United Technologies Corporation

Raytheon Technologies Corporation

Rtx Corporation

Daniel A Snyder

Edwin Otero

Christopher Corcoran

Kyle C Lana

Kurt R Heinemann

Tracy A Propheter-Hinckley

Dominic J Mongillo, Jr

Brandon W Spangler

Scott Daniel Virkler

Jonathan Perry Sandoval

Jaimie Taraskevich

Di Wu

Wendell V Twelves

Jesse R Boyer

Evan J Butcher

Russell J Bergman

Glenn Levasseur

Jeffrey S Beattie

Lexia Kironn

Andrew S Aggarwala

Lane Mikal Thornton

Keith A Santeler

Mark A Boeke

Alex J Simpson

Timothy Charles Nash

Randall Joyner

David A Krause

Robert E Erickson

James Romanelli

Amanda Jean Learned

Jaimie Schweitzer

Perry Bowes

Evan Selin

Kun Tong

David J Yudichak

Jr Wendell V Twelves

Kyle Charles Lana

A qualification system for gas turbine engine components includes a computer system configured to receive a set of measured parameters for each gas turbine engine component in a plurality of substantially identical gas turbine engine components, and determine a variation model based on the set of measured parameters. The computer system includes at least one simulated engine model configured to determine a predicted operation of each gas turbine engine component in the plurality of substantially identical gas turbine engine components, a correlation system configured to correlate variations in the set of parameters for each of the gas turbine engine components with a set of the predicted operations of each gas turbine engine, thereby generating a predictive model based on the variations. The computer system also includes a qualification module configured to generate a qualification formula based on the predictive model. The qualification formula is configured to receive a set of measured parameters of an as-manufactured gas turbine engine component and determine when the as manufactured gas turbine engine component is qualified for inclusion in at least one engine.

Aircraft component qualification system and process

An exemplary method for qualifying a gas turbine engine component includes creating a first set of substantially identical gas turbine engine components via a uniform manufacturing procedure, determining a set of as-manufactured parameters of each gas turbine engine component in the first set of substantially identical gas turbine engine components, determining a variance model of the first set of substantially identical gas turbine engine components, and determining a plurality of predicted response models based at least in part on the variance model, each of the predicted response models corresponding to one of an engine type and an engine assembly, and each of the predicted response models being configured to determine a predicted response of including a gas turbine engine component from the first set of substantially identical gas turbine engine components in the corresponding one of the engine type and the engine assembly. The system then identifies as-manufactured parameters of a second engine component, and applies the as-manufactured parameters of the second engine component to each of the predicted response models, thereby generating a predicted response output from each of the predicted response models. An optimum predicted response from each of the generated predicted response models is identified and the engine type or engine assembly that corresponds with the optimum predicted response is associated with a unique part identifier of the second engine component.

Aircraft component qualification system and process for target based inventory qualification

A turbomachine airfoil element comprises an airfoil having: an inboard end; an outboard end; a leading edge; a trailing edge; a pressure side; and a suction side. A span between the inboard end and the outboard end is 1.35-1.65 inches. A chord length at 50% span is 1.20-1.60 inches. At least two of: a first mode resonance frequency is 2858&#xb1;10% Hz; a second mode resonance frequency is 4916&#xb1;10% Hz; a third mode resonance frequency is 7160&#xb1;10% Hz; a fourth mode resonance frequency is 10268&#xb1;10% Hz; a fifth mode resonance frequency is 14235&#xb1;10% Hz; and a sixth mode resonance frequency is 15088&#xb1;10% Hz.

Turbomachine blade

A method for qualifying a gas turbine engine component includes creating a first set of substantially identical gas turbine engine components via a uniform manufacturing procedure, determining a set of as-manufactured parameters of each gas turbine engine component in the first set, and determining a variance model of the first set. The variance model includes a representative parameter profile, which includes a plurality of component parameter profiles. The sum of each of the component parameter profiles is the representative parameter profile. The method also includes determining at least one predicted response models based at least in part on the variance model, identifying as-manufactured parameters of a second engine component, applying the as-manufactured parameters of the second engine component to the at least one predicted response models, thereby generating a predicted response output, and qualifying the second engine component for usage in at least one gas turbine engine corresponding to the at least one predicted response model.

Aircraft component qualification system and process including variation modeling

An example turbine includes a turbine disk and a turbine blade. The turbine disk includes a plurality of lugs and a plurality of slots. Each of the plurality of lugs is located between two of the plurality of slots. Each of the plurality of lugs includes a tab that extends radially outwardly from an end of the lug. The tab has a first section having an upper surface and a second section having an upper surface. The upper surface of the first section is inclined greater than the upper surface of the second section, and the upper surface of the first section of the tab defines a contour. The turbine blade includes a root received in one of the plurality of slots and a platform, and the platform has a lower surface defining a contour. When the turbine blade is received in one of the plurality of slots, the lower surface of the platform is located above a portion of the upper surface of the first section of the tab, and the contour of the upper surface of the first section of the tab matches the contour of the lower surface of the platform.

Turbine blade damper seal

A component according to an exemplary aspect of the present disclosure includes, among other things, a wall and a hollow vascular engineered lattice structure formed inside of the wall. The hollow vascular engineered lattice structure has an inlet hole and an outlet hole that communicate fluid into and out of the hollow vascular structure. The hollow vascular engineered lattice structure further has at least one resupply inlet hole between the inlet hole and the outlet hole with respect to a dimension of the component to communicate additional fluid into the hollow vascular engineered lattice structure.

Gas turbine engine component having engineered vascular structure

An airfoil stage of a turbine engine includes an upstream airfoil assembly, a downstream airfoil assembly in rotational relationship to the upstream airfoil assembly and a rim seal assembly integrated therebetween. The rim seal assembly may include a sloped downstream portion of a platform of the upstream airfoil assembly, an upstream segment of a platform of the downstream airfoil assembly and a nub that projects radially outward from the upstream segment. The downstream portion and the upstream segment are spaced from one-another defining a cooling cavity therebetween for the flow of cooling air. The portion and segment overlap axially such that the nub is axially aligned to the downstream portion for improved cooling effectiveness and a reduction of core airflow into the cooling cavity.

Low loss airfoil platform rim seal assembly

A segmented sideplate for use in a gas turbine engine is described. The segmented sideplate includes a first plate having a first circumferential edge configured to interface with a complementary circumferential edge. The segmented sideplate also includes a second plate having a second circumferential edge configured to interface with the first circumferential edge.

Turbine rotor segmented sideplates with anti-rotation

A flowpath component for a gas turbine engine includes a first platform including a vascular engineered lattice structure, a body extending from, and supported by the first platform. The body is configured to at least partially span a flowpath in an installed position and the vascular engineered lattice structure including at least one purge air inlet, and at least one spent air outlet.

Gas turbine engine component having vascular engineered lattice structure

A damper seal received in a cavity of a turbine blade located between a platform and a retention shelf damper seal according to an exemplary aspect of the present disclosure includes, among other things, a central body portion having a first end region, an opposing second end region, and a width. The damper seal further includes a first portion extending from the first end region of the central body portion, and a first end region of the first portion includes first outwardly extending tabs that define a first enlarged portion that has a first width greater than the width of the central body portion and a second portion extending from the opposing second end region of the central body portion.

A rotatable sealing structure for a gas turbine engine includes a first blade and a second blade. A seal is arranged between the blades and has a body that is configured for operative association with the first and second blades in a first orientation and in a second orientation to seal a gap defined between the blades.

Reversible blade rotor seal

A component according to an exemplary aspect of the present disclosure includes, among other things, a wall and a hollow vascular engineered lattice structure formed inside of the wall. The hollow vascular engineered lattice structure has an inlet hole and an outlet hole that communicate fluid into and out of the hollow vascular structure. The hollow vascular engineered lattice structure further has at least one resupply inlet hole between the inlet hole and the outlet hole to communicate additional fluid into the hollow vascular engineered lattice structure.

A seal and damper system is provided. The system comprises a blade assembly seal with a central portion defining a recess and a depression. The system also includes a damper comprising a center body and an arm extending from the center body and configured to engage the depression. A blade assembly is also provided comprising a blade platform, a seal configured to press against the blade platform, and a damper clipped onto the seal.

Seal and clip-on damper system and device

An attachment root of an airfoil is provided. The attachment root may comprise a serration profile, a symmetry plane bisecting the serration profile, and a line perpendicular to the symmetry plane. A first lobe of the serration profile may have a first contact face angled 135 degrees from the line. A second lobe of the serration profile may have a second contact face angled 135 degrees from the line. A third lobe of the serration profile may have a third contact face angled 135 degrees from the line.

Turbine airfoil attachment with serration profile

An exemplary gas turbine engine assembly includes a damping device having a first side and a second side facing away from the first side. The first side is configured to hold a seal when the second side engages an extension from a gas turbine engine component. The first side is further configured to engage the extension when the second side holds the seal.

Growing community of inventors

Manchester, CT, United States of America

Daniel A Snyder