Abstract: A method for assembling a gas turbine engine combustor facilitates reducing costs and time required for assembly. The combustor includes a spectacle plate, a plurality of swirlers, and a plurality of deflector plates. The method includes coupling an assembly fixture to at least one swirler, coupling the assembly fixture to the spectacle plate such that the swirler is maintained in alignment with respect to the spectacle plate during assembly of the combustor, and attaching the swirler to the spectacle plate.

Abstract: Methods for error proofing, which also improve material, or information, flow through a process, are described. In one embodiment, the method includes the steps of identifying a process responsible for at least one error, process mapping the identified process, identifying at least one step in the identified process at which scrap and/or nonconformance occurred, and razing the identified process. The step of razing, in the one embodiment, is performed by determining whether a plurality of errors throughout the identified process occurs, and if a plurality of errors are identified throughout the identified process, determining whether a different process can be substituted for the identified process, and if the different process can be substituted for the identified process, then substituting the different process for the identified process.

Abstract: A combustor for a gas turbine engine that facilitates reducing combustor chamber dump pressure losses to improve combustor and engine performance is described. The combustor includes a diffuser that diffuses airflow directed into the combustor. The diffuser includes an outer wall, an inner wall, and a plurality of splitter vanes between the outer wall and the inner wall. The splitter vanes are spaced radially apart, and each is spaced radially from a respective diffuser wall such that an outer passage, an inner passage, and a central passage are defined by the splitter vanes.

Abstract: A coupling for coupling a component to a hanger including a first member and a second member. The coupling includes a first member and a second member. The first member includes a pair of extending portions and a pair of extending tabs. The second member includes a pair of circumferential openings and a pair of opposed notches. The openings and notches receive the extending portions and tabs. The component is supportable from the hanger and is substantially restricted from independent movement relative to the hanger when the first member is connected to the second member.

Abstract: A method for fabricating a rotor blade for a gas turbine engine facilitates reducing operating temperatures of a tip portion of the rotor blade. The method comprises forming an airfoil including a first sidewall and a second sidewall connected at a leading edge and a trailing edge to define a cavity therein, wherein the first and second sidewalls extend radially between a rotor blade root and a rotor blade tip plate, and forming a first tip wall extending from the rotor blade tip plate along the first sidewall, such that at least a portion of the first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. A second tip wall is formed to extend from the rotor blade tip plate along the second sidewall.

Abstract: A gas turbine engine includes a non-rotatable member that includes a honeycomb seal that reduces wear to rotor seal teeth disposed within the gas turbine engine. The gas turbine engine also includes a rotatable annular member including a sealing assembly disposed between rotor and stator components. The rotatable annular member includes seal teeth that extend radially outward from the rotatable annular member. The stator components include a non-rotatable member that includes a honeycomb seal that extends radially inward. The honeycomb seal is fabricated from a material that has a melting temperature less than approximately 2000° F.

Abstract: A method for assembling a turbine nozzle for a gas turbine engine facilitates improving cooling efficiency of the turbine nozzle. The method includes providing a hollow doublet including a leading airfoil and a trailing airfoil coupled by at least one platform, wherein each airfoil includes a first sidewall and a second sidewall that extend between a respective leading and trailing edge. The method also includes inserting an insert into at least one of the airfoils, wherein the insert includes a first sidewall including a first plurality of cooling openings that extending therethrough, and a second sidewall including a second plurality of cooling openings extending therethrough, and wherein the first plurality of cooling openings facilitate more cooling of the airfoil than the second plurality of cooling openings.

Abstract: A web-based system standardizes data from a plurality of different programs to facilitate design review comparisons in a cost-effective and reliable manner. The system includes a client system including a browser, a data storage device for storing information, and a server system coupled to the client system and the database. The server system is configured to store data in the data storage device in a standardized format, generate characteristics unique to the data input using pre-determined calculations, compare the characteristics to pre-determined target values, quantify shifts of the characteristics from respective target values, and generate standardized graphical representations of the data using based on the generated characteristics.

Abstract: A method for controlling an aircraft engine using a control system enables an engine to be operated at low power while hail is ingested. The control system includes a processor coupled to the aircraft engine. The method comprises receiving a plurality of sensed inputs from the aircraft engine, receiving current engine environmental operating conditions, determining a reference value for a controlled variable from a fuel flow schedule, generating a fuel-air-ratio for the reference value based on current engine environmental operating conditions, and enriching combustor fuel staging in response to the fuel-air-ratio.

Abstract: A gas turbine engine including in-line intercooling wherein compressor intercooling is achieved without removing the compressor main flow airstream from the compressor flowpath is described. In an exemplary embodiment, a gas turbine engine suitable for use in connection with in-line intercooling includes a low pressure compressor, a high pressure compressor, and a combustor. The engine also includes a high pressure turbine, a low pressure turbine, and a power turbine. For intercooling, fins are located in an exterior surface of the compressor struts in the compressor flowpath between the outlet of the low pressure compressor and the inlet of the high pressure compressor. Coolant flowpaths are provided in the compressor struts, and such flowpaths are in flow communication with a heat exchanger. In operation, air flows through the low pressure compressor, and compressed air is supplied from the low pressure compressor to the high pressure compressor.

Abstract: An airfoil for a gas turbine engine including an inflection that facilitates enhancing film cooling of the airfoil, without adversely affecting aerodynamic efficiency of airfoil is described. The airfoil includes a generally concave first sidewall and a generally convex second sidewall joined at a leading edge and at a trailing edge of the airfoil. A plurality of cooling openings extend between an internal cooling chamber and an external surface of the first sidewall. One cooling opening extends from the cooling chamber into the inflection at a relatively shallow injection angle with respect the airfoil external surface.

Abstract: A gas turbine engine wash process that facilitates reducing a formation of particulate matter within a gas turbine engine is described. The engine wash process includes injecting a first liquid into the engine to remove particulate matter formed within the engine and adversely affecting engine operation and performance. A second liquid is then injected into the engine to facilitate reducing a rate of formation of particulate matter within the gas turbine engine as the engine is operated in the future. More specifically, the second liquid is an anti-static liquid that coats compressor blades within the gas turbine engine.

Abstract: A method for fabricating a three-dimensional turbine engine combustor liner model includes coupling at least a first member to a second member to form an assembly that has an inner surface. The assembly inner surface simulates an inner surface of the aircraft engine combustor liner. The first member is at least one of a pre-formed conical member and a pre-formed cylindrical member, and the second portion is at least one of a pre-formed conical member and a pre-formed cylindrical member. The method also includes coupling the assembly to a baseplate, and coupling a plurality of templates to the assembly.

Abstract: A method and system for pacing a vehicle along a path of travel is described. The method includes determining a geographical location of the vehicle, displaying a vehicle position icon representative of the geographical location, determining an optimal position for the vehicle, displaying a pace icon representative of the optimal position for the vehicle, and operating the vehicle to maintain a vehicle position icon displayed on the operator pace display substantially coincident with the pace icon displayed on the operator pace display.

Abstract: A control strategy for a gas turbine engine which exchanges future lifetime of the engine for present thrust. Gas turbine engines, such as those used in aircraft, sometimes incur damage, as when they ingest birds, or are struck with ballistic objects fired by an enemy. The invention detects the damage, and invokes a control strategy wherein the engine is operated in a more harsh manner, thereby sacrificing a significant part of the remaining lifetime of the engine, in order to obtain thrust currently.

Abstract: A fuel nozzle including a nozzle stem having an annular overhang and a heat shield secured to the overhang is described. More specifically, and in one embodiment, the nozzle stem includes an upstream end and a downstream end. The annular overhang is intermediate to the upstream end and the downstream end of the stem. The heat shield includes a first end and a second end, and the heat shield is welded to the annular overhang at the heat shield first end. An annular air gap is between the nozzle stem and the heat shield.

Abstract: A rotor overspeed protection system for a gas turbine engine controls engine fuel flow to prevent an engine rotor from over-speeding. The engine fuel metering system includes a fuel metering valve in flow communication with a fuel shutoff valve and a fuel bypass valve. The rotor protection system includes an overboost servovalve and a soleniod valve coupled to the fuel metering system and to an independent speed sensing system. In operation, servovalve can control the metered fuel flow to engine independently of the fuel metering valve to facilitate reducing rotor overspeeds and overboosts.

Abstract: A method for assembling an articulated fan front frame for a gas turbine engine facilitates improving engine performance. The method includes forming a strut including a pair of sidewalls connected at a leading edge and a trailing edge, forming a flap including a first sidewall and a second sidewall connected at a leading edge and a trailing edge, and extending in radial span between a root endwall and a tip endwall, wherein each endwall extends between the first and second sidewalls, and wherein at least one of the root endwall and the tip endwall is contoured in a radial direction extending between the flap leading and trailing edges, and pivotally coupling the flap downstream from the strut such that a gap is defined between the flap and the strut.

Abstract: In one aspect of the invention a method for processing a fluoroscopic image is provided. The method includes scanning an object with an imaging system including at least one radiation source and at least one detector array, acquiring a plurality of dark images to generate a baseline image, acquiring a plurality of lag images subsequent to the baseline image, determining a plurality of parameters of a power law using at least one lag image and at least one baseline image, and performing a log—log extrapolation of the power law including the determined parameters.

Abstract: A communication system is provided that includes a plurality of remote terminal units (RTUs) that communicate with a server via a network. The network provides communication paths to the RTUs via both a sessionless protocol and a switched connection protocol. Generally, the RTUs communicate via the network using the sessionless protocol. The network also includes a server. The RTUs share at least one account identifier that is coordinated by the server. Short messages are sent via the sessionless protocol. Longer messages are sent via the switched connection protocol using a shared account identifier. A corresponding method of operating the network is also provided. The use of a shared account identifier avoids the requirement of obtaining access accounts to the switched connection protocol for each of the RTUs, and is especially efficient and economical when the need to communicate longer messages is infrequent, but where each of a large number of RTUs require this capability.