United States Patent: 10731563

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United States Patent	10,731,563
Munsell , et al.	August 4, 2020

Compressed air bleed supply for buffer system

Abstract

A method of designing a buffer system for a gas turbine engine according to an example of the present disclosure includes, among other things, configuring a heat exchanger to define a first inlet and a first outlet fluidly coupled to each other and a second inlet and a second outlet fluidly coupled to each other, configuring the first inlet to receive air from first and second air sources that are selectively fluidly coupled to the first inlet, configuring the second inlet to receive air from a third air source fluidly that is coupled to the second inlet, and configuring the first outlet to provide cooled pressurized air to multiple fluid-supplied areas that are located remotely from one another and that are fluidly coupled to the first outlet, the multiple fluid-supplied areas including a bearing compartment of a gas turbine engine.

Inventors:

Munsell; Peter M. (Granby, CT), Stripinis; Philip S. (Rocky Hill, CT)

Applicant:

Name	City	State	Country	Type
United Technologies Corporation	Farmington	CT	US

Assignee:

RAYTHEON TECHNOLOGIES CORPORATION (Farmington, CT)

Family ID:

48869054

Appl. No.:

15/882,058

Filed:

January 29, 2018

Prior Publication Data


	Document Identifier	Publication Date
	US 20180149088 A1	May 31, 2018

Related U.S. Patent Documents


Application Number	Filing Date	Patent Number	Issue Date
14246260	Apr 7, 2014	9879602
13362499	Jan 31, 2012

Current U.S. Class: 1/1

Current CPC Class: F02C 7/185 (20130101); F02C 6/08 (20130101); F01D 25/183 (20130101); F02C 7/18 (20130101); F01D 25/22 (20130101); Y10T 29/49359 (20150115); Y02T 50/60 (20130101)

Current International Class: F02C 1/00 (20060101); F01D 25/18 (20060101); F02C 7/18 (20060101); F01D 25/22 (20060101); F02C 6/08 (20060101)

References Cited [Referenced By]

U.S. Patent Documents


2258792	April 1941	New
2936655	May 1960	Peterson et al.
3021731	February 1962	Stoeckicht
3194487	July 1965	Tyler et al.
3273340	September 1966	Hull, Jr.
3287906	November 1966	McCormick
3352178	November 1967	Lindgren et al.
3412560	November 1968	Gaubatz
3487993	January 1970	Rannenberg
3527054	September 1970	Hemsworth
3747343	July 1973	Rosen
3754484	August 1973	Roberts
3820719	June 1974	Clark
3842597	October 1974	Ehrich
3892358	July 1975	Gisslen
3932058	January 1976	Harner et al.
3935558	January 1976	Miller et al.
3988889	November 1976	Chamay et al.
4130872	December 1978	Harloff
4284174	August 1981	Salvana et al.
*4478551*	October 1984	Honeycutt, Jr. et al.
4542623	September 1985	Hovan et al.
4550561	November 1985	Coffinberry
4574584	March 1986	Hovan
4645415	February 1987	Hovan et al.
4649114	March 1987	Miltenburger et al.
4696156	September 1987	Burr et al.
4709880	December 1987	Bradfield et al.
4979362	December 1990	Vershure, Jr.
5102379	April 1992	Pagluica et al.
5141400	August 1992	Murphy et al.
5305616	April 1994	Coffinberry
5317877	June 1994	Stuart
5414992	May 1995	Glickstein
5433674	July 1995	Sheridan et al.
5447411	September 1995	Curley et al.
5466198	November 1995	McKibbin et al.
5524847	June 1996	Brodell et al.
5581996	December 1996	Koch et al.
5677060	October 1997	Terentieva et al.
5778659	July 1998	Duesler et al.
5857836	January 1999	Stickler et al.
5915917	June 1999	Eveker et al.
5975841	November 1999	Lindemuth et al.
5985470	November 1999	Spitsberg et al.
6035627	March 2000	Liu
6223616	May 2001	Sheridan
6305156	October 2001	Lui
6315815	November 2001	Spadaccini et al.
6318070	November 2001	Rey et al.
6387456	May 2002	Eaton, Jr. et al.
6412270	July 2002	Mortzheim et al.
6517341	February 2003	Brun et al.
6607165	August 2003	Manteiga et al.
6709492	March 2004	Spadaccini et al.
6732502	May 2004	Seda et al.
6814541	November 2004	Evans et al.
7021042	April 2006	Law
7328580	February 2008	Lee et al.
7334392	February 2008	Moniz et al.
7374403	May 2008	Decker et al.
7591631	September 2009	Hendricks et al.
7591754	September 2009	Duong et al.
7662059	February 2010	McCune
7694505	April 2010	Schilling
7806651	October 2010	Kennepohl et al.
7824305	November 2010	Duong et al.
7926260	April 2011	Sheridan et al.
7997868	August 2011	Liang et al.
8074440	December 2011	Kohlenberg et al.
8205432	June 2012	Sheridan
2006/0213202	September 2006	Fukutani
2008/0003096	January 2008	Kohli et al.
2008/0116009	May 2008	Sheridan et al.
2008/0310956	December 2008	Jain
2008/0317588	December 2008	Grabowski et al.
2009/0056343	March 2009	Suciu et al.
2009/0293445	December 2009	Ress, Jr.
2010/0092116	April 2010	Franconi
2010/0105516	April 2010	Sheridan et al.
2010/0148396	June 2010	Xie et al.
2010/0192593	August 2010	Brown
2010/0212281	August 2010	Sheridan
2010/0218483	September 2010	Smith
2010/0331139	December 2010	McCune
2011/0159797	June 2011	Beltman et al.
2011/0203293	August 2011	Giahn
2011/0293423	December 2011	Bunker et al.
2012/0124964	May 2012	Hasel et al.
2013/0133334	May 2013	Strecker et al.

Foreign Patent Documents


0127562	May 1984	EP
0791383	Aug 1997	EP
1142850	Oct 2001	EP
1516041	Jun 1978	GB
2041090	Sep 1980	GB
2152148	Jul 1985	GB
2426792	Dec 2006	GB
2007038674	Apr 2007	WO

Other References

Honeywell LF507. Jane's Aero-engines, Aero-engines--Turbofan. Feb. 9, 2012. cited by applicant .
Honeywell TFE731. Jane's Aero-engines, Aero-engines--Turbofan. Jul. 18, 2012. cited by applicant .
NASA Conference Publication. Quiet, powered-lift propulsion. Cleveland, Ohio. Nov. 14-15, 1978. pp. 1-420. cited by applicant .
"Civil Turbojet/Turbofan Specifications", Jet Engine Specification Database (Apr. 3, 2005). cited by applicant .
Kandebo, S.W. (1993). Geared-turbofan engine design targets cost, complexity. Aviation Week & Space Technology, 148(8). Start p. 32. cited by applicant .
Hendricks, E.S. and Tong, M.T. (2012). Performance and weight estimates for an advanced open rotor engine. NASA/TM-2012-217710. pp. 1-13. cited by applicant .
Guynn, M. D., Berton, J.J., Fisher, K. L, Haller, W.J., Tong, M. T., and Thurman, D.R. (2011). Refined exploration of turbofan design options for an advanced single-aisle transport. NASA/TM-2011-216883. pp. 1-27. cited by applicant .
Zalud, T. (1998). Gears put a new spin on turbofan performance. Machine Design, 70(20), p. 104. cited by applicant .
Kurzke, J. (2008). Preliminary Design, Aero-engine design: From state of the art turbofans towards innovative architectures. pp. 1-72. cited by applicant .
Zamboni, G. and Xu, L. (2009). Fan root aerodynamics for large bypass gas turbine engines: Influence on the engine performance and 3D design. Proceedings of ASME Turbo Expo 2009: Power for Land, Sea and Air. Jun. 3-12, 2009, Orlando, Florida, USA. pp. 1-12. cited by applicant .
Han, J., Dutta, S., and Ekkad, S.V. (2000). Gas turbine heat transfer and cooling technology. New York, NY: Taylor & Francis. pp. 1-25, 129-157, and 160-249. cited by applicant .
Mattingly, J.D. (1996). Elements of gas turbine propulsion. New York, New York: McGraw-Hill, Inc. pp. 1-18, 60-62, 85-87, 95-104, 121-123, 223-234, 242-245, 278-280, 303-309, 323-326, 462-479, 517-520, 563-565, 673-675, 582-685, 697-699, 703-705, 802-805, 862-864, and 923-925. cited by applicant .
Declaration of Reza Abhari, Ph.D. In re U.S. Pat. No. 8,844,265. Executed Jun. 28, 2016. pp. 1-91. cited by applicant .
Declaration of John Eaton, Ph.D. In re U.S. Pat. No. 8,869,568. Executed Mar. 28, 2016. pp. 1-87. cited by applicant .
Declaration of Reza Abhari. In re U.S. Pat. No. 8,695,920. Executed Nov. 30. pp. 1-67. cited by applicant .
Declaration of Reza Abhari. In re U.S. Pat. No. 8,448,895. Executed Nov. 28. pp. 1-81. cited by applicant .
Declaration of Reza Abhari. In re U.S. Pat. No. 8,695,920, claims 1-4, 7-14, 17 and 19. Executed Nov. 29. pp. 1-102. cited by applicant .
Declaration of Dr. Magdy Attia. In re U.S. Pat. No. 8,313,280. Executed Oct. 21, 2016. pp. 1-88. cited by applicant .
Lord, W.K., MacMartin, D.G., and Tillman, T.G. (2000). Flow control opportunities in gas turbine engines. American Institute of Aeronautics and Astronautics. pp. 1-15. cited by applicant .
Daly, M. Ed. (2010). Jane's Aero-Engine. Issue Twenty-seven. Mar. 2010. p. 633-636. cited by applicant .
Roux, E. (2007). Turbofan and turbojet engines database handbook. Editions Elodie Roux. Blagnac: France. pp. 1-595. cited by applicant .
Wilfert, G. (2008). Geared fan. Aero-Engine Design: From State of the Art Turbofans Towards Innovative Architectures, von Karman Institute for Fluid Dynamics, Belgium, Mar. 3-7, 2008. pp. 1-26. cited by applicant .
Declaration of Dr. Magdy Attia. In re U.S. Pat. No. 8,517,668. Executed Dec. 8, 2016. pp. 1-81. cited by applicant .
Cramoisi, G. Ed. (2012). Death in the Potomac: The crash of Air Florida Flight 90. Air Crash Investigations. Accident Report NTSB/AAR-82-8. p. 45-47. cited by applicant .
Norton, M. and Karczub, D. (2003). Fundamentals of noise and vibration analysis for engineers. Press Syndicate of the University of Cambridge. New York: New York. p. 524. cited by applicant .
U.S. Department of Transportation: Federal Aviation Administration Advisory Circular. Runway overrun prevention. dated Nov. 6, 2007. p. 1-8 and Appendix 1 p. 1-15, Appendix 2 p. 1-6, Appendix 3 p. 1-3, and Appendix 4 p. 1-5. cited by applicant .
U.S. Department of Transportation: Federal Aviation Administration Advisory Circular. Standard operating procedures for flight deck crewmembers. dated Feb. 27, 2003.. p. 1-6 and Appendices. cited by applicant .
Vasudevan, A.K. and Petrovic, J.J. (1992). A comparative overview of molybedenum disilicide composites. Materials Science and Engineering, A155, 1992. pp. 1-17. cited by applicant .
Clarke, D.R. and Levi, C.G. (2003). Materials design for the next generation thermal barrier coatings. Annual. Rev. Mater. Res. vol. 33. 2003. pp. 383-417. cited by applicant .
Lee, K.N. (2000). Current status of environmental barrier coatings for Si-Based ceramics. Surface and Coatings Technology 133-134, 2000. pp. 1-7. cited by applicant .
Bornstein, N. (1993). Oxidation of advanced intermetallic compounds. Journal de Physique IV, 1993, 03 (C9), pp. C9-367-C9-373. cited by applicant .
Krenkel, W., Naslain, R., and Schneider, H. Eds. (2001). High temperature ceramic matrix composites pp. 224-229. Weinheim, DE: Wiley-VCH Verlag GmbH. cited by applicant .
Gibala, R., Ghosh, A.K., Van Aken, D.C., Srolovitz, D.J., Basu, A., Chang, H., . . . Yang, W. (1992). Mechanical behavior and interface design of MoSi2-based alloys and composites. Materials Science and Engineering, A155, 1992. pp. 147-158. cited by applicant .
Shah, D.M. (1992). MoSi2 and other silicides as high temperature structural materials. Superalloys 1992. The Minerals, Metals, & Materials Society. pp. 409-422. cited by applicant .
Zhao, J.C. and Westbrook, J.H. (2003). Ultrahigh-temperature materials for jet engines. MRS Bulletin. vol. 28(9). Sep. 2003. pp. 622-630. cited by applicant .
Tsirlin, M., Pronin, Y.E., Florina, E.K., Mukhametov, S. Kh., Khatsernov, M.A., Yun, H.M., . . . Kroke, E. (2001). Experimental investigation of multifunctional interphase coatings on SiC fibers for non-oxide high temperature resistant CMCs. High Temperature Ceramic Matrix Composites. 4th Int'l Conf. on High Temp. Ceramic Matrix Composites. Oct. 1-3, 2001. pp. 149-156. cited by applicant .
Jacobson, N.S. (1993). Corrosion of silicon-based ceramics in combustion environments. J. Am. Ceram. Soc. 76(1). pp. 3-28. cited by applicant .
Jorgensen, P.J., Wadsworth, M.E., and Cutler, I.B. (1961). Effects of water vapor on oxidation of silicon carbide. J. Am. Ceram. Soc. 44(6). pp. 248-261. cited by applicant .
Xu, Y., Cheng, L., Zhang, L., Ying, H., and Zhou, W. (1999). Oxidation behavior and mechanical properties of C/3iC composites with Si-MoSi2 oxidation protection coating. J. of Mat. Sci. vol. 34. 1999. pp. 6009-6014. cited by applicant .
Sundaram, S.K., Hsu, J-Y., Speyer, R.F. (1995). Molten glass corrosion resistance of immersed combustion-heating tube materials in e-glass. J. Am. Ceram. Soc. 78(7). pp. 1940-1946. cited by applicant .
Jeng, Y.-L., Lavernia, E.J. (1994). Processing of molybdenum disilicide. J. of Mat. Sci. vol. 29. 1994. pp. 2557-2571. cited by applicant .
Suzuki, Y., Morgan, P.E.D., and Niihara, K. (1998). Improvement in mechanical properties of powder-processed MoSi2 by the addition of Sc2O3 and Y2O3. J. Am. Ceram. Soci. 81(12). pp. 3141-3149. cited by applicant .
Webster, J.D., Westwood, M.E., Hayes, F.H., Day, R.J., Taylor, R., Duran, A., . . . Vogel, W.D. (1998). Oxidation protection coatings for C/SiC based on yttrium silicate. Journal of European Ceramic Society vol. 18. 1998. pp. 2345-2350. cited by applicant .
Petrovic, J.J., Castro, R.G., Vaidya, R.U., Peters, M.I., Mendoza, D., Hoover, R.C., and Gallegos, D. E. (2001). Molybdenum disilicide materials for glass melting sensor sheaths. Ceramic Engineering and Science Proceedings. vol. 22(3). 2001. pp. 59-64. cited by applicant .
Kahn, H., Tayebi, N., Ballarini, R., Mullen, R.L., Heuer, A.H. (2000). Fracture toughness of polysilicon MEMS devices. Sensors and Actuators vol. 82. 2000. pp. 274-280. cited by applicant .
Muhlstein, C.L., Stach, E.A., and Ritchie, R.O. (2002). A reaction-layer mechanism for the delayed failure of micron-scale polycrystalline silicon structural films subjected to high-cycle fatigue loading. Acta Materialia vol. 50. 2002. pp. 3579-3595. cited by applicant .
Sundaram, S.K., Hsu, J-Y., Speyer, R.F. (1994). Molten glass corrosion resistance of immersed combustion-heating tube materials in soda-lime-silicate glass. J. Am. Ceram. Soc. 77(6). pp. 1613-1623. cited by applicant .
Leckie, F.A. and Dal Bello, D.J. (2009). Strength and stiffness of engineering systems. Mechanical Engineering Series. Springer. pp. 1-3. cited by applicant .
El-Sayad, A.F. (2008). Aircraft propulsion and gas turbine engines. Boca Raton, FL: CRC Press. pp. 215-219 and 855-860. cited by applicant .
Bunker, R.S. (2005). A review of shaped hole turbine film-cooling technology. Journal of Heat Transfer vol. 127. Apr. 2005. pp. 441-453. cited by applicant .
International Search Report and Written Opinion for International Application No. PCT/US2013020746 completed dated Aug. 29, 2013. cited by applicant .
Extended European Search Report for European Application No. 13768600.2 dated Nov. 12, 2015. cited by applicant .
Davies, D. and Miller, D.C. (1971). A variable pitch fan for an ultra quiet demonstrator engine. 1976 Spring Convention: Seeds for Success in Civil Aircraft Design in the Next Two Decades. pp. 1-18. cited by applicant .
Middleton, P. (1971). 614: VFW's jet feederliner. Flight International, Nov. 4, 1971. p. 725, 729-732. cited by applicant .
Schaefer, J.W., Sagerser, D.R., and Stakolich, E.G. (1977). Dynamics of high-bypass-engine thrust reversal using a variable-pitch fan. Technical Report prepare for NASA. NASA-TM-X-3524. May 1, 1977. pp. 1-31. cited by applicant .
Savelle, S.A. and Garrard, G.D. (1996). Application of transient and dynamic simulations to the U.S. Army T55-L-712 helicopter engine. The American Society of Mechanical Engineers. Presented Jun. 10-13, 1996. pp. 1-8. cited by applicant .
Drago, R.J. and Margasahayam, R.N. (1987). Stress analysis of planet gears with integral bearings; 3D finite-element model development and test validation. 1987 MSC NASTRAN World Users Conference. Los Angeles, CA. Mar. 1987. pp. 1-14. cited by applicant .
Baker, R.W. (2000). Membrane technology and applications. New York, NY: McGraw-Hill. pp. 87-153. cited by applicant .
Cheryan, M. (1998). Ultrafiltration and microfiltration handbook. Lancaster, PA: Tecnomic Publishing Company, Inc. pp. 171-236. cited by applicant .
Seader, J.D. and Henley, E.J. (1998). Separation process principles. New York, NY: John Wiley & Sons, Inc. pp. 722-726 and 764-771. cited by applicant .
Spadaccini, L.J., and Huang, H. (2002). On-line fuel deoxygenation for coke suppression. ASME, Jun. 2002. pp. 1-7. cited by applicant .
Darrah, S. (1987). Jet fuel deoxygenation. Interim Report for Period Mar. 1987-Jul. 1988. pp. 1-22. cited by applicant .
Bucknell, R.L. (1973). Influence of fuels and lubricants on turbine engine design and performance, fuel and lubricant analyses. Final Technical Report, Mar. 1971-Mar. 1973. pp. 1-252. cited by applicant .
Hazlett, R.N. (1991). Thermal oxidation stability of aviation turbine fuels. Philadelphia, PA: ASTM. pp. 1-163. cited by applicant .
Taylor, W.F. (1974). Deposit formation from deoxygenated hydrocarbons. I. General features. Ind. Eng. Chem., Prod. Res. Develop., vol. 13(2). 1974. pp. 133-38. cited by applicant .
Taylor, W.F. (1974). Deposit formation from deoxygenated hydrocarbons. II. Effect of trace sulfur compounds. Ind. Eng. Chem., Prod. Res. Dev., vol. 15(1). 1974. pp. 64-68. cited by applicant .
Taylor, W.F. and Frankenfeld, J.W. (1978). Deposit fromation from deoxygenated hydrocarbons. 3. Effects of trace nitrogen and oxygen compounds. Ind. Eng. Chem., Prod. Res. Dev., vol. 17(1). 1978. pp. 86-90. cited by applicant .
Frankenfeld, J.W. and Taylor, W.F. (1980). Deposit fromation from deoxygenated hydrocarbons. 4. Studies in pure compound systems. Ind. Eng. Chem., Prod. Res. Dev., vol. 19(1). 1978. pp. 65-70. cited by applicant .
Hemighaus, G., Boval, T., Bacha, J., Barnes, F., Franklin, M., Gibbs, L., . . . Morris, J. (2007). Aviation fuels: Techincal review. Chevron Products Company. pp. 1-94. Retrieved from: https://www.cgabusinessdesk.com/document/aviation_tech_review.pdf. cited by applicant .
Spadaccini, L.J., Sobel, D.R., and Huang, H. (2001). Deposit formation and mitigation in aircraft fuels. Journal of Eng. for Gas Turbine and Power, vol. 123. Oct. 2001. pp. 741-746. cited by applicant .
Edwards, T. and Zabarnick, S. (1993). Supercritical fuel deposition mechanisms. Ind. Eng. Chem. Res. vol. 32. 1993. pp. 3117-3122. cited by applicant .
Huang, H., Sobel, D.R., and Spadaccini, L.J. (2002). Endothermic heat-sink of hydrocarbon fuels for scramjet cooling. AIAA/ASME/SAE/ASEE, Jul. 2002. pp. 1-7. cited by applicant .
Bessarabov, D.G., Jacobs, E.P., Sanderson, R.D., and Beckman, I.N. (1996). Use of nonporous polymeric flat-sheet gas-separation membranes in a membrane-liquid contactor: experimental studies. Journal of Membrane Sciences, vol. 113. 1996. pp. 275-284. cited by applicant .
Matsumoto, T., Toshiro, U., Kishida, A., Tsutomu, F., Maruyama, I., and Akashi, M. (1996). Novel functional polymers: Poly (dimethylsiloxane)-polyamide multiblock copolymer. VII. Oxygen permeability of aramid-silicone membranes in a gas-membrane-liquid system. Journal of Applied Polymer Science, vol. 64(6). May 9, 1997. pp. 1153-1159. cited by applicant .
Technical Data. Teflon. WS Hampshire Inc. Retrieved from: http://catalog.wshampshire.com/Asset/psg_teflon_pffe.pdf. cited by applicant .
Anderson, N.E., Loewenthal, S.H., and Black, J.D. (1984). An analytical method to predict efficiency of aircraft gearboxes. NASA Technical Memorandum prepared for the Twentieth Joint Propulsion Conference. Cincinnati, OH. Jun. 11-13, 1984. pp. 1-25. cited by applicant .
Edkins, D.P., Hirschkron, R., and Lee, R. (1972). TF34 turbofan quiet engine study. Final Report prepared for NASA. NASA-CR-120914. Jan. 1, 1972. pp. 1-99. cited by applicant .
Waters, M.N. and Schairer, E.T. (1977). Analysis of turbofan propulsion system weight and dimensions. NASA Technical Memorandum. Jan. 1977. pp. 1-65. cited by applicant .
Meyer, A.G. (1988). Transmission development of TEXTRON Lycoming's geared fan engine. Technical Paper. Oct. 1988. pp. 1-12. cited by applicant .
Dudley, D.W., Ed. (1962). Gear handbook. New York, NY: McGraw-Hill. pp. 14-17 (TOC, Preface, and Index). cited by applicant .
Hughes, C. (2002). Aerodynamic performance of scale-model turbofan outlet guide vanes designed for low noise. Prepared for the 40th Aerospace Sciences Meeting and Exhibit. Reno, NV. NASA/TM-2001-211352. Jan. 14-17, 2002. pp. 1-38. cited by applicant .
Kaplan, B., Nicke, E., Voss, C. (2006), Design of a highly efficient low-noise fan for ultra-high bypass engines. Proceedings of GT2006 for ASME Turbo Expo 2006: Power for Land, Sea and Air. Barcelona, SP. May 8-11, 2006. pp. 1-10. cited by applicant .
Gates, D. Bombardier flies at higher market. Seattle Times. Jul. 13, 2008. pp. C6. cited by applicant .
Decker, S. and Clough, R. (2016). Ge wins shot at voiding pratt patent in jet-engine clash. Bloomberg Technology. Retrieved from: https://www.bloomberg.com/news/articles/2016-06-30/ge-wins-shot-to-invali- date-pratt-airplane-engine-patent-in-u-s. cited by applicant .
Trembley, Jr., H.F. (1977). Determination of effects of ambient conditions on aircraft engine emissions. ALF 502 combustor rig testing and engine verification test. Prepared for Environmental Protection Agency. Sep. 1977. pp. 1-256. cited by applicant .
Lewicki, D.G., Black, J.D., Savage, M., and Coy, J.J. (1985). Fatigue life analysis of a turboprop reduction gearbox. NASA Technical Memorandum. Prepared for the Design Technical Conference (ASME). Sep. 11-13, 1985. pp. 1-26. cited by applicant .
McCune, M.E. (1993). Initial test results of 40,000 horsepower fan drive gear system for advanced ducted propulsion systems. AIAA 29th Joint Conference and Exhibit. Jun. 28-30, 1993. pp. 1-10. cited by applicant .
Wright, G.H. and Russell, J.G. (1990). The M.45SD-02 variable pitch geared fan engine demonstrator test and evaluation experience. Aeronautical Journal., vol. 84(836). Sep. 1980. pp. 268-277. cited by applicant .
Drago, R.J. (1974). Heavy-lift helicopter brings up drive ideas. Power Transmission Design. Mar. 1987. pp. 1-15. cited by applicant .
Krantz, T.L. (1990). Experimental and analytical evaluation of efficiency of helicopter planetary stage. NASA Technical Paper. Nov. 1990. pp. 1-19. cited by applicant .
Heingartner, P., MBA, D., Brown, D. (2003). Determining power losses in the helical gear mesh; Case Study ASME 2003 Design Engineering Technical Conferences. Chicago, IL. Sep. 2-6, 2003. pp. 1-7. cited by applicant .
Thulin, R.D., Howe, D.C., and Singer, I.D. (1982). Energy efficient engine: High pressure turbine detailed design report. Prepared for NASA. NASA CR-165608. Received Aug. 9, 1984. pp. 1-178. cited by applicant .
Reshotko, M., Karchmer, A., Penko, P.F. (1977). Core noise measurements on a YF-102 turbofan engine Prepared for Aerospace Sciences Meeting sponsored by the American Institute of Aeronautics and Astronautics. Jan. 24-26, 2977. cited by applicant .
McMillian, A. (2008) Material development for fan blade containment casing. Abstract. p. 1. Conference on Engineering and Physics: Synergy for Success 2006. Journal of Physics: Conference Series vol. 105. London, UK. Oct. 5, 2006. cited by applicant .
Kurzke, J. (2009). Fundamental differences between conventional and geared turbofans. Proceedings of ASME Turbo Expo: Power for Land, Sea, and Air. 2009, Orlando, Florida. pp. 145-151. cited by applicant .
Agarwal, B.D and Broutman, L.J. (1990). Analysis and performance of fiber composites, 2nd Edition. John Wiley & Sons, Inc. New York: New York. pp. 1-30, 50-51, 56-58, 60-61, 64-71, 87-89, 324-329, 436-437. cited by applicant .
Carney, K., Pereira, M. Revilock, and Matheny, P. (2003). Jet engine fan blade containment using two alternate geometries. 4th European LS-DYNA Users Conference. pp. 1-10. cited by applicant .
Brines, G.L. (1990). The turbofan of tomorrow. Mechanical Engineering: The Journal of the American Society of Mechanical Engineers,108(8), 65-67. cited by applicant .
Faghri, A. (1995). Heat pipe and science technology. Washington, D.C.: Taylor & Francis. pp. 1-60. cited by applicant .
Hess, C. (1998). Pratt & Whitney develops geared turbofan. Flug Revue 43(7). Oct. 1998. cited by applicant .
Grady, J.E., Weir, D.S., Lamoureux, M.G., and Martinez, M.M. (2007). Engine noise research in NASA's quiet aircraft technology project. Papers from the International Symposium on Air Breathing Engines (ISABE). 2007. cited by applicant .
Griffiths, B. (2005). Composite fan blade containment case. Modern Machine Shop. Retrieved from: http://www.mmsonline.com/articles/composite-fan-blade-containment-case pp. 1-4. cited by applicant .
Hall, C.A. and Crichton, D. (2007). Engine design studies for a silent aircraft. Journal of Turbomachinery, 129, 479-487. cited by applicant .
Haque, A. and Shamsuzzoha, M., Hussain, F., and Dean, D. (2003). S20-glass/epoxy polymer nanocomposites: Manufacturing, structures, thermal and mechanical properties. Journal of Composite Materials, 37(20), 1821-1837. cited by applicant .
Brennan, P.J. and Kroliczek, E.J. (1979). Heat pipe design handbook. Prepared for National Aeronautics and Space Administration by B & K Engineering, Inc. Jun. 1979. pp. 1-348. cited by applicant .
Horikoshi, S. and Serpone, N. (2013). Introduction to nanoparticles. Microwaves in nanoparticle synthesis. Wiley-VCH Verlag GmbH & Co. KGaA. pp. 1-24. cited by applicant .
Kerrebrock, J.L. (1977). Aircraft engines and gas turbines. Cambridge, MA: The MIT Press. p. 11. cited by applicant .
Xie, M. (2008). Intelligent engine systems: Smart case system. NASA/CR-2008-215233. pp. 1-31. cited by applicant .
Knip, Jr., G. (1987). Analysis of an advanced technology subsonic turbofan incorporating revolutionary materials. NASA Technical Memorandum. May 1987. pp. 1-23. cited by applicant .
Willis, W.S. (1979). Quiet clean short-haul experimental engine (QCSEE) final report NASA/CR-159473 pp. 1-289. cited by applicant .
Kojima, Y., Usuki, A. Kawasumi, M., Okada, A., Fukushim, Y., Kurauchi, T., and Kamigaito, O. (1992). Mechanical properties of nylon 6-clay hybrid. Journal of Materials Research, 8(5), 1185-1189. cited by applicant .
Kollar, L.P. and Springer, G.S. (2003). Mechanics of composite structures. Cambridge, UK: Cambridge University Press. p. 465. cited by applicant .
Ramsden, J.M. (Ed). (1978). The new European airliner. Flight International, 113(3590). Jan. 7, 1978. pp. 39-43. cited by applicant .
Langston, L. and Faghri, A. Heat pipe turbine vane cooling. Prepared for Advanced Turbine Systems Annual Program Review. Morgantown, West Virginia. Oct. 17-19, 1995. pp. 3-9. cited by applicant .
Oates, G.C. (Ed). (1989). Aircraft propulsion systems and technology and design. Washington, D.C.: American Institute of Aeronautics, Inc. pp. 341-344. cited by applicant .
Lau, K., Gu, C., and Hui, D. (2005). A critical review on nanotube and nanotube/nanoclay related polymer composite materials. Composites: Part B 37(2006) 425-436. cited by applicant .
Shorter Oxford English dictionary, fith Edition. (2007). vol. 2, N-Z. p. 1888. cited by applicant .
Lynwander, P. (1983). Gear drive systems: Design and application. New York, New York: Marcel Dekker, Inc. pp. 145, 355-358. cited by applicant .
Sweetman, B. and Sutton, O. (1998). Pratt & Whitney's surprise leap. Interavia Business & Technology, 53.621, p. 25. cited by applicant .
Mattingly, J.D. (1996). Elements of gas turbine propulsion. New York, New York: McGraw-Hill, Inc. pp. 8-15. cited by applicant .
Pyrograf-III Carbon Nanofiber. Product guide. Retrieved Dec. 1, 2015 from: http://pyrografproducts.com/Merchant5/merchant.mvc?Screen=cp_nanofiber. cited by applicant .
Nanocor Technical Data for Epoxy Nanocomposites using Nanomer 1.30E Nanoclay. Nnacor, Inc. Oct. 2004. cited by applicant .
Ratna, D. (2009). Handbook of thermoset resins. Shawbury, UK: iSmithers. pp. 187-216. cited by applicant .
Wendus, B.E., Stark, D.F., Holler, R.P., and Funkhouser, M.E. (2003). Follow-on technology requirement study for advanced subsonic transport. NASA/CR-2003-212467. pp. 1-37. cited by applicant .
Silverstein, C.C., Gottschlich, J.M., and Meininger, M. The feasibility of heat pipe turbine vane cooling. Presented at the International Gas Turbine and Aeroengine Congress and Exposition, The Hague, Netherlands. Jun. 13-16, 1994.pp. 1-7. cited by applicant .
Merriam-Webster's collegiate dictionary, 11th Ed. (2009). p. 824. cited by applicant .
Merriam-Webster's collegiate dictionary, 10th Ed. (2001). p. 1125-1126. cited by applicant .
Whitaker, R. (1982). ALF 502: plugging the turbofan gap. Flight International, p. 237-241, Jan. 30, 1982. cited by applicant .
Hughes, C. (2010). Geared turbofan technology. NASA Environmentally Responsible Aviation Project. Green Aviation Summit. NASA Ames Research Center. Sep. 8-9, 2010. pp. 1-8. cited by applicant .
Gliebe, P.R. and Janardan, B.A. (2003). Ultra-high bypass engine aeroacoustic study. NASA/CR-2003-21252. GE Aircraft Engines, Cincinnati, Ohio. Oct. 2003. pp. 1-103. cited by applicant .
Moxon, J. How to save fuel in tomorrow's engines. Flight International. Jul. 30, 1983. 3873(124). pp. 272-273. cited by applicant .
File History for U.S. Appl. No. 12/131,876. cited by applicant .
Cusick, M. (1981). Avco Lycoming's ALF 502 high bypass fan engine. Society of Automotive Engineers, inc. Business Aircraft Meeting & Exposition. Wichita, Kansas. Apr. 7-10, 1981. pp. 1-9. cited by applicant .
Fledderjohn, K.R. (1983). The TFE731-5: Evolution of a decade of business jet service. SAE Technical Paper Series. Business Aircraft Meeting & Exposition. Wichita, Kansas. Apr. 12-15, 1983. pp. 1-12. cited by applicant .
Dickey, T.A. and Dobak, E.R. (1972). The evolution and development status of ALF 502 turbofan engine. National Aerospace Engineering and Manufacturing Meeting. San Diego, California. Oct. 2-5, 1972. pp. 1-12. cited by applicant .
Gunston, B. (Ed.) (2000). Jane's aero-engines, Issue seven. Coulsdon, Surrey, UK: Jane's Information Group Limited. pp. 510-512. cited by applicant .
Ivchenko-Progress D-436. Jane's Aero-engines, Aero-engines--Turbofan. Feb. 8, 2012. cited by applicant .
Ivchenko-Progress AI-727M. Jane's Aero-engines, Aero-engines--Turbofan. Nov. 27, 2011. cited by applicant .
Ivchenko-Progress D-727. Jane's Aero-engines, Aero-engines--Turbofan. Feb. 7, 2007. cited by applicant .
Turbomeca Aubisque. Jane's Aero-engines, Aero-engines--Turbofan. Nov. 2, 2009. cited by applicant .
Aviadvigatel D-110. Jane's Aero-engines, Aero-engines--Turbofan. Jun. 1, 2010. cited by applicant .
Rolls-Royce M45H. Jane's Aero-engines, Aero-engines--Turbofan. Feb. 24, 2010. cited by applicant .
Honeywell LF502. Jane's Aero-engines, Aero-engines--Turbofan. Feb. 9, 2012. cited by applicant.

Primary Examiner: Sung; Gerald L
Attorney, Agent or Firm: Carlson, Gaskey & Olds, P.C.

Parent Case Text

CROSS-REFERENCE TO RELATED APPLICATION

This application is a divisional of U.S. application Ser. No. 14/246,260, filed Apr. 7, 2014, which is a continuation of U.S. application Ser. No. 13/362,499, filed Jan. 31, 2012.

Claims

What is claimed is:

1. A method of designing a buffer system for a gas turbine engine, the method comprising the steps of: configuring a heat exchanger to define a first inlet and a first outlet fluidly coupled to each other and a second inlet and a second outlet fluidly coupled to each other; configuring the first inlet to receive air from first and second air sources that are selectively fluidly coupled to the first inlet; configuring the second inlet to receive air from a third air source that is fluidly coupled to the second inlet; configuring the first outlet to provide cooled pressurized air to multiple fluid-supplied areas that are located remotely from one another and that are fluidly coupled to the first outlet, the multiple fluid-supplied areas including a bearing compartment of the gas turbine engine; configuring a valve to be fluidly coupled to the first and second inlets; designing the valve to regulate air from the first and second air sources, wherein the first and second air sources have different pressures; designing the first, second and third air sources to be different than one another; configuring the first air source to provide air at a first pressure and temperature state; configuring the second air source to provide air at a higher pressure and temperature state than the first air source; configuring a compressor section to compress air and deliver the air into a combustion section, the compressor section including first and second compressors; configuring the first compressor to provide the third air source; and configuring the second compressor to provide the second air source.

2. The method as recited in claim 1, comprising the steps of: configuring a mid-stage to provide a flow path between first and second compressors; and configuring the mid-stage to provide the first air source.

3. The method as recited in claim 1, wherein the third air source is provided by the compressor section.

4. The method as recited in claim 1, wherein the step of configuring the compressor section includes arranging the first compressor to be upstream of the second compressor.

5. The method as recited in claim 4, wherein the step of configuring the compressor section includes arranging the first compressor and the second compressor downstream of a fan.

6. The method as recited in claim 5, wherein the third air source is provided by the compressor section.

7. The method as recited in claim 6, wherein a pressure ratio across the fan is less than 1.45.

8. The method as recited in claim 7, comprising the step of: configuring a turbine section to drive the compressor section, the turbine section including a fan drive turbine configured to drive the fan.

9. The method as recited in claim 8, wherein the bearing compartment includes a plurality of bearing compartments.

10. The method as recited in claim 8, wherein the multiple fluid-supplied areas include one of a vane, a blade and a clearance control device.

11. The method as recited in claim 1, comprising the steps of: configuring a mid-stage to provide a flow path between the first and second compressors; and configuring the mid-stage to provide the first air source.

12. The method as recited in claim 1, comprising the step of: designing the multiple fluid-supplied areas to include a component of the gas turbine engine, the component being at least one of a rotor of the second compressor and a shaft of a fan drive turbine driven by the compressor section.

13. The method as recited in claim 12, wherein the multiple fluid-supplied areas includes the shaft.

14. The method as recited in claim 1, comprising the step of: configuring the second outlet to be fluidly coupled to a bypass flow path, the bypass flow path being arranged between core and fan nacelles of the gas turbine engine.

15. The method as recited in claim 14, comprising the steps of: configuring a fan to deliver a portion of air into the compressor section, and a portion of air into a bypass duct defining the bypass flow path; and configuring a turbine section to drive the compressor section, the turbine section including a fan drive turbine configured to drive the fan.

Description

BACKGROUND

This disclosure relates to a buffer system for a gas turbine engine.

Gas turbine engines typically require air from one of the compressor stages to provide buffer air to a bearing compartment, for example. One proposed system utilizes a valve in the buffer system that regulates the fluid flow from a high compressor stage and a low compressor stage. The regulated air is provided to a cooler and a bearing compartment.

Another system utilizes a dedicated centrifugal compressor to provide pressurized air to multiple bearing compartments and provide shaft ventilation. The centrifugal compressor is a separate compressor, discrete from the high pressure and low pressure compressor stages of the gas turbine engine.

SUMMARY

A gas turbine engine according to an example of the present disclosure includes a fan, a compressor section fluidly connected to the fan, a combustor fluidly connected to the compressor section, a turbine section fluidly connected to the combustor, and a buffer system. The buffer system includes a heat exchanger having a first inlet, a first outlet, a second inlet, and a second outlet. The first outlet is configured to provide a cooled pressurized fluid, first and second air sources that are selectively fluidly coupled to the first inlet, a third air source fluidly coupled to the second inlet, and multiple fluid-supplied areas located remotely from one another and fluidly coupled to the first outlet. The multiple fluid-supplied areas include a bearing compartment.

A further embodiment of any of the foregoing embodiments includes a first compressor and a second compressor. The second air source is provided by the second compressor, and the third air source is provided by the first compressor.

In a further embodiment of any of the foregoing embodiments, the first, second and third air sources are different than one another.

In a further embodiment of any of the foregoing embodiments, the first air source is configured to provide air at a first pressure and temperature state, and the second air source is configured to provide air at a higher pressure and temperature state than the first air source.

In a further embodiment of any of the foregoing embodiments, a valve is fluidly coupled to at least one of the first and second air sources. The valve is configured to regulate fluid flow from at least one of the one first and second air sources to the heat exchanger.

In a further embodiment of any of the foregoing embodiments, a mid-stage is provided fluidly between the first and second compressors. The first air source is provided by the mid-stage.

In a further embodiment of any of the foregoing embodiments, the second compressor includes a rotor, and the multiple fluid-supplied areas include the rotor.

In a further embodiment of any of the foregoing embodiments, the turbine section comprises a fan drive turbine including a shaft, and the multiple fluid-supplied areas include the shaft.

In a further embodiment of any of the foregoing embodiments, a pressure ratio across the fan is less than about 1.45.

In a further embodiment of any of the foregoing embodiments, fan drive turbine is configured to drive the fan. A pressure ratio across the fan drive turbine is greater than about 5.

In a further embodiment of any of the foregoing embodiments, the first inlet is fluidly connected to the first outlet, and the second inlet is fluidly connected to the second outlet.

In a further embodiment of any of the foregoing embodiments, the multiple fluid-supplied areas include a component of the gas turbine engine.

In a further embodiment of any of the foregoing embodiments, the component is one of a vane, a blade and a clearance control device.

A further embodiment of any of the foregoing embodiments includes a bypass flow path arranged between core and fan nacelles, the second outlet fluidly being coupled to the bypass flow path.

In a further embodiment of any of the foregoing embodiments, the fan is configured to deliver a portion of air into the compressor section, and a portion of air into the bypass duct. A bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, is greater than about 6.

A method of designing a buffer system for a gas turbine engine according to an example of the present disclosure includes the steps of configuring a heat exchanger to define a first inlet and a first outlet fluidly coupled to each other and a second inlet and a second outlet fluidly coupled to each other, configuring the first inlet to receive air from first and second air sources that are selectively fluidly coupled to the first inlet, configuring the second inlet to receive air from a third air source fluidly that is coupled to the second inlet, and configuring the first outlet to provide cooled pressurized air to multiple fluid-supplied areas that are located remotely from one another and that are fluidly coupled to the first outlet. The multiple fluid-supplied areas include a bearing compartment of a gas turbine engine.

A further embodiment of any of the foregoing embodiments includes the steps of configuring a valve to be fluidly coupled to the first and second inlets, and designing the valve to regulate air from the first and second air sources, wherein the first and second air sources have different pressures.

A further embodiment of any of the foregoing embodiments includes the step of designing the first, second and third air sources to be different than one another.

A further embodiment of any of the foregoing embodiments includes the steps of configuring the first air source to provide air at a first pressure and temperature state, and configuring the second air source to provide air at a higher pressure and temperature state than the first air source.

A further embodiment of any of the foregoing embodiments includes the steps of configuring a compressor section to compress air and deliver it into a combustion section, the compressor section including first and second compressors; configuring the first compressor to provide the third air source; and configuring the second compressor to provide the second air source.

In a further embodiment of any of the foregoing embodiments, the step of configuring a compressor section includes arranging the first compressor to be upstream of the second compressor.

A further embodiment of any of the foregoing embodiments includes the steps of configuring a mid-stage to provide a flow path between the first and second compressors, and configuring the mid-stage to provide the first air source.

A further embodiment of any of the foregoing embodiments includes the step of designing the multiple fluid-supplied areas to include a component of the gas turbine engine. The component is at least one of a rotor of the second compressor and a shaft of a fan drive turbine driven by the compressor section.

A further embodiment of any of the foregoing embodiments includes the step of configuring the second outlet to be fluidly coupled to a bypass flow path. The bypass flow path is arranged between core and fan nacelles of a gas turbine engine.

A buffer system for a gas turbine engine according to an example of the present disclosure includes a heat exchanger configured for securing to a static structure. The heat exchanger has a first inlet and outlet and a second inlet and outlet, first and second air sources selectively fluidly coupled to the first inlet, and a third air source fluidly coupled to the second inlet and being provided by a first compressor. The first outlet is configured to provide a cooled pressurized fluid to multiple fluid-supplied areas fluidly coupled to the first outlet. The multiple fluid-supplied areas are located remotely from one another and include a bearing compartment.

In a further embodiment of any of the foregoing embodiments, the first air source is configured to provide air at a low pressure and temperature state, and the second air source is configured to provide air at a higher pressure and temperature state than the first air source.

In a further embodiment of any of the foregoing embodiments, the second air source is provided by a second compressor.

A further embodiment of any of the foregoing embodiments includes a valve configured for securing to the static structure, the valve being fluidly coupled to at least one of the first and second air sources. The valve is configured to regulate fluid flow from the first and second air sources to the heat exchanger.

In a further embodiment of any of the foregoing embodiments, the first inlet is fluidly connected to the first outlet, and the second inlet is fluidly connected to the second outlet.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the following detailed description when considered in connection with the accompanying drawings wherein:

FIG. 1 schematically illustrates a gas turbine engine.

FIG. 2 is a schematic of an example buffer system for a gas turbine engine.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbofan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Alternative engines might include an augmentor section (not shown) among other systems or features. The fan section 22 drives air along a bypass flowpath B while the compressor section 24 drives air along a core flowpath C for compression and communication into the combustor section 26 then expansion through the turbine section 28. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures.

The engine 20 generally includes a low speed spool 30 and a high speed spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing systems 38. It should be understood that various bearing systems 38 at various locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor 44 and a low pressure turbine 46. The inner shaft 40 is connected to the fan 42 through a geared architecture 48 to drive the fan 42 at a lower speed than the low speed spool 30. The high speed spool 32 includes an outer shaft 50 that interconnects a high pressure compressor 52 and high pressure turbine 54. A combustor 56 is arranged between the high pressure compressor 52 and the high pressure turbine 54. A mid-turbine frame 57 of the engine static structure 36 is arranged generally between the high pressure turbine 54 and the low pressure turbine 46. The mid-turbine frame 57 supports one or more bearing systems 38 in the turbine section 28. The inner shaft 40 and the outer shaft 50 are concentric and rotate via bearing systems 38 about the engine central longitudinal axis A, which is collinear with their longitudinal axes.

The core airflow C is compressed by the low pressure compressor 44 then the high pressure compressor 52, mixed and burned with fuel in the combustor 56, then expanded over the high pressure turbine 54 and low pressure turbine 46. The mid-turbine frame 57 includes airfoils 59 which are in the core airflow path. The turbines 46, 54 rotationally drive the respective low speed spool 30 and high speed spool 32 in response to the expansion.

The engine 20 in one example a high-bypass geared aircraft engine. In a further example, the engine 20 bypass ratio is greater than about six (6), with an example embodiment being greater than ten (10), the geared architecture 48 is an epicyclic gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.3 and the low pressure turbine 46 has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine 20 bypass ratio is greater than about ten (10:1), the fan diameter is significantly larger than that of the low pressure compressor 44, and the low pressure turbine 46 has a pressure ratio that is greater than about 5:1. Low pressure turbine 46 pressure ratio is pressure measured prior to inlet of low pressure turbine 46 as related to the pressure at the outlet of the low pressure turbine 46 prior to an exhaust nozzle. The geared architecture 48 may be an epicycle gear train, such as a planetary gear system or other gear system, with a gear reduction ratio of greater than about 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section 22 of the engine 20 is designed for a particular flight condition--typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, with the engine at its best fuel consumption--also known as "bucket cruise Thrust Specific Fuel Consumption (`TSFC`)"--is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. "Low fan pressure ratio" is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane ("FEGV") system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.45. "Low corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tambient deg R)/518.7){circumflex over ( )}0.5]. The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about 1150 ft/second.

An example buffer system 60 is illustrated in FIG. 2. The system 60 includes first, second and third air sources 62, 64, 66, which are different than one another in the example. In one example, one of the first and second air sources 62, 64 is provided by the high pressure compressor 52. The third air source 66 is provided by the low pressure compressor 44. A pressurized region 100 between the exit of the low pressure compressor 44 and the entrance of high pressure compressor 52 is referred to as "mid-stage" (100 in FIG. 1) and is arranged fluidly with the low and high pressure compressors 44, 52 within the compressor section 24. In the example, the first and second air sources 62, 64 are provided within the high pressure compressor but could be representative of any two differing pressure supplies. Thus, in one example, the first air source 62 provides air at a low pressure and temperature state, the second air source 64 provides air at a higher pressure and temperature state, and the third air source 66 provides low pressure compressor air.

The first and second air sources 62, 64 are selectively fluidly coupled to a heat exchanger 72. In one example, a valve 68 is fluidly coupled to the first and second air sources 62, 64 and is configured to regulate fluid flow from the first and second air sources 62, 64 to the heat exchanger 72.

The heat exchanger 72 includes a first inlet and outlet 69, 70 and a second inlet and outlet 71, 73. The first outlet 70 provides cooled pressurized air 74. Passages are provided between respective inlets and outlets and are configured to transfer heat between the passages. In the example, the valve 68 is fluidly coupled to the first inlet 69, and the third air source 66 is fluidly coupled to the second inlet 71. The second outlet 73 is fluidly coupled to the bypass flow B so that low pressure compressor air expelled from the heat exchanger 72 may be used to supplement the thrust provided by the bypass flow B.

The first outlet 70 is fluidly coupled to multiple fluid-supplied areas that are located remotely from one another. The multiple fluid-supplied areas include multiple bearing compartments 76A, 76B, 76C, 76D. The multiple fluid-supplied areas may also include a component 78 that requires thermal conditioning, such as a vane, blade or clearance control device. The multiple fluid-supplied areas may also include a rotor 80 in the high pressure compressor 52 and/or a low pressure turbine section shaft 82.

In operation, a method of providing pressurized air to the gas turbine engine 10 includes selectively providing pressurized air from multiple air sources to the heat exchanger 72 to cool the pressurized air. The cooled pressurized air 74 is distributed to multiple fluid-supplied areas within the gas turbine engine. In one example, air from first and second air sources 62, 64, which are different from one another, are regulated. The pressurized air is cooled by providing low pressure compressor air from a third air source 66 to the heat exchanger 72. The cooled pressurized air 74 is distributed to multiple bearing compartments 76A, 76B, 76C, 76D, high pressure compressor section rotor 80, low pressure turbine section shaft 82, and/or a component 78.

Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of the claims. For that reason, the following claims should be studied to determine their true scope and content.

* * * * *

Current U.S. Class:	1/1
Current CPC Class:	F02C 7/185 (20130101); F02C 6/08 (20130101); F01D 25/183 (20130101); F02C 7/18 (20130101); F01D 25/22 (20130101); Y10T 29/49359 (20150115); Y02T 50/60 (20130101)
Current International Class:	F02C 1/00 (20060101); F01D 25/18 (20060101); F02C 7/18 (20060101); F01D 25/22 (20060101); F02C 6/08 (20060101)