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US-12618332-B2 - Gas turbine engine

US12618332B2US 12618332 B2US12618332 B2US 12618332B2US-12618332-B2

Abstract

A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (A HPCExit ) in square inches; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn Total ) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: Fn Total ×EGT/(A HPCExit 2 ×1000).

Inventors

  • Daniel Alan NIERGARTH
  • Jeffrey Donald Clements
  • Jeffrey S. Spruill
  • Erich Alois Krammer
  • Matthew Kenneth MacDonald
  • Scott Alan Schimmels
  • David P. Mourer

Assignees

  • GENERAL ELECTRIC COMPANY

Dates

Publication Date
20260505
Application Date
20250204

Claims (20)

  1. 1 . A gas turbine engine comprising: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (A HPCExit ) in square inches; and a component within the turbomachine, wherein the component has a coating on an external surface of the component that is exposed to a hot gas flow path in the turbomachine, and wherein the coating comprises: cobalt, chromium, aluminum, yttrium, and nickel, wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn Total ) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific thrust determined as follows: Fn Total ×EGT/(A HPCExit 2 ×1000).
  2. 2 . The gas turbine engine of claim 1 , wherein the component is in a high pressure turbine portion of the turbine section.
  3. 3 . The gas turbine engine of claim 1 , wherein the component is in a low pressure turbine portion of the turbine section.
  4. 4 . The gas turbine engine of claim 1 , wherein the coating comprises a distribution of pinning agents.
  5. 5 . The gas turbine engine of claim 4 , wherein the pinning agents are located at the interfaces between grains defined in the coating.
  6. 6 . The gas turbine engine of claim 5 , wherein the grains have an average grain size of 0.1 microns to 5 microns.
  7. 7 . The gas turbine engine of claim 1 , wherein the component comprises a base metal having a base chromium content, and wherein the coating has a coating chromium content that is greater than the base chromium content of the base metal.
  8. 8 . The gas turbine engine of claim 1 , wherein the coating comprises 15 wt % to 45 wt % cobalt; 20 wt % to 40 wt % chromium; 2 wt % to 15 wt % aluminum; 0.1 wt % to 1 wt % yttrium; and nickel.
  9. 9 . The gas turbine engine of claim 8 , wherein the coating comprises 30 wt % to 40 wt % cobalt, and wherein the coating comprises 55 wt % to 75 wt % of a combined amount of nickel and cobalt.
  10. 10 . The gas turbine engine of claim 8 , wherein the coating comprises 5 wt % to 14 wt % aluminum.
  11. 11 . The gas turbine engine of claim 8 , wherein the coating has a thickness on the external surface that is 5 μm to 100 μm.
  12. 12 . The gas turbine engine of claim 1 , wherein the coating further comprises at least one of lanthanum, cerium, zirconium, magnesium, a rare earth metal, or a combination thereof.
  13. 13 . The gas turbine engine of claim 1 , wherein the coating further comprises: 0 wt % to 10 wt % tungsten; 0 wt % to 10 wt % tantalum; 0 wt % to 0.5 wt % hafnium; and 0 wt % to 0.5 wt % silicon.
  14. 14 . The gas turbine engine of claim 1 , wherein the coating further comprises: tungsten, molybdenum, tantalum, rhenium, titanium, niobium, vanadium, a platinum group metal, or a combination thereof, wherein a total combined amount of these elements is 20 wt % or less.
  15. 15 . The gas turbine engine of claim 1 , wherein the coating has a thickness on the component that is 10 μm to 90 μm.
  16. 16 . The gas turbine engine of claim 1 , wherein the EGT is greater than 1000 degrees Celsius and less than 1300 degrees Celsius.
  17. 17 . The gas turbine engine of claim 1 , wherein the EGT is greater than 1100 degree Celsius and less than 1250 degrees Celsius.
  18. 18 . The gas turbine engine of claim 1 , wherein the EGT is greater than 1000 degree Celsius and less than 1300 degrees Celsius, and wherein the corrected specific thrust is greater than or equal to 45.
  19. 19 . The gas turbine engine of claim 1 , wherein the EGT is greater than 1000 degree Celsius and less than 1300 degrees Celsius, and wherein the corrected specific thrust is greater than or equal to 50.
  20. 20 . The gas turbine engine of claim 1 , wherein the turbine section comprises a high pressure turbine having a first stage of high pressure turbine rotor blades, and wherein the gas turbine engine further comprises: a cooled cooling air system in fluid communication with the first stage of high pressure turbine rotor blades.

Description

CROSS-REFERENCE TO RELATED APPLICATIONS This application is a continuation-in-part patent application of U.S. application Ser. No. 18/481,515 filed Oct. 5, 2023, which is a continuation-in-part application of U.S. application Ser. No. 17/978,629 filed Nov. 1, 2022. Each of these applications are hereby incorporated by reference in their entirety. FIELD The present disclosure relates to a gas turbine engine. BACKGROUND A gas turbine engine typically includes a fan and a turbomachine. The turbomachine generally includes an inlet, one or more compressors, a combustor, and at least one turbine. The compressors compress air which is channeled to the combustor where it is mixed with fuel. The mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the turbine(s) which extracts energy from the combustion gases for powering the compressor(s), as well as for producing useful work to propel an aircraft in flight. The turbomachine is mechanically coupled to the fan for driving the fan during operation. BRIEF DESCRIPTION OF THE DRAWINGS A full and enabling disclosure of the present disclosure, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended FIGS., in which: FIG. 1 is a schematic cross-sectional view of a three-stream engine in accordance with an exemplary embodiment of the present disclosure. FIG. 2 is a close-up, schematic view of the exemplary three-stream engine of FIG. 1 with a cooled cooling air system in accordance with an exemplary embodiment of the present disclosure. FIG. 3 is a close-up view of an aft-most stage of high pressure compressor rotor blades within the exemplary three-stream engine of FIG. 1. FIG. 4 is a close-up, schematic view of the exemplary three-stream engine of FIG. 1 showing the cooled cooling air system of FIG. 2. FIG. 5 is a schematic view of a thermal transport bus of the present disclosure. FIG. 6 is a table depicting numerical values showing the relationships between various parameters in accordance with various example embodiments of the present disclosure. FIG. 7 is a graph depicting a range of corrected specific thrust values and redline exhaust gas temperature values of gas turbine engines in accordance with various example embodiments of the present disclosure. FIG. 8 is a schematic view of a ducted turbofan engine in accordance with an exemplary aspect of the present disclosure. FIG. 9 is a schematic, close-up view of a gas turbine engine having a cooled cooling air system in accordance with another exemplary aspect of the present disclosure. FIG. 10 is a schematic, close-up view of a gas turbine engine having a cooled cooling air system in accordance with yet another exemplary aspect of the present disclosure. FIG. 11 is a schematic, close-up view of a gas turbine engine having a cooled cooling air system in accordance with still another exemplary aspect of the present disclosure. FIG. 12 is a schematic view of a turbofan engine in accordance with another exemplary aspect of the present disclosure. FIG. 13 is a cross-sectional view of a portion of the turbine section of an exemplary gas turbine engine according to an embodiment of the disclosure. FIG. 14 is a perspective view of an exemplary turbine disk of a type used in gas turbine engines according to an embodiment of the disclosure. FIG. 15 schematically represents a cross-sectional view of a corrosion and oxidation-resistant coating on a surface of one or more of the turbine components in FIG. 13 according to an embodiment of the disclosure. FIG. 16 is a micrograph showing pitting at 1300° F. of an uncoated René® 104 sample after 1 cycle. FIG. 17 is a micrograph showing that a CoNiCrAlY coated sample of René® 104 at 1300° F. did not exhibit pitting after 10 cycles. FIG. 18 is a plot of corrosion pitting as a function of chromium content in wt % in Ni-based alloys. FIG. 19 is a plot of chromium loss after about 815° C./450 hr (1500° F./450 hr) air exposure for different coating compositions. FIG. 20 is a plot of cobalt levels in coatings after thermal exposure for coatings without cobalt. FIG. 21 is a plot of cobalt levels in coatings after thermal exposure for coatings with cobalt. FIG. 22 is a cross-section of a coating having 0.18 wt % Al after about 815° C. (1500° F.) isothermal exposure. FIG. 23 is a cross-section of a coating having 2.5 wt % Al after about 815° C. (1500° F.) isothermal exposure. FIG. 24 is a graph showing coating crack initiation or failure (cycles) at 760° C. (1400° F.) LCF for various coatings. FIG. 25 is a graph illustrating fatigue life of a NiCr coating as a function of coating thickness. FIG. 26 is a graph illustrating fatigue life of a CoNiCrAlY according to the disclosure coating as a function of coating thickness. FIG. 27 is a micrograph showing cracking in a coarse grain coating tested in fatigue at about 705° C./0.713% strain range (1300° F./0.713