Florida Turbine Technologies, Inc.

Siemens Aktiengesellschaft

Siemens Energy, Inc.

Christopher K Rawlings

Ching-Pang Lee

James Preston Downs

Kok-Mun Tham

Robert L Memmen

Russell B Jones

Vincent Paul Laurello

Gm Salam Azad

Darryl S Eng

Jae Y Um

Abdullatif M Chehab

Manjit Shivanand

John A Fedock

Friedrich T Rogers

Anthony J Malandra

Shantanu P Mhetras

Bryan C Bernier

Judson J Krueger

Kenneth K Landis

Thomas F Pechette

Phillip M Redman

Gloria E Goebel

Robert M Dysert

Billie E Sealey

Jose Paulino

Darryl Eng

Thomas Pechette

An apparatus and a process for experimentally determining a heat transfer coefficient of a surface which includes a duct having an outer passage for water flow and an inner passage for air flow, the two passages separated by a thin wall membrane such that the water flow establishes a datum temperature on the thin wall membrane, and where air flowing through the second passage and over the thin wall membrane surface which can include heat transfer enhancements features will be heated by the features, and where the heat transfer coefficient can be determined from the surface temperature of the thin wall membrane and the change in temperature of the air flow. The duct with the thin wall membrane and heat transfer enhancement features is produced using a plastic or metallic additive manufacture process for low cost and quit turnaround time.

Apparatus and process for determining a convective heat transfer coefficient between a moving fluid and a bounding surface

A shroud cooling system configured to cool a shroud adjacent to an airfoil within a gas turbine engine is disclosed. The turbine engine shroud may be formed from shroud segments that include a plurality of cooling air supply channels extending through a forward shroud support for impingement of cooling air onto an outer radial surface of the shroud segment with respect to the inner turbine section of the turbine engine. The channels may extend at various angles to increase cooling efficiency. The backside surface may also include various cooling enhancement components configured to assist in directing, dispersing, concentrating, or distributing cooling air impinged thereon from the channels to provide enhanced cooling at the backside surface. The shroud cooling system may be used to slow down the thermal response by isolating a turbine vane carrier from the cooling fluids while still providing efficient cooling to the shroud.

Shroud cooling system for shrouds adjacent to airfoils within gas turbine engines

A turbine blade airfoil () with a center of mass (ACM) that is laterally offset from the center of mass (PCM) of a platform () to which the airfoil is attached. Respective offsets (d, d) balance these centers of mass (ACM, PCM) about an attachment plane () of the blade root (), providing balanced centrifugal loading on opposite lobes () or other attachment surfaces of the root. The attachment plane () may be a plane of bilateral symmetry of the root, and/or it may include an attachment axis () that passes through the root center of mass (RCM) along a radius of rotation of the airfoil. The airfoil and platform centers of mass (ACM, PCM) may be dynamically balanced about the attachment axis () and/or the attachment plane ().

Turbine blade with laterally biased airfoil and platform centers of mass

A gas turbine having rotor discs (), a disc cavity () and a stator stage () extending to the disc cavity (). Seal housing flanges () extend from a seal housing () of the stator stage (). Rotor flanges () extend from a rotor disk (-). An inner rotor flange () and first seal housing flange () are inward from a second seal housing flange (). One rotor flange () is outward from the second seal housing flange (). The inner rotor flange () and first seal housing flange () extend toward one another to limit movement of main gas flow (). An inlet () injects air () between the outward rotor flange () and second seal housing flange ().

Flow discourager integrated turbine inter-stage U-ring

A turbine stator vane for an industrial engine, the vane having two impingement cooling inserts that produce a series of impingement cooling from the pressure side to the suction side of the vane walls. Each insert includes a spar with a row of alternating impingement cooling channels and return air channels extending in a radial direction. Impingement cooling plates cover the two sides of the insert and having rows of impingement cooling holes aligned with the impingement cooling channels and return air openings aligned with the return air channel.

Industrial stator vane with sequential impingement cooling inserts

A bi-metallic strip seal used to seal the gap between adjacent shroud segments in a gas turbine engine. The bi-metallic strip seal is formed from one or more layers of bi-metallic materials with different coefficients of thermal expansion. In a cold state the bi metallic strip seal will fit within the slot easily, while in the hot state the strip seal will expand to fit tightly within the slot and block leakage flow. The strip seal can be formed with the bi-metallic layers oriented similarly or oppositely in order to control the seal force and deflection properties. In one embodiment, the strip seal is formed from four bi-metallic layers in which the middle is bonded while the adjacent layers are free from shear.

Bi-metallic strip seal for a turbine shroud

A turbine airfoil with a relatively thick TBC applied over the airfoil surface. The airfoil has a surface contour on the trailing edge region with a forward end having a large amount of taper and a rearward end with less taper such that a substantially constant wall thickness is formed in the rearward end. The TBC is applied over the airfoil surface and tapers off at the trailing edge ends on the pressure side and the suction side walls to produce an ideal surface contour on the airfoil. In another embodiment, the airfoil surface includes a tapered section at the trailing edge region, and the TBC is applied over the taper so that an over-coating is formed. The TBC over-coating is then removed and a smooth and ideal surface contour is produced along the airfoil surface. Small raised bumps each having a height of the desired thickness of the TBC to be applied over the respective bump is used to control the finished thickness of the TBC.

Turbine airfoil with trailing edge

The present invention is a process for designing the internal cooling passages with these small scale features using a meshed solid model in a FEA program to perform the CFD analysis. Instead of modeling these small scale features (cooling holes, impingement holes, and turbulators) in the meshed FEA solid model, most of the small scale features are replaced by grid extraneous source terms in which the small scale feature is eliminated from the solid model and replaced with point source terms. The source terms can duplicate the effect of the feature within the analysis without requiring the complex analysis that such features would require in the analysis. Text files for each of the cooling holes, impingement holes and turbulators are inputted and then translated into grid extraneous source terms that include position, energy, and continuity. A conjugate CFD solver performs an analysis and produces new values for the thermal and boundary conditions.

Process for conjugate CFD analysis of a turbine airfoil

Growing community of inventors

Stuart, FL, United States of America

Christopher K Rawlings