Abstract: A seal includes first and second axially opposite loops integrally joined together by a coextensive web in a collective ring having a circumferential split. The two loops are arcuate in section, with radially outer and inner sealing lands. And, the web is disposed radially between the outer and inner lands. The loops are disposed in corresponding grooves of adjoining members for effecting a seal therebetween and accommodating differential radial and axial thermal movement.

Abstract: A gas turbine engine fan assembly has a plurality of axially swept fan exit guide vanes circumferentially disposed around an axially extending centerline, a plurality of fan frame struts having axially swept strut leading edges and circumferentially disposed around the centerline directly aft of the exit guide vanes, and an axially extending annular gap between the trailing edges of the fan exit guide vanes and the strut leading edges. Each of the fan exit guide vanes has pressure and suction sides, vane trailing edges, and is circumferentially leaned such that the pressure side facing radially inward. Preferably, the axially swept vane trailing edges and the axially swept strut leading edges generally conform to each other in shape. An annular splitter is radially disposed at a spanwise position to split airflow from the fan into fan bypass airflow and core engine airflow and includes a splitter leading edge preferably positioned aft of the strut leading edges.

Abstract: A method controls the stroke of three circumferentially spaced apart actuators attached at respective joints to an actuation ring positionable to vector thrust in a gas turbine engine exhaust nozzle. The circumferential angular location of each of the actuator joints is specified. And, the stroke for each of the actuators is independently calculated using the joint locations in response to pitch and yaw commands. The actuators are then operated to effect the respective strokes therefor to position the nozzle. In a preferred embodiment, the angular locations are specified once, and trigonometric functions thereof are calculated and stored in memory. The stored values may then be used subsequently in calculating stroke as the pitch and yaw commands change.

Abstract: A turbine shroud includes a panel having a forward end, an opposite aft end, and a middle therebetween. Cooling holes extend through the panel for channeling cooling air therethrough. The holes are arranged in a pattern of greater density aft of the panel middle than forward of the panel middle for effecting preferential cooling.

Abstract: A gas turbine engine airfoil is manufactured by forming an internal retention seat in two complementary airfoil parts. An insert is fabricated for retention in the seat. The two parts are assembled with the insert disposed in the seat therebetween. The parts are then bonded together to trap the insert therein to collectively define the airfoil. The insert and seat may be precisely fabricated for improving the efficiency of the airfoil.

Abstract: A gas turbine engine rotor blade 18 includes a dovetail 22 and integral airfoil 24. The airfoil includes a pair of sidewalls 28,30 extending between leading and trailing edges 32,34, and longitudinally between a root 36 and tip 38. The sidewalls are spaced laterally apart to define a flow channel 40 for channeling cooling air through the airfoil. The tip includes a floor 48 atop the flow channel, and a pair of ribs 50,52 laterally offset from respective sidewalls. The ribs are longitudinally tapered for increasing cooling conduction thereof.

Abstract: The flow of cooling fluid through a core tie hole formed between a pair of internal cooling passageways of an airfoil component is reduced by providing a restriction that meters coolant flow through the inlet passage of one of the passageways so that the pressures in the two passageways are equalized, thereby minimizing the flow of cooling fluid through the hole. The restriction can be a metering plate disposed at the entrance of the inlet passage or a restriction integrally formed in the inlet passage.

Abstract: A gas turbine engine airfoil includes first and second sidewalls joined together at opposite leading and trailing edges, and spaced apart from each other therebetween to define a leading edge channel extending longitudinally from a root to a tip of the airfoil. A plurality of film cooling holes extend through the leading edge and are disposed in flow communication with the leading edge channel. An isolation plenum extends along the first sidewall and adjacent the leading edge channel, and is separated therefrom by a partition having a plurality of inlet holes. A plurality of film cooling gill holes extend through the first sidewall, and are disposed in flow communication with the isolation plenum. Cooling air is channeled from the leading edge channel to the isolation plenum for feeding the gill holes with reduced pressure air.

Abstract: A turbine nozzle includes a plurality of vanes joined at opposite ends to outer and inner bands. The inner band has a forward hook which is segmented to reduce thermal mismatch. And, in additional embodiments the vane includes an impingement baffle having preferential cooling.

Abstract: A turbine blade includes a platform and an airfoil having a plurality of trailing edge slots separated by land areas larger than the slots. The slots have an exit diffusion half angle of about two degrees. Cooling air flows through the slots and over a trailing edge of the airfoil. The platform includes a plurality of openings extending through the platform at an angle. The openings are positioned between a suction side of the blade and a second end of the platform. The openings transport disk post cooling air to a surface of the platform and provide cooling for the platform.

Abstract: A turbine airfoil includes first and second sidewalls joined together at leading and trailing edges, and spaced apart to define first and second cavities separated by a septum therebetween. An aft bridge extends along the first cavity and includes a row of outlet holes therein. The septum includes a row of inlet holes. And, turbulators are disposed in rows inside the first cavity, and extend from the first sidewall toward the second sidewall. The turbulators are aligned with the inlet holes for multiple impingement cooling thereof.

Abstract: A turbine blade including an airfoil having a plurality of trailing edge slots and a plurality of trailing edge openings for use in a turbine engine is described. The trailing edge slots extend along a pressure side wall to the trailing edge at an upper portion of the blade. The trailing edge openings extend through the trailing edge at a lower portion of the blade. The trailing edge at the upper portion of the airfoil is relatively thin in comparison to the trailing edge at the lower portion of the airfoil, which allows for high aerodynamic efficiency and cooling of the airfoil.

Abstract: A turbine blade damper in the form of a sheet metal body includes a concave notch along one edge thereof, and a projecting side tab along an opposite edge thereof.

Abstract: A turbine blade includes an airfoil having a leading edge flow chamber disposed behind a leading edge in flow communication with an isolation flow channel disposed therebehind. The isolation channel is bounded on opposite sides by a pair of side flow channels.

Abstract: A turbine blade includes a hollow airfoil having a squealer tip rib extending outboard from a tip cap enclosing the airfoil. An impingement baffle is spaced inboard from the tip cap to define a tip plenum therebetween. Impingement holes extend through the baffle and are directed at the junction of the rib and cap for impinging a coolant jet thereagainst.

Abstract: A turbine airfoil includes first and second sidewalls joined together at leading and trailing edges, and extending longitudinally from root to tip. An internal cooling circuit is disposed between the sidewalls for channeling a coolant therein. A column of longitudinally spaced apart ejection slots extend through the first sidewall along the trailing edge, and are disposed in flow communication with the cooling circuit for discharging the coolant toward the trailing edge. Some of the slots are disposed between a pitchline and the tip and are inclined at different ejection angles longitudinally outwardly from the pitchline. And, some of the slots are disposed between the pitchline and root and extend at least parallel with the pitchline without longitudinally inward inclination.

Abstract: A method of making a hollow airfoil includes machining first and second parts to produce internal features thereof. The parts are then co-machined simultaneously at complementary joining surfaces. The parts are then bonded together at the co-machined joining surfaces.

Abstract: A gas turbine blade includes an airfoil, platform, and dovetail. The blade is cast with a chamfer along one edge of the platform thereof. The platform edge is then machined to truncate the chamfer in a simple and precise method.

Abstract: A concentrated reduced dosage spray pump delivery system for dispensing fluids containing actives and volatile organic compounds with reduced emissions of the volatile organic compounds is disclosed. The system includes a housing storing a concentrated fluid containing volatile organic compounds and an effective component, wherein the effective component is concentrated within the fluid. The system also includes a spray pump in fluid communication with the concentrated fluid containing the volatile organic compounds such that the spray pump dispenses the concentrated fluid at a reduced dosage per pump stroke. The combination of the concentrated fluid containing volatile organic compounds and the reduced dosage per pump stroke reduces the emission of volatile organic compounds, while the quantity of actives applied per square area of application surface substantially remains the same.

Abstract: The present invention provides a gas turbine engine blade having a dovetail, a shank extending radially outward from the dovetail, and a platform joined to the shank. An inner surface of the platform faces radially inwardly and an opposite outer surface of the platform faces radially outwardly. An airfoil extends radially outwardly from the platform and has pressure and suction airfoil sides that define pressure and suction blade sides of the blade. The platform extends axially between leading and trailing platform edges and transversely between pressure and suction side platform edges of the platform. At least one transversely extending bracing rib is in a corner of the shank and the platform between one of the blade sides and the inner surface of the platform. The preferred embodiment further includes the bracing rib, the shank, and the platform being integrally cast. The bracing rib is preferably wider along the platform and the shank than at a distal edge of the rib.