Patents by Inventor Ching-Pang Lee

Ching-Pang Lee has filed for patents to protect the following inventions. This listing includes patent applications that are pending as well as patents that have already been granted by the United States Patent and Trademark Office (USPTO).

  • Patent number: 10787911
    Abstract: A gas turbine engine airfoil includes an outer wall including a suction side, a pressure side, a leading edge, and a trailing edge, the outer wall defining an interior chamber of the airfoil. The airfoil further includes cooling structure provided in the interior chamber. The cooling structure defines an interior cooling cavity and includes a plurality of cooling fluid outlet holes, at least one of which is in communication with a pressure side cooling circuit and at least one of which is in communication with a suction side cooling circuit. At least one of the pressure and suction side cooling circuits includes: a plurality of rows of airfoils, wherein radially adjacent airfoils within a row define segments of cooling channels. Outlets of the segments in one row align aerodynamically with inlets of segments in an adjacent downstream row such that the cooling channels have a serpentine shape.
    Type: Grant
    Filed: June 11, 2018
    Date of Patent: September 29, 2020
    Assignees: SIEMENS ENERGY, INC., MIKRO SYSTEMS, INC.
    Inventors: Ching-Pang Lee, Benjamin Heneveld
  • Publication number: 20200291787
    Abstract: A turbine airfoil (10) includes a trailing edge coolant cavity (41f) located in an airfoil interior (11) between a pressure sidewall (14) and a suction sidewall (16). The trailing edge coolant cavity (41f) is positioned adjacent to a trailing edge (20) of the turbine airfoil (10) and is in fluid communication with a plurality of coolant exit slots (28) positioned along the trailing edge (20). At least one framing passage (70, 80) is formed at a span-wise end of the trailing edge coolant cavity (41f). The airfoil (10) further includes framing features (72A-B, 82A-B) located in the framing passage (70, 80). The framing features are configured as ribs (72A-B, 82A-B) protruding from the pressure sidewall (14) and/or the suction sidewall (16). The ribs (72A-B, 82A-B) extend partially between the pressure sidewall (14) and the suction sidewall (16).
    Type: Application
    Filed: October 24, 2016
    Publication date: September 17, 2020
    Inventor: Ching-Pang Lee
  • Patent number: 10697306
    Abstract: A core structure (10) includes a first core element (16) including a leading edge section (30), a tip section (32), and a turn section (34) joining the leading edge and tip sections (30, 32). The first core element (16) is adapted to be used to form a leading edge cooling circuit (102) in a gas turbine engine airfoil (100). The leading edge cooling circuit (102) includes a cooling fluid passage (104) having a leading edge portion (106) formed by the first core element leading edge section (30), a tip portion (108) formed by the first core element tip section (32), and a turn portion (110) formed by the first core element turn section (34). Each of the leading edge portion (106), the tip portion (108), and the turn portion (110) of the cooling fluid passage (104) are formed concurrently in the airfoil (100) by the first core element (16).
    Type: Grant
    Filed: September 18, 2014
    Date of Patent: June 30, 2020
    Assignee: SIEMENS AKTIENGESELLSCHAFT
    Inventors: Ching-Pang Lee, Jae Y. Um, Gerald L. Hillier, Wayne J. McDonald, Erik Johnson, Anthony Waywood, Eric Schroeder, Zhengxiang Pu
  • Publication number: 20200157951
    Abstract: A blade (10) for a turbine engine that includes an internal cooling system (56) formed from at least one cavity (58) positioned within a generally elongated airfoil (12). A squealer tip (36) and at least one densified oxide dispersion strengthened layer (38) extend radially from a radially outer tip cap (70) of the blade (10), the tip cap (70) having a tip cap upper surface (50).
    Type: Application
    Filed: May 1, 2018
    Publication date: May 21, 2020
    Inventors: Kok-Mun Tham, Ching-Pang Lee, Li Shing Wong, Sin Chien Siw
  • Publication number: 20200095869
    Abstract: An integrated airfoil and platform cooling system (30) for a turbine rotor blade (10) includes an inlet (38, 48) located at the root (24) for receiving a supply of a coolant (K), and at least one cooling leg (32a, 32c, 42a, 42c) fluidly connected to the inlet (38, 48) and configured for conducting the coolant (K) in a radially outboard direction. The cooling leg (32a, 32c, 42a, 42c) is defined at least partially by a span-wise extending internal cavity (26) within a blade airfoil (12). An entrance of the cooling leg (32a, 32c, 42a, 42c) comprises a flow passage (92, 102) that extends radially outboard and laterally into a blade platform (50), so as to direct a radially outboard flowing coolant (K) to impinge on an inner side (60) of a radially outer surface (52) of the blade platform (50), before leading the coolant (K) into the cooling leg (32a, 32c, 42a, 42c).
    Type: Application
    Filed: March 20, 2018
    Publication date: March 26, 2020
    Inventors: Ching-Pang Lee, Anthony Waywood, Steven Koester
  • Publication number: 20200032659
    Abstract: A bladed rotor system (10) for a turbomachine includes a row of blades (14) mounted on a rotor disc (12). Each blade (14) includes an airfoil (16) with a mid-span snubber (30). The row of blades (14) includes first (H) and second (L) sets of blades (14). The blades (14) of the second set (L) are distinguished from the blades (14) of the first set (H) by a geometry of the snubber (30) that is unique to the respective set (H, L). In particular, the snubbers (30) of the second set (L) are attached to the respective airfoils (16) at a different span-wise height than that of the snubbers (30) of the first set (H). Thereby the natural frequency of a blade (14) in the second set (L) differs from the natural frequency of a blade (14) in the first set (H) by a predetermined amount. Blades (14) of the first set (H) and the second set (L) are positioned alternately in the row of blades (14), to provide a frequency mistuning to stabilize flutter of the blades (14).
    Type: Application
    Filed: February 26, 2018
    Publication date: January 30, 2020
    Inventors: Yuekun Zhou, Ching-Pang Lee
  • Patent number: 10526900
    Abstract: A turbine component including a shrouded airfoil with a flow conditioner configured to direct leakage flow and coolant to be aligned with main hot gas flow is provided. The flow conditioner is positioned on a shroud base radially adjacent to the tip of the airfoil and includes a ramped radially outer surface positioned further radially inward than a radially outer surface of the shroud base. The ramped radially outer surface extends from a first edge to a second edge in a direction generally from the suction side to the pressure side of the airfoil, such that the first edge is positioned further radially inward than the second edge. Multiple coolant ejection holes are positioned on the ramped radially outer surface. The coolant ejection holes are connected fluidically to an interior of the airfoil.
    Type: Grant
    Filed: June 29, 2015
    Date of Patent: January 7, 2020
    Assignee: SIEMENS AKTIENGESELLSCHAFT
    Inventors: Kok-Mun Tham, Ching-Pang Lee, Eric Chen, Steven Koester
  • Patent number: 10393132
    Abstract: A compressor (10) configured for use in a gas turbine engine (12) and having a rotor assembly (14) with a pumping system (16) positioned on a rotor drum (18) to counteract reverse leakage flow at a gap (20) formed between one or more stator vane tips (22) and a radially outer surface (24) of the rotor drum (18). The pumping system (16) may be from pumping components (26) positioned radially inward of one or more stator vane tips (22) to reduce, if not completely eliminate, reverse leakage flow at the stator vane tips (22). In at least one embodiment, the pumping component (26) may be formed from one or more cutouts (28) in the outer surface (24) of the rotor drum (18). In another embodiment, the pumping component (26) may be formed from at least one pumping fin (30) extending from the radially outer surface (24) of the rotor drum (18). In at least one embodiment, rows (32) of pumping components (26) may be aligned with rows (34) of stator vanes (36) within the compressor (10).
    Type: Grant
    Filed: August 8, 2014
    Date of Patent: August 27, 2019
    Assignee: SIEMENS AKTIENGESELLSCHAFT
    Inventors: Ching-Pang Lee, Kok-Mun Tham
  • Patent number: 10221716
    Abstract: Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with ridges projecting from the abradable surface, separated by grooves. The ridges have one or both sidewalls inclined against the opposing turbine blade tip rotational direction for redirecting and/or dissipating blade tip gap leakage airflow energy. In some embodiments the ridge tip and/or groove base have inclined profiles for redirecting airflow leakage away from the blade tip gap. In some embodiments, the inclined ridge tip profile provides a progressive wear zone that increases abradable surface area as the inclined ridge is abraded by the rotating blade tip.
    Type: Grant
    Filed: February 18, 2015
    Date of Patent: March 5, 2019
    Assignee: SIEMENS AKTIENGESELLSCHAFT
    Inventors: Kok-Mun Tham, Ching-Pang Lee
  • Patent number: 10196906
    Abstract: A turbine blade including a pressure sidewall (24) and a suction sidewall (26), and at least one partition rib (34) extends between the pressure and suction sidewalls (24, 26) to define a serpentine cooling path (35) having adjacent cooling channels (36a, 36b, 36c) extending in the spanwise direction (S) within the airfoil (12). A flow turning guide structure (50) extends around an end of the at least one partition rib (34) and includes a first element (52) extending from the pressure sidewall (24) to a lateral location in the cooling path between the pressure and suction sidewalls (24, 26), a second element (54) extending from the suction sidewall (26) to the lateral location in the cooling path between the pressure and suction sidewalls (24, 26). The first and second elements (52, 54) include respective distal edges (52d, 54d) that laterally overlap each other at the lateral location.
    Type: Grant
    Filed: March 17, 2015
    Date of Patent: February 5, 2019
    Assignee: SIEMENS ENERGY, INC.
    Inventors: Ching-Pang Lee, Yuekun Zhou, Adhlere Coffy, Steven Koester
  • Patent number: 10189082
    Abstract: Turbine and compressor casing abradable components for turbine engines include abradable surfaces with a zonal system of forward (zone A) and rear or aft sections (zone B) surface features. The zone A surface profile comprises an array pattern of non-directional depression dimples, or upwardly projecting dimples, or both, in the abradable surface. The dimpled forward zone A surface features reduce surface solidity in a controlled manner, to help increase abradability during blade tip rubbing incidents, yet they provide sufficient material to resist incoming hot working fluid erosion of the abradable surface. In addition, the dimples provide generic forward section aerodynamic profiling to the abradable surface, compatible with different blade airfoil-camber profiles. The aft zone B surface features comprise an array pattern of ridges and grooves.
    Type: Grant
    Filed: December 9, 2015
    Date of Patent: January 29, 2019
    Assignee: SIEMENS AKTIENGESELLSCHAFT
    Inventors: Kok-Mun Tham, Ching-Pang Lee
  • Patent number: 10190435
    Abstract: Turbine and compressor casing abradable component embodiments for turbine engines vary localized porosity or abradability through use of holes or dimple depressions of desired polygonal profiles that are formed into the surface of otherwise monolithic abradable surfaces or rib structures. Abradable porosity within a rib is varied locally by changing any one or more of hole/depression depth, diameter, array pitch density, and/or volume. In various embodiments, localized porosity increases and corresponding abradability increases axially from the upstream or forward axial end of the abradable surface to the downstream or aft end of the surface. In this way, the forward axial end of the abradable surface has less porosity to counter hot working gas erosion of the surface, while the more aft portions of the abradable surface accommodate blade cutting and incursion with lower likelihood of blade tip wear.
    Type: Grant
    Filed: December 9, 2015
    Date of Patent: January 29, 2019
    Assignee: SIEMENS AKTIENGESELLSCHAFT
    Inventors: Ching-Pang Lee, Ramesh Subramanian, Kok-Mun Tham
  • Publication number: 20190003319
    Abstract: A gas turbine engine airfoil includes an outer wall including a suction side, a pressure side, a leading edge, and a trailing edge, the outer wall defining an interior chamber of the airfoil. The airfoil further includes cooling structure provided in the interior chamber. The cooling structure defines an interior cooling cavity and includes a plurality of cooling fluid outlet holes, at least one of which is in communication with a pressure side cooling circuit and at least one of which is in communication with a suction side cooling circuit. At least one of the pressure and suction side cooling circuits includes: a plurality of rows of airfoils, wherein radially adjacent airfoils within a row define segments of cooling channels. Outlets of the segments in one row align aerodynamically with inlets of segments in an adjacent downstream row such that the cooling channels have a serpentine shape.
    Type: Application
    Filed: June 11, 2018
    Publication date: January 3, 2019
    Inventors: Ching-Pang Lee, Benjamin Heneveld
  • Publication number: 20180318919
    Abstract: A method of forming an airfoil (12), including: abutting end faces (72) of cantilevered film hole protrusions (64) extending from a ceramic core (50) against an inner surface (80) of a wax die (68) 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 (34) 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 (102) formed by the film hole protrusions during the casting step.
    Type: Application
    Filed: July 16, 2018
    Publication date: November 8, 2018
    Inventors: Ching-Pang Lee, Nan Jiang
  • Publication number: 20180298763
    Abstract: The present disclosure provides a turbine blade (12) comprising a leading edge cooling circuit (30), a trailing edge cooling circuit (34), a mid-section cooling circuit (32) comprising a first channel (32a), an intermediate channel (32b), and a final channel (32c), and an axial tip cooling circuit (56). The leading edge, mid-section, and trailing edge cooling circuits (30, 32, 34) each receive a cooling airflow (CF) from a cooling air supply. A radially outer portion of each of the leading edge and mid-section cooling circuits (30, 32) further comprises at least one outlet (62, 64) in fluid communication with the axial tip cooling circuit (56) such that substantially all of a leading edge cooling airflow (LEF) exiting the leading edge cooling circuit (30) and substantially all of a mid-section cooling airflow (MSF) exiting the mid-section cooling circuit (32) is directed to the axial tip cooling circuit (56).
    Type: Application
    Filed: November 11, 2014
    Publication date: October 18, 2018
    Inventors: Ching-Pang Lee, Jae Y. Um, Gerald L. Hillier, Eric Schroeder, Erik Johnson, Dustin Muller
  • Publication number: 20180266254
    Abstract: A trailing edge cooling feature for a turbine airfoil (10) includes a plurality of pins (22a-l) positioned in an airfoil interior (11) toward the trailing edge 20), each extending from the pressure side (14) to the suction side (16) and further being elongated in a radial direction (R). The pins (22a-l) are arranged in multiple radial rows (A-L) spaced along the chordal axis (30), with the pins (22a-l) in each row (A-L) being interspaced to define coolant passages (24a-l) therebetween. A row of radially spaced apart partition walls (26) are positioned aft of the pins (22a-l). Each partition wall (26) extends from the pressure side (14) to the suction side (16) and is elongated in a generally axial direction, extending along the chordal axis (30) to terminate at the trailing edge (20).
    Type: Application
    Filed: October 30, 2015
    Publication date: September 20, 2018
    Inventor: Ching-Pang Lee
  • Patent number: 10060270
    Abstract: An airfoil (10) is disclosed for a gas turbine engine in which the airfoil (10) includes an internal cooling system (14) with one or more converging-diverging exit slots (20) configured to increase the effectiveness of the cooling system (14) at the trailing edge (34) of the airfoil (10) by increasing the contact of cooling fluids with internal surfaces (24, 30) of the pressure and suction sides (36, 38) of the airfoil (10). In at least one embodiment, the trailing edge cooling channel (18) may include one or more converging-diverging exit slots (20) to further pressurize the trailing edge cooling channel (18) and may be formed by a first and second ribs (80, 82) extending between an outer walls (13, 12) forming the pressure and suction sides (36, 38).
    Type: Grant
    Filed: March 17, 2015
    Date of Patent: August 28, 2018
    Assignee: SIEMENS ENERGY, INC.
    Inventors: Ching-Pang Lee, Caleb Myers, Erik Johnson, Steven Koester
  • Patent number: 10053993
    Abstract: A shrouded turbine airfoil (10) with a leakage flow conditioner (12) configured to direct leakage flow to be aligned with main hot gas flow is disclosed. The leakage flow conditioner (12) may be positioned on a radially outer surface (18) of an outer shroud base (20) of the outer shroud (22) on a tip (24) of an airfoil (10). The leakage flow conditioner (12) may include a radially outer surface (28) that is positioned further radially inward than the radially outer surface (18) of the outer shroud base (20) creating a radially outward extending wall surface (30) that serves to redirect leakage flow. In at least one embodiment, the radially outward extending wall surface (30) may be aligned with a pressure side (38) of the shrouded turbine airfoil (10) to increase the efficiency of a turbine engine by redirecting leakage flow to be aligned with main hot gas flow to reduce aerodynamic loss upon re-introduction to the main gas flow.
    Type: Grant
    Filed: March 17, 2015
    Date of Patent: August 21, 2018
    Assignee: SIEMENS ENERGY, INC.
    Inventors: Kok-Mun Tham, Ching-Pang Lee, Li Shing Wong, Andrew S. Lohaus, Farzad Taremi
  • Patent number: 10022790
    Abstract: A method of forming an airfoil (12), including: abutting end faces (72) of cantilevered film hole protrusions (64) extending from a ceramic core (50) against an inner surface (80) of a wax die (68) 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 (34) 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 (102) formed by the film hole protrusions during the casting step.
    Type: Grant
    Filed: June 18, 2014
    Date of Patent: July 17, 2018
    Assignees: SIEMENS AKTIENGESELLSCHAFT, MIKRO SYSTEMS, INC.
    Inventors: Ching-Pang Lee, Nan Jiang
  • Publication number: 20180179900
    Abstract: A turbine component (10) including a shrouded airfoil (32) with a flow conditioner (70, 70a, 70b) configured to direct leakage flow and coolant to be aligned with main hot gas flow is provided. The flow conditioner (70, 70a, 70b) is positioned on a shroud base (20) radially adjacent to the tip of the airfoil and includes a ramped radially outer surface (72) positioned further radially inward than a radially outer surface (25) of the shroud base (20). The ramped radially outer surface (72) extends from a first edge (74) to a second edge (76) in a direction generally from the suction side (40) to the pressure side (38) of the airfoil (32), such that the first edge (74) is positioned further radially inward than the second edge (76). Multiple coolant ejection holes (80) are positioned on the ramped radially outer surface (72). The coolant ejection holes (80) are connected fluidically to an interior (81) of the airfoil (32).
    Type: Application
    Filed: June 29, 2015
    Publication date: June 28, 2018
    Inventors: Kok-Mun Tham, Ching-Pang Lee, Eric Chen, Steven Koester