Nan Shi Jiang

US Patent:

20200240276, Jul 30, 2020

Filed:

Aug 14, 2017

Appl. No.:

16/638260

Inventors:

- München, DE
Nan Jiang - Charlotte NC, US

International Classification:

F01D 5/20
F01D 5/18

Abstract:

A turbine blade tip includes a squealer tip wall having a forward surface that is continuous with an outer surface of an airfoil pressure sidewall. A plurality of cooling channels are provided spaced apart along a contour of the squealer tip wall. Each cooling channel includes: an inlet configured for receiving a coolant from airfoil internal cavity; an upstream section including a closed channel extending from the inlet to the forward surface of the squealer tip wall; and a downstream section including an open channel formed by a slot on the forward surface of the squealer tip wall. The slot extends radially outward in a downstream direction so as to guide the coolant along the forward surface toward a radially outermost tip of the squealer tip wall.

US Patent:

20180318919, Nov 8, 2018

Filed:

Jul 16, 2018

Appl. No.:

16/036611

Inventors:

Ching-Pang Lee - Cincinnati OH, US
Nan Jiang - Charlotte NC, US

International Classification:

B22D 25/02
F01D 5/18
B22C 9/04
B22C 9/10
B22C 7/02
B22D 29/00
B22C 21/14
B22C 9/24

Abstract:

A method of forming an airfoil (), including: abutting end faces () of cantilevered film hole protrusions () extending from a ceramic core () against an inner surface () of a wax die () to hold the ceramic core in a fixed positional relationship with the wax die; casting an airfoil including a superalloy around the ceramic core; and machining film cooling holes () in the airfoil after the casting step to form an pattern of film cooling holes comprising the film cooling holes formed by the machining step and the cast film cooling holes () formed by the film hole protrusions during the casting step.

US Patent:

20170081960, Mar 23, 2017

Filed:

Jun 5, 2014

Appl. No.:

15/303569

Inventors:

- Orlando FL, US
Ralph W. Matthews - Oviedo FL, US
Nan Jiang - Charlotte NC, US
Gm Salam Azad - Oviedo FL, US

International Classification:

F01D 5/18

Abstract:

A cooling system () positioned within a turbine airfoil () useable in a turbine engine and having cooling channels () positioned within a platform () of the turbine airfoil () with exhaust outlets () at the pressure and suction side edges () to prevent hot gas ingestion under the platform () is disclosed. The cooling channels () may be formed from main channels () extending from cooling fluid supply channels () aligned with the airfoil () and branch channels () extending between the main channels () and the pressure or suction side edges (). The cooling system () reduces the cooling surface area adjacent to the airfoil fillet () at the intersection () of the platform () and airfoil () and increases cooling surface area adjacent to the pressure side and suction side mate faces () as compared with conventional designs. Such configuration of the cooling system () yields a more uniform platform temperature distribution, colder and higher pressure cooling air for platform cooling and less manufacturing expense than conventional designs.

US Patent:

20170087630, Mar 30, 2017

Filed:

Jun 18, 2014

Appl. No.:

15/307073

Inventors:

- Orlando FL, US
Nan Jiang - Charlotte NC, US

International Classification:

B22D 25/02
B22C 9/10
F01D 5/18
B22C 9/04
B22D 29/00
B22C 7/02
B22C 9/24

Abstract:

A method of forming an airfoil (), including: abutting end faces () of cantilevered film hole protrusions () extending from a ceramic core () against an inner surface () of a wax die () to hold the ceramic core in a fixed positional relationship with the wax die; casting an airfoil including a superalloy around the ceramic core; and machining film cooling holes () in the airfoil after the casting step to form an pattern of film cooling holes comprising the film cooling holes formed by the machining step and the cast film cooling holes () formed by the film hole protrusions during the casting step.

US Patent:

20170058678, Mar 2, 2017

Filed:

Aug 31, 2015

Appl. No.:

14/840572

Inventors:

- Orlando FL, US
Nan Jiang - Charlotte NC, US
Jae Y. Um - Winter Garden FL, US
Harry Holloman - Cincinnati OH, US
Steven Koester - Sharonville OH, US

International Classification:

F01D 5/18
F01D 5/14

Abstract:

A turbine rotor blade includes at least two integrated cooling circuits that are formed within the blade that include a leading edge circuit having a first cavity and a second cavity and a trailing edge circuit that includes at least a third cavity located aft of the second cavity. The trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity. The last cavity is located along a trailing edge of the blade. A tip axial cooling channel connects to the first cavity of the leading edge circuit and the penultimate cavity of the trailing edge circuit. At least one crossover hole connects the penultimate cavity to the last cavity substantially near the tip end of the blade.