Siemens Aktiengesellschaft

Siemens Energy, Inc.

Erik Johnson

Ching-Pang Lee

Eric Joseph Schroeder

Kok-Mun Tham

Gm Salam Azad

Steven Koester

Nicholas F Martin, Jr

Jae Y Um

Zhihong Gao

Manjit Shivanand

Zhengxiang Pu

Wayne J McDonald

Anthony Waywood

Gerald L Hillier

Caleb Myers

Dustin Jaime Muller

Steven Coppess

Kahwai G Muriithi

Steven Koester, Jr

A core structure () includes a first core element () including a leading edge section (), a tip section (), and a turn section () joining the leading edge and tip sections (). The first core element () is adapted to be used to form a leading edge cooling circuit () in a gas turbine engine airfoil (). The leading edge cooling circuit () includes a cooling fluid passage () having a leading edge portion () formed by the first core element leading edge section (), a tip portion () formed by the first core element tip section (), and a turn portion () formed by the first core element turn section (). Each of the leading edge portion (), the tip portion (), and the turn portion () of the cooling fluid passage () are formed concurrently in the airfoil () by the first core element ().

Gas turbine airfoil including integrated leading edge and tip cooling fluid passage and core structure used for forming such an airfoil

An airfoil () is disclosed for a gas turbine engine in which the airfoil () includes an internal cooling system () with one or more converging-diverging exit slots () configured to increase the effectiveness of the cooling system () at the trailing edge () of the airfoil () by increasing the contact of cooling fluids with internal surfaces () of the pressure and suction sides () of the airfoil (). In at least one embodiment, the trailing edge cooling channel () may include one or more converging-diverging exit slots () to further pressurize the trailing edge cooling channel () and may be formed by a first and second ribs () extending between an outer walls () forming the pressure and suction sides (). The converging-diverging exit slot () may be formed from a first converging section () having an inlet () with a larger cross-sectional area than an outlet () and is formed from a second diverging section () having an inlet () with a smaller cross-sectional area than an outlet ().

Internal cooling system with converging-diverging exit slots in trailing edge cooling channel for an airfoil in a turbine engine

Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with asymmetric forward and aft ridge surface area density. The forward ridges have greater surface area density than the aft ridges to compensate for greater ridge erosion in the forward zone during engine operation and reduce blade tip wear in the aft zone. Some abradable component embodiments increase forward zone ridge surface area density by incorporating wider ridges than those in the aft zone.

Turbine abradable layer with compound angle, asymmetric surface area ridge and groove pattern

Turbine and compressor casing abradable component embodiments for turbine engines, with composite, non-inflected, bi-angle, 'hockey stick' like pattern abradable surface ridges and grooves. Some embodiments include distinct forward upstream and aft downstream composite multi orientation groove and vertically projecting ridges planform patterns, to reduce, redirect and/or block blade tip airflow leakage downstream into the grooves rather than from turbine blade airfoil high to low pressure sides. In some embodiments the grooves are split or divided into multiple sections to interrupt flow traveling inside the groove and cause a local pressurization that reduces tip leakage flow. Some ridge or rib embodiments also have first lower and second upper wear zones. The lower zone optimizes engine airflow characteristics while the upper zone is optimized to minimize blade tip gap and wear by being more easily abradable than the lower zone.

Turbine abradable layer with composite non-inflected bi-angle ridges and grooves

A seal assembly between a disc cavity and a hot gas path in a gas turbine engine includes a stationary vane assembly and a rotating blade assembly axially upstream from the vane assembly. A platform of the blade assembly has a radially outwardly facing first surface, an axially downstream facing second surface defining an aft plane, and a plurality of grooves extending into the second surface such that the grooves are recessed from the aft plane The grooves are arranged such that a circumferential space is defined between adjacent grooves During operation of the engine, the grooves impart a circumferential velocity component to purge air flowing out of a disc cavity through the grooves to guide the purge air toward a hot gas path such that the purge air flows in a desired direction with reference to a direction of hot gas flow through the hot gas path.

Seal assembly including grooves in an aft facing side of a platform in a gas turbine engine

Growing community of inventors

Cedar Park, TX, United States of America

Erik Johnson