IP Library › Granted Patent US 12,618,332
Granted Patent B2
US 12,618,332 · App. 19/045,244 · Granted May 5, 2026

Gas turbine engine

Inventors: Daniel Alan Niergarth (Norwood, OH); Jeffrey Donald Clements (Mason, OH); Jeffrey S. Spruill (Hillsboro, OH); Erich Alois Krammer (West Chester, OH); Matthew Kenneth MacDonald (Austin, TX); Scott Alan Schimmels (Miamisburg, OH); David P. Mourer (Beverly, MA)
Assignee: General Electric Company
F01D5/288C22C19/007C22C19/07C22C30/00F02C9/28F05D2300/175F05D2300/611
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Quick Facts
Patent No.
US 12,618,332
App. No.
19/045,244
Granted
May 5, 2026
Kind
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).

Claims (24)

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 . The gas turbine engine of claim 1 , wherein the component is in a high pressure turbine portion of the turbine section.

3 . The gas turbine engine of claim 1 , wherein the component is in a low pressure turbine portion of the turbine section.

4 . The gas turbine engine of claim 1 , wherein the coating comprises a distribution of pinning agents.

5 . The gas turbine engine of claim 4 , wherein the pinning agents are located at the interfaces between grains defined in the coating.

6 . The gas turbine engine of claim 5 , wherein the grains have an average grain size of 0.1 microns to 5 microns.

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 . 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 . 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 . The gas turbine engine of claim 8 , wherein the coating comprises 5 wt % to 14 wt % aluminum.

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 . 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 . 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 . 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 . The gas turbine engine of claim 1 , wherein the coating has a thickness on the component that is 10 μm to 90 μm.

16 . The gas turbine engine of claim 1 , wherein the EGT is greater than 1000 degrees Celsius and less than 1300 degrees Celsius.

17 . The gas turbine engine of claim 1 , wherein the EGT is greater than 1100 degree Celsius and less than 1250 degrees Celsius.

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 . 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 . 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.

Assignments (1)
ASSIGNMENT OF ASSIGNOR'S INTEREST Recorded Sep 30, 2025
From: NIERGARTH, DANIEL ALAN; CLEMENTS, JEFFREY DONALD; SPRUILL, JEFFREY S.; KRAMMER, ERICH ALOIS; MACDONALD, MATTHEW KENNETH; SCHIMMELS, SCOTT ALAN; MOURER, DAVID P.
To: GENERAL ELECTRIC COMPANY
Reel/Frame 072424/0731 →
Continuity (3)
Continuation In Part 18481515 · Oct 5, 2023
Continuation In Part 17978629 · Nov 1, 2022
Related Publication 20250179923A1 · Jun 5, 2025
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