Siemens Energy, Inc.

Mikro Systems, Inc.

Siemens Aktiengesellschaft

Siemens Energy Global Gmbh & Co. Kg

Rtx Corporation

Nan Jiang

Ching-Pang Lee

John Joseph Marra

George P Liang

Ronald J Rudolph

Stephen Williamson

Steven Koester

Jae Y Um

Samuel R Miller, Jr

Zhihong Gao

Friedrich T Rogers

John P Dalton

Harry Holloman

John Harrington

Hubert Paprotna

Lieke Wang

Shantanu P Mhetras

Mehmet Gungor

A turbine blade includes a blade platform, a blade airfoil that extends from the blade platform toward a blade tip, the blade airfoil having a pressure side wall and a suction side wall joined at a blade leading edge and a blade trailing edge, a tip cap surface defined at an end of the blade airfoil facing the blade tip, a squealer tip wall that extends along a portion of the pressure side wall and a portion of the suction side wall from the tip cap surface to the blade tip and from the blade leading edge toward the blade trailing edge, and a chamfered surface formed as a part of the squealer tip wall at a region that is adjacent to the blade trailing edge.

Turbine blade squealer tip wall with chamfered surface

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.

Turbine blade and corresponding method of servicing

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.

Turbine airfoil cooling system with leading edge impingement cooling system turbine blade investment casting using film hole protrusions for integral wall thickness control

A cooling system for a turbine airfoil of a gas turbine engine is disclosed, whereby the cooling system includes spanwise extending midflow blockers positioned within one or more cooling channels to maintain an internal through flow channel Mach number. One or more cooling channels may have a larger cross-sectional area proximate to an outer end of the airfoil than at an inner end. One or more cooling channels include midflow blockers extending into the cooling channel. In at least one embodiment, the midflow blocker may extend radially inward from the outer end of the airfoil. The midflow blocker may limit movement of cooling fluid from the pressure side to the suction side or vice versa. The midflow blocker may increase in size moving radially outward as the cross-sectional area of the cooling channel increases as well. Such configuration keeps the internal through flow channel Mach number within design limits.

Turbine airfoil cooling system with spanwise extending fins

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.

Integrated circuit cooled turbine blade

A continuous serpentine cooling circuit forming a progression of radial passages (A,A) between pressure and suction side walls () in a MID region of a turbine airfoil (). The circuit progresses first axially, then tangentially, ending in a last radial passage (A) adjacent to the suction side () and not adjacent to the pressure side (). The passages of the axial progression () may be adjacent to both the pressure and suction side walls of the airfoil. The next to last radial passage (A) may be adjacent to the pressure side wall and not adjacent to the suction side wall. The last two radial passages (A,A) may be longer along the pressure and suction side walls respectively than they are in a width direction, providing increased direct cooling surface area on the interiors of these hot walls.

Integrated axial and tangential serpentine cooling circuit in a turbine airfoil

Multi-scale turbulation features, including first turbulators () on a cooling surface (), and smaller turbulators () on the first turbulators. The first turbulators may be formed between larger turbulators (). The first turbulators may be alternating ridges () and valleys (). The smaller turbulators may be concave surface features such as dimples () and grooves (), and/or convex surface features such as bumps () and smaller ridges (). An embodiment with convex turbulators () in the valleys () and concave turbulators () on the ridges () increases the cooling surface area, reduces boundary layer separation, avoids coolant shadowing and stagnation, and reduces component mass.

Compound cooling flow turbulator for turbine component

A ring segment for a gas turbine engine includes an outer panel defining a structural body for the ring segment. An outer side of an inner panel is attached to an inner side of the outer panel at an interface, and an inner side of the inner panel defines a portion of a hot gas path through the gas turbine engine. An outer side of the outer panel, opposite from the interface, is in communication with a source of cooling air. A plurality of impingement holes extend through the outer panel from the outer side to the inner side of the outer panel for directing impingement air to the outer side of the inner panel. The outer and inner panels define a plurality of flow channels at the interface for effecting convective cooling of the outer panel along the flow channels between the outer and inner panels.

Ring segment with impingement and convective cooling

A cooling system having one or more pin fins with reduced mass for a gas turbine engine is disclosed. The cooling system may include one or more first surfaces defining at least a portion of the cooling system. The pin fin may extend from the surface defining the cooling system and may have a noncircular cross-section taken generally parallel to the surface and at least part of an outer surface of the cross-section forms at least a quartercircle. A downstream side of the pin fin may have a cavity to reduce mass, thereby creating a more efficient turbine airfoil.

Cooling system having reduced mass pin fins for components in a gas turbine engine

A turbine vane for a gas turbine engine with an internal cooling system formed from a serpentine cooling channel with one or more flow blocking ribs is disclosed. The serpentine cooling channels may be configured to receive cooling fluids from internal cooling fluids supply channels. The serpentine cooling channels may include flow blocking ribs to form concurrent flow channels to reduce the cross-sectional area within the midchord region of the airfoil to maintain the internal through flow channel Mach number. The flow blocking ribs may include slots therein and may have any appropriate configuration. In at least one embodiment, the flow blocking ribs may be have a nonuniform taper or a uniformed taper.

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Jupiter, FL, United States of America

Nan Jiang