2013/0186059 At* 7/2013 Epstein ..................... F02C 3/22 60/205 OTHER PUBLICATIONS 2013/0219915 At* 8/2013 Prociw .................... F02C 7/224 60/782 English language translation of FR2482196, Boudigues, Nov. 13, 2013/0239542 At* 9/2013 Dasgupta ................ F02C 7/143 1981.

60/39.093 2014/0216045 At * 8/2014 Gueh ........................ F02C 6/10 * cited by examiner 60/772

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U.S. Patent Sep. 26,2017 US 9,771,867 B2

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---------------------- rr ---~ 18 --- --74- --- FIG. 1 COMPRESSOR STAGE COMPRESSOR STAGE HEAT EXCHANGER 36 26 FUEL SUPPLY FUEL INJECTORS FIG. 2

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U.S. Patent Sep. 26,2017 US 9,771,867 B2

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GAS TURBINE ENGINE WITH AIR/FUEL DETAILED DESCRIPTION HEAT EXCHANGER For purposes of promoting an understanding of the prin- ciples of the invention, reference will now be made to the CROSS REFERENCE TO RELATED 5 embodiments illustrated in the drawings, and specific lan- APPLICATIONS guage will be used to describe the same. It will nonetheless be understood that no limitation ofthe scope ofthe invention The present application is a continuation of PCT Appli- is intended by the illustration and description of certain cation No.PCT/US2012/072117,filed Dec. 28, 2012, which embodiments of the invention. In addition, any alterations claims the benefit ofU.S. Provisional Patent Application No.

io and/or modifications of the illustrated and/or described 61/581,850 filed Dec. 30, 2011, each of which is incorpo- embodiment(s) are contemplated as being within the scope rated herein by reference.

of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or GOVERNMENT RIGHTS described herein, as would normally occur to one skilled in 15 the art to which the invention pertains, are contemplated as The present application was made with United States being within the scope of the present invention.

government support under Contract No. NNH08ZEA001 N Referring to the drawings, and in particular FIG. 1, some Amendment #2, awarded by NASA. The United States aspects of a non-limiting example of an engine 10 in government may have certain rights in the present applica- accordance with an embodiment ofthe present invention are tion.

20 schematically depicted. Engine 10 is an aircraft propulsion gas turbine engine. Engine 10 includes a compressor system FIELD OF THE INVENTION 12, a combustion system 14 in fluid communication with compressor system 12, and a turbine system 16 in fluid The present invention relates to gas turbine engines, and communication with combustion system 14. In one form, more particularly, to gas turbine engines with heat exchange 25 compressor system 12, combustion system 14 and turbine systems.

system 16 are disposed about an engine centerline 18, e.g., the axis of rotation of compressor system 12 and turbine BACKGROUND system 16. In other embodiments, other arrangements may be employed. In various embodiments, engine 10 may be a Gas turbine heat exchange systems that effectively trans- 30 single spool engine or a multi-spool engine. In various fer heat from pressurized compressor air to fuel remain an embodiments, engine 10 may or may not have a turbine area of interest. Some existing systems have various short- system, or may have additional turbomachinery components comings, drawbacks, and disadvantages relative to certain in addition to a compressor system and/or a turbine system, applications. Accordingly, there remains a need for further e.g., a fan system. In some embodiments, engine 10 may be contributions in this area of technology.

35 a direct propulsion engine that produces thrust directly from combustion system 14. In other embodiments, combustion SUMMARY system 14 may form a gas generator for a gas turbine propulsion system, or may be employed in a gas turbine One embodiment of the present invention is a unique engine topping cycle. In still other embodiments, engine 10 aircraft propulsion gas turbine engine. Another embodiment 40 may be one or more of other types of gas turbine engines, is a unique gas turbine engine. Another embodiment is hybrid engines and/or combined cycle engines.

another unique gas turbine engine. Other embodiments Compressor system 12 includes a compressor case 20 that include apparatuses, systems, devices, hardware, methods, houses stationary and rotating compressor system 12 com- and combinations for gas turbine engines with heat ponents. In various embodiments, compressor case 20 may exchange systems. Further embodiments, forms, features, 45 be formed of one or more individual compressor case aspects, benefits, and advantages of the present application structures, e.g., depending on the number, size and location will become apparent from the description and figures of compressor stages and/or the number of spools employed provided herewith.

in engine 10. Combustion system 14 includes a combustor BRIEF DESCRIPTION OF THE DRAWINGS case 22, a combustor 24 and a plurality of fuel injectors 26.

50 Combustor 24 receives pressurized air from compressor The description herein makes reference to the accompa- system 12. Fuel injectors 26 are configured to inject fuel into nying drawings wherein like reference numerals refer to like combustor 24. Combustor 24 is configured to combust the parts throughout the several views, and wherein: fuel injected therein by fuel injectors 26 with pressurized air FIG. 1 schematically depicts some aspects of a non- received from compressor system 12. Turbine system 16 limiting example of a gas turbine engine in accordance with 55 includes a turbine case 28 that houses stationary and rotating an embodiment of the present invention. turbine system 16 components. In various embodiments, FIG. 2 schematically illustrates some aspects of non- turbine case 28 may be formed of one or more individual limiting examples of a heat exchanger and a gas turbine turbine case structures, e.g., depending on the number, size engine in accordance with an embodiment of the present and location of turbine stages and/or the number of spools invention. 60 employed in engine 10.

FIG. 3 schematically illustrates some aspects of non- Engine 10 includes a heat exchanger 30 fluidly disposed limiting examples of a heat exchanger and a compressor between two compressor stages and in fluid communication system in accordance with an embodiment of the present with fuel injectors 26. Heat exchanger 30 is configured to invention. cool pressurized airflow in compressor system 12 by heat FIG. 4 schematically illustrates some aspects of a non- 65 exchange with the fuel supplied to fuel injectors 26, and to limiting example of a heat exchanger in accordance with an heat the fuel by heat exchange with the pressurized airflow embodiment of the present invention. in compressor system 12.

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Referring to FIG. 2, in conjunction with FIG. 1, some surized air flow from compressor stage 32 to heat exchanger aspects of non-limiting examples of heat exchanger 30 and 30 and then to compressor stage 34. Primary flowpath 38 is engine 10 in accordance with an embodiment of the present disposed within engine 10, that is, disposed within one or invention are schematically depicted. Heat exchanger 30 is more engine 10 cases, as opposed to being disposed external an air/fuel heat exchanger that is configured to exchange 5 to gas turbine engine 10. By being disposed internal to heat between compressor system 12 air and the fuel supplied engine 10, primary flowpath 38 does not include any engine to fuel injectors 26 without the use of an intermediate heat 10 external ducting to duct pressurized air to and from heat transfer fluid. Heat exchanger 30 is disposed within engine exchanger 30 from and to compressor stages 32 and 34 10. That is, disposed within one or more engine 10 cases, as within engine 10. By being internal to engine 10, primary opposed to being disposed external to gas turbine engine 10, io flowpath 38 does not increase the frontal area of engine 10, which requires external ducting to duct the pressurized air to which would otherwise adversely impact the flight charac- and from heat exchanger 30 from and to compressor stages teristics or frontal area drag ofthe aircraft or air-vehicle into within engine 10. By being internal to engine 10, heat which engine 10 is installed as a propulsion power plant.

exchanger 30 does not increase the frontal area ofengine 10, In one form, primary flowpath 38 has a maximum radial which would otherwise adversely impact the flight charac- 15 extent 40, relative to engine centerline 18, and heat teristics and frontal area drag of the aircraft or air-vehicle exchanger 30 has a maximum radial extent 42, relative to into which engine 10 is installed as a propulsion power engine centerline 18,that do not exceed the maximum radial plant. In one form, heat exchanger 30 is an annular heat extent 44 relative to engine centerline 18, (FIG. 1), of exchanger, extending annularly around engine centerline 18. turbine case 28. In one form, maximum radial extent 40 of In other embodiments, heat exchanger 30 may take other 20 primary flowpath 38 and/or maximum radial extent 42 of forms. heat exchanger 30 do not exceed the maximum radial extent Heat exchanger 30 is in fluid communication with and 46, relative to engine centerline 18 (FIG. 1), of a turbine fluidly disposed between a compressor stage 32 and a blade tip diameter of turbine system 16. In one form, compressor stage 34. Combustor 24 is fluidly disposed maximum radial extent 40 of primary flowpath 38 and downstream of compressor stage 34. Compressor stage 32 is 25 maximum radial extent 42 of heat exchanger 30 do not a lower pressure compressor stage than compressor stage 34. exceed the maximum radial extent 48, relative to engine Compressor stage 32 is configured to produce a pressurized centerline 18 (FIG. 1), of combustor case 22. In one form, airflow, which is received by compressor stage 34 after maximum radial extent 40 of primary flowpath 38 and having passed through heat exchanger 30. In one form, maximum radial extent 42 of heat exchanger 30 do not compressor stage 34 is a final compressor stage, and com- 30 exceed the maximum radial extent 50, relative to engine bustor 24 is configured to receive compressor discharge air centerline 18 (FIG. 1), of compressor case 20, e.g., a high from compressor stage 34 for combustion, e.g., via a dif- pressure (HP) compressor case 52, which surrounds and fuser. In other embodiments, compressor stage 34 may not houses compressor stage 34. In other embodiments, maxi- be a final compressor stage. mum radial extent 40 of primary flowpath 38 and maximum Heat exchanger 30 is also in fluid communication with a 35 radial extent 42 of heat exchanger 30 may be disposed fuel supply 36 and fuel injectors 26. Fuel supply 36 is within the radial extents of other engine 10 components.

operative to supply fuel to heat exchanger 30 for subsequent Primary flowpath 38 includes a diffuser portion 54 and a delivery to fuel injectors 26 after having performed heat converging portion 56. Diffuser portion 54 is fluidly dis- exchange between pressurized air from compressor stage 32 posed upstream of heat exchanger 30. Diffuser portion 54 is and the fuel prior to delivery of the fuel to fuel injectors 26. 40 configured to diffuse the air pressurized by compressor stage Heat exchanger 30 is configured to receive the pressurized 32 prior to entry of the pressurized air into heat exchanger air flow from compressor stage 32, to discharge the pres- 30. Converging portion 56 is fluidly disposed downstream of surized air flow to compressor stage 34; to heat the fuel by heat exchanger 30. Converging portion 56 is configured to heat exchange with the pressurized air flow prior to delivery reduce the flow area in primary flowpath 38 and to increase ofthe fuel to fuel injector 26; and to cool the pressurized air 45 the velocity of the air pressurized by compressor stage 32 flow by heat exchange with the fuel prior to delivery of the after the pressurized air has passed through heat exchanger pressurized air flow to compressor stage 34. 30, prior to delivery of the pressurized air to compressor Referring to FIG. 3, some aspects of non-limiting stage 34.

examples of heat exchanger 30 and compressor system 12 in In one form, disposed within diffuser portion 54 is a flow accordance with an embodiment of the present invention are 50 splitter 58. Some embodiments may not include a flow schematically depicted. In one form, compressor stage 32 is splitter. Flow splitter 58 is configured to prevent or reduce an axial compressor stage, whereas compressor stage 34 is separation of the pressurized air flow from the walls of a centrifugal compressor stage. In other embodiments, com- diffuser portion 54 upstream of heat exchanger 30. In one pressor stages 32 and 34 take other forms, e.g., including form,flow splitter 58 is configured to enable a more aggres- both compressor stages 32 and 34 being axial compressor 55 sive diffusion angle in diffuser portion 54 than the diffusion stages; both compressor stages 32 and 34 being centrifugal angle of a diffuser portion not having a flow splitter, e.g., compressor stages; or compressor stage 32 being a centrifu- which allows a reduction in the length of diffuser portion 54 gal compressor stage and compressor stage 34 being an axial relative to embodiments not equipped with flow splitter 58.

compressor stage. In one form,flow splitter 58 is positioned proximate to heat Disposed between and fluidly coupling compressor stage 60 exchanger 30, e.g., immediately adjacent to heat exchanger 32 and compressor stage 34 is a primary flowpath 38. In one 30, to prevent recirculation of the pressurized air down- form, primary flowpath 38 is annular, extending annularly stream offlow splitter 58 (between flow splitter 58 and heat around engine centerline 18 and forming an annulus therein. exchanger 30), e.g., owing to potential pressure differentials In other embodiments, primary flowpath 38 may take other between locations above and below splitter 58, e.g., which forms. Heat exchanger 30 is disposed within primary flow- 65 may otherwise yield an effective flow blockage. In some path 38, between compressor stage 32 and compressor stage embodiments, a seal 60 is disposed between flow splitter 58 34. Primary flowpath 38 is configured to deliver the pres- and heat exchanger 30 in order to further prevent recircu-

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lation downstream offlow splitter 58. Seal 60 may take any pre-split in a manner that does not adversely affect the latent form suitable for fitment and sealing between flow splitter heating value of the fuel. In one form, endothermic fuel has 58 and heat exchanger 30.

a temperature limit of approximately 900° F. As the allow- Referring to FIG. 4 in conjunction with FIG. 3, some able fuel temperature increases, more heat can be transferred aspects of a non-limiting example of heat exchanger 30 in 5 to the fuel from the pressurized air flow provided by accordance with an embodiment of the present invention are compressor stage 32 via heat exchanger 30, which increases schematically depicted. For clarity of illustration, only heat the specific fuel consumption (SFC) benefit to engine 10 exchanger 30 is illustrated in FIG. 4. As set forth previously, from the use of heat exchanger 30, relative to the use of in one form, heat exchanger 30 is an annular heat exchanger.

lower temperature-capable fuels. Additionally, by using a In a particular form, heat exchanger 30 is formed of a to high temperature capable fuel, purging of heat exchanger 30 plurality of individual heat exchanger modules 62. In one after engine 10 shutdown may not be required to avoid fuel form, heat exchanger modules 62 are equally spaced apart coking at high temperatures, e.g., high operating tempera- circumferentially and arranged annularly within primary tures and hot soak-back conditions. Hence, such embodi- flowpath 38, between diffuser portion 54 and converging 15 ments may not require a purge system,reducing the cost and portion 56. In other embodiments, heat exchanger modules weight of engine 10 relative to systems that do require a 62 may be arranged differently.

purge system.

In one form, heat exchanger 30 includes sixteen heat In other embodiments, the fuel used by engine 10 is a exchanger modules 62. In other embodiments, the number deox fuel, and combustor 24, fuel injectors 26 and air/fuel and size of heat exchanger modules 62 may vary with the 20 heat exchanger 30 are configured for use with the deox fuel.

needs of the particular application. In one form, heat exchanger modules 62 are plate-and-fin heat exchanger Deox fuel is a fuel that has been processed to remove oxygen modules. In other embodiments, heat exchanger modules 62 from the fuel. In one form, deox fuel has a temperature limit may take other forms. In one form, heat exchanger modules of approximately 600° F. As the allowable fuel temperature 62 are configured for cross-flow heat exchange between the increases, more heat can be transferred to the fuel from the fuel and the air pressurized by compressor stage 32. In other 25 pressurized air flow provided by compressor stage 32 via embodiments, other heat exchange configurations may be heat exchanger 30, which increases the SFC benefit to employed in place of or in addition to cross-flow, e.g., engine 10 from the use of heat exchanger 30, relative to the counter-flow, parallel flow and/or mixed flow. use of lower temperature-capable fuels. Additionally, by Disposed between heat exchanger modules 62 are fuel using a higher temperature-capable fuel, purging of heat distribution manifolds 64. Fuel distribution manifolds 64 are so exchanger 30 after engine 10 shutdown may not be required in fluid communication with adjacent heat exchanger mod- to avoid fuel coking at high temperatures, e.g., high oper- ules 62, and are configured to transmit fuel between the ating temperatures and hot soak-back conditions. Hence, adjacent heat exchanger modules 62. In one form, fuel such embodiments may not require a purge system,reducing distribution manifolds are pie-shaped, owing to the shape of the cost and weight of engine 10 relative to systems that do heat exchanger modules 62. In other embodiments, other 35 require a purge system.

suitable shapes may be employed. In one form, heat In still other embodiments, the fuel used by engine 10 is exchanger 30 is effectively split into two parallel heat a conventional gas turbine engine fuel, e.g., JP-8, and exchanger halves with a fuel inlet 66 and a fuel outlet 68 for combustor 24, fuel injectors 26 and air/fuel heat exchanger distributing fuel in a generally circumferential direction 70 30 are configured for use with the conventional fuel. In one through one side of heat exchanger 30; and with a fuel inlet 40 form, conventional fuel has a temperature limit of approxi- 72 and a fuel outlet 74 for distributing fuel in a generally mately 450° F.

circumferential direction 76 through the other side of heat Embodiments of the present invention include an aircraft exchanger 30. By effectively splitting heat exchanger 30 into propulsion gas turbine engine, comprising: a first compres- two parallel heat exchangers,the circumferential variation in sor stage configured to produce a pressurized air flow; a heat transfer to the pressurized air flow provided by com- 45 second compressor stage disposed downstream of the first pressor stage 32 is reduced. In other embodiments, only a compressor stage; a primary annular flowpath fluidly cou- single fuel inlet and a single fuel outlet may be employed, pling the first compressor stage and the second compressor e.g., for distributing the fuel around the entire heat stage, wherein the primary annular flowpath is disposed exchanger 30. In still other embodiments, a plurality offuel within the aircraft propulsion gas turbine engine; a combus- inlets and/or fuel outlets may be employed to distribute fuel 50 for disposed downstream of the second compressor stage; a in parallel through smaller segments of heat exchanger 30, fuel injector configured to inject a fuel into the combustor, e.g., to further reduce the circumferential variation in heat wherein the combustor is configured to combust the fuel transferred to the pressurized air flow provided by compres- injected therein by the fuel injector; and an air/fuel heat sor stage 32. exchanger disposed in the primary annular flowpath, In one form, disposed immediately upstream of each fuel 55 wherein the air/fuel heat exchanger is in fluid communica- distribution manifold 64 is a leading transition 78. Leading tion with the fuel injector, the first compressor stage and the transitions 78 are configured to guide the pressurized airflow second compressor stage; and wherein the air/fuel heat around fuel distribution manifolds 64 and into heat exchanger is configured to receive the pressurized air flow exchanger modules 62, which reduces pressure losses in the from the first compressor stage, to discharge the pressurized pressurized air flow from compressor stage 32. In some 6o air flow to the second compressor stage, to heat the fuel by embodiments, trailing transitions may also be positioned heat exchange with the pressurized air flow prior to delivery downstream of fuel distribution manifolds 64 to reduce ofthe fuel to the fuel injector, and to cool the pressurized air pressure losses in air flow exiting heat exchanger 30. flow by heat exchange with the fuel.

In one form, the fuel used by engine 10 is an endothermic In a refinement, the air/fuel heat exchanger is an annular fuel, and combustor 24, fuel injectors 26 and air/fuel heat 65 heat exchanger.

exchanger 30 are configured for use with the endothermic In another refinement, the annular heat exchanger fuel. Endothermic fuel is a fuel having the fuel molecules includes a plurality of individual heat exchanger modules

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arranged annularly within the primary annular flowpath to Embodiments of the present invention include a gas form the annular heat exchanger. turbine engine, comprising: a first compressor stage config- ured to produce a pressurized air flow; a second compressor In yet another refinement, the air/fuel heat exchanger stage disposed downstream of the first compressor stage; a includes a plate-and-fin heat exchanger.

5 combustor disposed downstream of the second compressor In still another refinement, the fuel is a deox fuel; and the stage; a fuel injector configured to inject a fuel into the combustor, the fuel injector and the air/fuel heat exchanger combustor, wherein the combustor is configured to combust are configured for use with the deox fuel.

the fuel injected therein by the fuel injector; and means for In yet still another refinement, the fuel is an endothermic cooling the pressurized air flow prior to delivery of the fuel; and the combustor,the fuel injector and the air/fuel heat io pressurized air flow to the second compressor stage and for exchanger are configured for use with the endothermic fuel.

heating the fuel prior to delivery of the fuel to the fuel In a further refinement, the aircraft propulsion gas turbine injector.

engine further comprises an engine case, wherein a maxi- In a refinement, the gas turbine engine further comprising mum radial extent of the primary annular flowpath is less an engine case, wherein a maximum radial extent of the than a maximum radial extent of the engine case.

15 means for cooling and for heating is less than a maximum In a yet further refinement, the engine case is one of a radial extent of the engine case; and wherein the gas turbine compressor case, a combustor case and a turbine case.

engine is configured as an aircraft propulsion gas turbine In a still further refinement, the engine case is an HP engine.

compressor case.

While the invention has been described in connection Embodiments of the present invention include a gas 20 with what is presently considered to be the most practical turbine engine, comprising: a first compressor stage config- and preferred embodiment, it is to be understood that the ured to produce a pressurized air flow; a second compressor invention is not to be limited to the disclosed embodiment stage disposed downstream of the first compressor stage; a (s), but on the contrary, is intended to cover various modi- combustor disposed downstream of the second compressor fications and equivalent arrangements included within the stage; a fuel injector configured to inject a fuel into the 25 spirit and scope of the appended claims, which scope is to combustor, wherein the combustor is configured to combust be accorded the broadest interpretation so as to encompass the fuel injected therein by the fuel injector; and an air/fuel all such modifications and equivalent structures as permitted heat exchanger fluidly disposed between the first compressor under the law. Furthermore it should be understood that stage and the second compressor stage, wherein the air/fuel while the use ofthe word preferable, preferably, or preferred heat exchanger is in fluid communication with the fuel 30 in the description above indicates that feature so described injector, the first compressor stage and the second compres- may be more desirable, it nonetheless may not be necessary sor stage; and wherein the air/fuel heat exchanger is con- and any embodiment lacking the same may be contemplated figured to receive the pressurized air flow from the first as within the scope ofthe invention, that scope being defined compressor stage, to discharge the pressurized air flow to the by the claims that follow. In reading the claims it is intended second compressor stage, to heat the fuel by heat exchange 35 that when words such as "a," "an," "at least one" and "at with the pressurized air flow prior to delivery of the fuel to least a portion" are used, there is no intention to limit the the fuel injector, and to cool the pressurized air flow by heat claim to only one item unless specifically stated to the exchange with the fuel prior to delivery ofthe pressurized air contrary in the claim. Further, when the language "at least a flow to the second compressor stage, wherein air/fuel heat portion" and/or "a portion" is used the item may include a exchanger is disposed within the gas turbine engine. 40 portion and/or the entire item unless specifically stated to the In a refinement, the gas turbine engine further comprises contrary.

a primary annular flowpath fluidly coupling the first com- What is claimed is: pressor stage and the second compressor stage, wherein the 1. An aircraft propulsion gas turbine engine, comprising: air/fuel heat exchanger is disposed within the primary annu- a first compressor stage configured to produce a pressur- lar flowpath. 45 ized air flow; In another refinement, the primary annular flowpath a second compressor stage disposed downstream of the includes a diffuser portion upstream of the air/fuel heat first compressor stage; exchanger and a converging portion downstream of the a primary annular flowpath fluidly coupling the first air/fuel heat exchanger. compressor stage and the second compressor stage, In yet another refinement, the gas turbine engine further 50 wherein the primary annular flowpath is disposed comprises a flow splitter disposed in the diffuser portion within the aircraft propulsion gas turbine engine; proximate to the air/fuel heat exchanger, wherein the flow a combustor disposed downstream of the second com- splitter is configured to prevent or reduce flow separation in pressor stage; the diffuser portion. a fuel injector configured to inject a fuel into the com- In still another refinement, the gas turbine engine further 55 bustor, wherein the combustor is configured to combust comprises a seal disposed between the flow splitter and the the fuel injected therein by the fuel injector; and air/fuel heat exchanger. an air/fuel heat exchanger disposed in the primary annular In yet another refinement, the gas turbine engine further flowpath, the air/fuel heat exchanger including a first comprises an engine case, wherein a maximum radial extent flowpath and a second flowpath, the first flowpath of the air/fuel heat exchanger is less than a maximum radial 60 having a first fuel inlet and a first fuel outlet for extent of the engine case. distributing fuel in a first direction through a first side In yet still another refinement, the engine case is one of a of the air/fuel heat exchanger, and the second flowpath compressor case, a combustor case and a turbine case. having a second fuel inlet and a second fuel outlet for In a further refinement, the engine case is an HP com- distributing fuel in a second direction through a second pressor case. 65 side of the air/fuel heat exchanger; In a yet further refinement, the gas turbine engine is wherein the air/fuel heat exchanger is an annular heat configured as an aircraft propulsion gas turbine engine. exchanger with the first direction being a first circum-

US 9,771,867 B2

9 10

ferential direction and the second direction being a discharge the pressurized air flow to the second com- second circumferential direction; pressor stage, to heat the fuel by heat exchange with the wherein the air/fuel heat exchanger is in fluid communi- pressurized air flow prior to delivery of the fuel to the cation with the fuel injector, the first compressor stage fuel injector, and to cool the pressurized air flow by and the second compressor stage; and 5 heat exchange with the fuel prior to delivery of the wherein the air/fuel heat exchanger is configured to pressurized air flow to the second compressor stage, receive the pressurized air flow from the first compres- wherein the air/fuel heat exchanger is disposed within sor stage, to discharge the pressurized air flow to the the gas turbine engine.

second compressor stage, to heat the fuel by heat 10. The gas turbine engine of claim 9, further comprising exchange with the pressurized air flow prior to delivery 10 a primary annular flowpath fluidly coupling the first com- of the fuel to the fuel injector, and to cool the pressur- pressor stage and the second compressor stage, wherein the ized air flow by heat exchange with the fuel.

air/fuel heat exchanger is disposed within the primary annu- 2. The aircraft propulsion gas turbine engine of claim 1, lar flowpath.

wherein the annular heat exchanger includes a plurality of 11. The gas turbine engine of claim 10, wherein the individual heat exchanger modules, the plurality of indi- 15 primary annular flowpath includes a diffuser portion vidual heat exchanger modules being spaced apart circum- upstream of the air/fuel heat exchanger and a converging ferentially and arranged annularly within the primary annu- portion downstream of the air/fuel heat exchanger.

lar flowpath to form the annular heat exchanger.

12. The gas turbine engine ofclaim 11,further comprising 3. The aircraft propulsion gas turbine engine of claim 1, 20 a flow splitter disposed in the diffuser portion proximate to wherein the air/fuel heat exchanger includes a plate-and-fin the air/fuel heat exchanger, wherein the flow splitter is heat exchanger.

configured to prevent or reduce flow separation in the 4. The aircraft propulsion gas turbine engine of claim 1, diffuser portion.

wherein the fuel is a deox fuel; and wherein the combustor, 13. The gas turbine engine ofclaim 12,further comprising the fuel injector and the air/fuel heat exchanger are config- a seal disposed between the flow splitter and the air/fuel heat ured for use with the deox fuel. 25 exchanger.

5. The aircraft propulsion gas turbine engine of claim 1, 14. The gas turbine engine of claim 9, further comprising wherein the fuel is an endothermic fuel; and wherein the an engine case, wherein the air/fuel heat exchanger is combustor, the fuel injector and the air/fuel heat exchanger narrower than the engine case.

are configured for use with the endothermic fuel.

15. The gas turbine engine of claim 14, wherein the 6. The aircraft propulsion gas turbine engine of claim 1, 30 engine case is one of a compressor case, a combustor case further comprising an engine case, wherein the primary and a turbine case.

annular flowpath is narrower than the engine case.

16. The gas turbine engine of claim 14, wherein the 7. The aircraft propulsion gas turbine engine of claim 6, engine case is a HP compressor case.

wherein the engine case is one of a compressor case, a 35 17. The gas turbine engine of claim 9, configured as an combustor case and a turbine case.

aircraft propulsion gas turbine engine.

8. The aircraft propulsion gas turbine engine of claim 6, 18. A gas turbine engine, comprising: wherein the engine case is a HP compressor case.

a first compressor stage configured to produce a pressur- 9. A gas turbine engine, comprising: ized air flow; a first compressor stage configured to produce a pressur- a second compressor stage disposed downstream of the ized air flow; 40 first compressor stage; a second compressor stage disposed downstream of the a combustor disposed downstream of the second com- first compressor stage; pressor stage; a combustor disposed downstream of the second com- a fuel injector configured to inject a fuel into the com- pressor stage; 45 bustor, wherein the combustor is configured to combust a fuel injector configured to inject a fuel into the com- the fuel injected therein by the fuel injector; and bustor, wherein the combustor is configured to combust a heat exchanger including a first flowpath and a second the fuel injected therein by the fuel injector; and flowpath, the first flowpath having a first fuel inlet and an air/fuel heat exchanger fluidly disposed between the a first fuel outlet for distributing fuel in a first direction first compressor stage and the second compressor stage, through a first side of the heat exchanger, and the the air/fuel heat exchanger including a first flowpath 50 second flowpath having a second fuel inlet and a and a second flowpath, the first flowpath having a first second fuel outlet for distributing fuel in a second fuel inlet and a first fuel outlet for distributing fuel in direction through a second side of the heat exchanger, a first direction through a first side of the air/fuel heat wherein the air/fuel heat exchanger is an annular heat exchanger, and the second flowpath having a second exchanger with the first direction being a first circum- fuel inlet and a second fuel outlet for distributing fuel 55 ferential direction and the second direction being a in a second direction through a second side of the second circumferential direction, and air/fuel heat exchanger, wherein the heat exchanger is configured to cool the wherein the air/fuel heat exchanger is an annular heat pressurized air flow prior to delivery of the pressurized exchanger with the first direction being a first circum- 60 air flow to the second compressor stage and to heat the ferential direction and the second direction being a fuel prior to delivery of the fuel to the fuel injector.

second circumferential direction, and 19. The gas turbine engine ofclaim 18,further comprising wherein the air/fuel heat exchanger is in fluid communi- an engine case, wherein the heat exchanger is narrower than cation with the fuel injector, the first compressor stage the engine case, and wherein the gas turbine engine is and the second compressor stage; and wherein the configured for aircraft propulsion.

air/fuel heat exchanger is configured to receive the 65 pressurized air flow from the first compressor stage, to