General Electric Company

United States of America As Represented by the Secretary of the Air Force

Thomas G Wakeman

Harvey M Maclin

Ambrose A Hauser

James S Kelm

Michael Leon Barron

Arthur L Ludwig

Donald L Bellia

William James Strock

Thomas C Hermans

Alan Majors Walker

David Edward Bulman

Peter W Mueller

Ivan H Peterson

James S Hoelle

James W Brantley

Stephen Eugene Melton

Michael Todd Summers

Steven K Roggenkamp

Jan Christopher Schilling

Robert H Roth

Robert J Corsmeier

Dudley O Nash

Larry W Plemmons

Dean Thomas Lenahan

Lawrence Butler

Michael Ralph Storage

Mark Kevin Meyer

Gary Craig Wollenweber

Thomas P Joseph

Michael Harvey Schneider

Rolf R Hetico

Michael A Umney

Edward Charles Vickers

Gregg Hunt Wilson

James Edward Cencula

Petr Hora

Robert M Ausdenmoore

Thomas Anthony Hauer

Alan Luis Arvidson

Eric E Baehre

Neil Walker

Thomas S Clayton

Edwin K Miller

Walter D Howard

Frederick W Tegarden

Kenneth J Lenhart

Hahn M Spofford

David L Joyce

Brian J Humke

Conrad D Wagenknecht

Andrew P Elovic

Donald F Keck

Albert P Sterman

Louis Lievestro

David Byron Morris

Eugene J Antuna

Jesse J Williams

Richard W Brown

Steward Joseph Cline

Peter Hora

Jeffrey B Fravel

Alan B Corbeil

A thrust bearing support assembly includes a first substantially trapezoidally-shaped side, and a second substantially rectangularly-shaped side coupled to the first side. The thrust bearing support assembly also includes a third substantially trapezoidally-shaped side coupled to the second side, and a fourth substantially rectangularly-shaped side coupled to the first and third sides.

Thrust bearing support assembly for a gas turbine engine

A method of reducing gas turbine engine distortion includes coupling a core gas turbine engine to a power turbine, coupling the power turbine to a thrust bearing support assembly through a frusto-conically shaped structural support such that the net axial thrust of the coupled core gas turbine engine and power turbine reacts through the thrust bearing support assembly at engine centerline to facilitate minimizing distortion of the core gas turbine engine casing, and supporting the thrust bearing support assembly such that all of the remaining thrust generated by the core gas turbine engine and the power turbine is reacted to ground.

Gas turbine engine assembly and method of assembling same

A method enables a variable vane assembly for a gas turbine engine to be coupled to an engine casing. The variable vane assembly includes a bushing assembly and at least one variable vane that includes a platform and a vane stem. The method comprises coupling a first bushing to the engine casing in a press fit, coupling a second bushing to the variable vane, and coupling the variable vane to the engine casing such that at least a portion of the first bushing is between the engine casing and the second bushing, and such that at least a portion of the second bushing is between the first bushing and the vane stem.

Methods and apparatus for sealing gas turbine engine variable vane assemblies

An exemplary embodiment of the invention is directed to a method for determining a transfer function relating a critical to quality parameter to key parameters in a design for six sigma process. The method includes determining a dimensionless group containing a plurality of key parameters. The key parameters may include key control parameters or key noise parameters that have an affect on the critical to quality parameter. A transfer function relating the dimensionless group to the critical to quality parameter is then generated.

Method for generating transfer functions

A method of developing a confidence level in reliability or producibility prediction scores that includes generating data for a set of factors, inserting the data into scorecards and dashboards, and calculating an overall Z score and Z confidence range. The scorecards and dashboards are used to calculate the confidence level in reliability scores. Each reliability or producibility factor may be weighted as to its importance to the overall confidence level. Factors included in the method include accuracy of the critical to quality flowdown; completeness of the critical to quality flowdown; percentage of the design analyzed; comprehensiveness of the design analysis; risk assessment analysis; percentage of component. subassembly or product reuse from other projects; percentage of the design complete; verification of the process capability; plant integration; order to remittance process integration; extent of pilot run; and effectiveness of a test plan.

Method and system of confidence scoring

A hollow actuating ring and apparatus for actuating and/or vectoring the flaps of an aircraft gas turbine engine nozzle by transferring actuation loads from a small number of actuators to a greater number of pivotal divergent flap members of the nozzle. The actuating ring has a generally hollow annular structure including axially spaced apart coaxial forward and aft walls, struts axially extending between and structurally joining together the forward and aft walls, and a ring stiffening for decreasing the effective ring bending and torsional length of at least one of the forward and aft walls between the struts. Stiffening structure is provided for spreading out loads transferred from the struts to one of the forward and aft walls using structural braces extending diagonally between the struts and one of the forward and aft walls.

Hollow nozzle actuating ring

A clearance control arrangement for the rotor blade-to-stator shroud clearance in an aircraft engine or other turbo machine. The disclosed arrangements use large area piston-like structures located circumferential to the turbo machine and capable of operation from pressures developed within the machine. One disclosed arrangement incorporates static adjustment capability that can be disposed to provide symmetrical or asymmetrical clearances. The disclosed clearance control can be used as a fast-acting supplement or in lieu of other clearance control arrangements and is arranged to be actuated upon sensing the onset of rotor-to-stator collision pinch point conditions in the turbo machine.

Turbo machine shroud-to-rotor blade dynamic clearance control

A gas turbine engine thermal backbone deflection control apparatus to counter the effects of backbone bending due to nacelle aerodynamic, thrust, and maneuver loads by introducing a controlled backbone counter-bending moment by applying an uneven temperature or thermal field around the backbone to induce controlled compression and/or tensioning forces in the backbone or casing between axially spaced apart frames connected by the backbone. One embodiment of the invention provides for flowing thermal control air along an axially and circumferentially sector of the engine casing. The thermal backbone deflection control apparatus is controlled by the engines digital electronic control system using input signals generated by engine thrust load, inlet moment load sensors and/or blade tip clearance sensors.

Aircraft gas turbine engine backbone deflection thermal control

An expansion joint for connecting first and second fluid-carrying pipes comprising first and second annular couplings which are operably affixable to respective ones of the pipes. In a preferred embodiment the couplings include cylindrical bores of substantially the same diameter disposed relative to one another to form a passage for providing flow communication between the pipes. An annular bellows extends between the couplings for accommodating relative axial movement and angular misalignment between the pipes and is disposed radially outward of the passage to capture any fluid leaking between the interfacing surfaces of the couplings. A second cylindrical bore of a first one of the couplings engages an outer spherical surface of the second coupling to allow the second coupling to pivot in a manner which accommodates angular misalignment between the pipes. Means are provided for limiting the pivoting of the second coupling and for preventing the first coupling from rotating about a centerline of the joint, thereby avoiding unacceptable bending and torsional loading of the bellows. The expansion joint further includes means for limiting the axial compression and axial extension of the bellows.

Expansion joint for fluid piping with rotation prevention member

A tube coupling assembly for use in connecting first and second fluid-carrying tubes comprising an annular sleeve member and an annular ferrule which are operably affixable to respective ones of the tubes. A tube connector and a coupling nut threadingly engage one another so as to force first and second spherical surfaces of the sleeve member against generally frusto-conical surfaces of the ferrule and coupling nut respectively, thereby forming a fluid sealing means and accommodating angular misalignment between the tubes. An internal locking means is provided for preventing the disengagement of the sealing means. The coupling nut and tube connector each include wrenching surfaces which must be used to torque the assembly, with the applied torque and counter-torque avoiding torsional loading of the tubes.

Tube coupling assembly

An aircraft engine thrust mount includes a pair of spaced apart thrust links pivotally joined at one end to an engine first frame and at opposite ends to a lever. A platform is fixedly joined to an aircraft pylon and includes first and second spaced apart clevises and a center clevis therebetween. The center clevis is pivotally joined to a center of the lever for carrying thrust force from the first frame through both thrust links and to the platform and pylon during normal operation. A pair of pins pivotally join the thrust links to the lever and extend through apertures in the platform first and second clevises to define predetermined gaps therebetween for allowing the lever to rotate relative to the platform center clevis up to a predetermined angular rotation. In a failure mode of operation the pins contact the platform first and second clevises for transmitting thereto either tension or compression thrust loads for providing an alternate failsafe loadpath. In a preferred embodiment, the platform further includes an aft mount joined to an engine second frame spaced aft of the engine first frame.

Aircraft engine thrust mount

A system for changing pitch in aircraft propfans is disclosed, in which a motor, located at a stationary position, changes pitch of the blades by of a gear train. Three features of the system are that a feather brake can override the motor and drive the blades toward feather; a unison brake can prevent pitch from undergoing excursions toward fine pitch when propeller speed exceeds a limit; and the motor is non-rotating in the absence of pitch change. A pair of such systems can be used in a counterrotating propfan system, in order to independently change pitch of the propfans.

Aircraft pitch change mechanism

In one type of aircraft propulsion system, propeller blades are mounted on a ring which surrounds a turbine. An annular space exists between the turbine and the ring. If a propeller blade should break free, the unbalanced centrifugal load tends to deform the ring. The invention reduces the deformation, as by locating spacers between the turbine and the ring.

Fan blade protection system

A gas turbine engine has a core frame separable from a vane frame for removably attaching the vane frame to a core structure through a mid ring. The mid ring includes an annular forward panel and an annular aft panel, each of which has a radially outer flange and a radially inner flange. The inner flanges are welded to the core frame and the outer flanges are bolted to the vane frame. A plurality of circumferentially spaced stiffeners extend between the forward and aft panels, with each of the stiffeners including a pair of stringers extending diagonally across the stiffeners. Panels may be incorporated with the stringers between the triangular shaped areas adjacent the forward and aft panels of the stiffener. The bolted connection between the mid ring and the vane frame includes a plurality of tongue and groove joints formed in the flanges of the mid ring and vane frame. The tongue and groove joints provide additional radial support for the core engine within the vane frame. Preferably, the tongue and groove joints are formed with angular sides so that conformance between the joints can be achieved by simply tightening the bolts attaching the joints. In one embodiment, the tongue and groove joints may only be used at preselected circumferential positions with the remainder of the flange connections using overlapping lap joints.

Apparatus for removably attaching a core frame to a vane frame with a

A gas turbine engine combustor is provided, having a single wall sheet metal liner which is generally annular in shape which may be corrugated and a multi-hole film cooling means which includes at least one pattern of a great many small closely spaced sharply downstream angled film cooling holes disposed essentially along portion of the liner to be cooled. Another embodiment provided a corrugated aircraft engine afterburner sheet metal liner having at least one pattern of great many small closely spaced sharply downstream angled film cooling holes disposed essentially along the entire surface of the liner.

Gas turbine engine multi-hole film cooled combustor liner and method of

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Lawrenceburg, IN, United States of America

Thomas G Wakeman