Abstract: A hollow cooled turbine airfoil is provided having pressure and suction side walls and a plurality of trailing edge cooling passages that feed cooling air bleed slots at the trailing edge. The airfoil trailing edge is selectively thickened in the root portion so as to allow shortened trailing edge slots, thereby improving trailing edge cooling and reducing mechanical and thermal stresses.

Abstract: A dovetail assembly including non-parallel relief faces that facilitates reduced pressure face brinelling in turbine engines. The assembly includes a plurality of rotor blades, each including a dovetail. Each dovetail includes at least a pair of blade tangs including blade relief faces. The dovetail assembly also includes a rotor disk including a plurality of dovetail slots, each sized to receive a dovetail. Each dovetail slot is defined by at least one pair of opposing disk tangs including disk relief faces. The disk relief faces are non-parallel to the blade relief faces when the dovetail is mounted in the dovetail slot.

Abstract: A cooling air cooling system is selectively operable to reduce fuel gum deposits within the cooling system during gas turbine engine operations. The cooling system includes a recirculating loop that includes at least three heat exchangers in fluid communication with the recirculating loop. A first heat exchanger cools cooling air supplied to the gas turbine engine. A second heat exchanger cools fluid exiting the first of the heat exchangers with fan discharge air prior to the fluid entering the third heat exchanger. A third heat exchanger uses combustor main fuel flow to cool the fluid circulating in the recirculating loop.

Abstract: A centerbody for a gas turbine engine includes a thermal control system which minimizes the thermal stresses between the centerbody and at least one stiffener. The centerbody stiffener extends radially inward from a centerbody shell. A cavity is defined within each centerbody stiffener. The thermal control system includes a plurality of openings circumferentially disposed around the centerbody. Each opening extends through the centerbody shell into each cavity. The openings include pairs of entrance openings and exit openings which permit circumferential flow to develop within the centerbody cavity.

Abstract: A gas turbine engine including a rotor shaft includes a bearing isolation system including an isolation shaft and a pair of pilots. The pilots are disposed between the isolation and rotor shafts and connect the isolation shaft to the rotor shaft to form a gap between the isolation and rotor shafts. The pilots are positioned axially along the rotor shaft and permit the isolation shaft to move axially with respect to the rotor shaft. Each pilot includes a plurality of openings to permit cooling air to flow within the gap defined between the isolation shaft and the rotor shaft.

Abstract: A cooling air cooling system operable to reduce fuel gum deposits within the cooling system when a gas turbine engine operates above a predefined percentage of rated engine power. The cooling system includes a recirculating loop including a plurality of heat exchanges in fluid communication with the recirculating loop. A first head exchange uses heat transfer fluid to cool cooling air used by the gas turbine engine. A second head exchange is a fluid-fuel head exchanger that uses combustor main fuel flow to cool the head transfer fluid circulating in the recirculating loop.

Abstract: A gas turbine engine includes rotor blades including airfoils that facilitates reducing manufacturing losses due to airfoil trailing edge scarfing. Each airfoil includes a first and second sidewall connected at a leading edge and a trailing edge. The sidewalls define a cooling cavity that includes at least a leading edge chamber bounded by the sidewalls and the airfoil leading edge, and a trailing edge chamber bounded by sidewalls and the airfoil trailing edge. The cooling cavity trailing edge chamber includes a tip region, a throat, and a passageway region connected in flow communication such that the throat is between the tip region and the passageway region. Furthermore, the tip region is bounded by the airfoil tip and extends divergently from the throat, such that a width of the tip region is greater than a width of the throat.

Abstract: A combustor for a gas turbine engine includes a deflector assembly that enhances heat transfer from the combustor and minimizes low cycle fatigue stresses induced within the combustor. The deflector assembly includes a plurality of deflectors secured to a spectacle plate. Each deflector has tapered edges and includes a plurality of cylindrical projections extending outward from the deflector to facilitate heat transfer. The projections include rounded edges and are arranged in a high density pattern. The deflector is coated with a thermal barrier coating and a bondcoat to minimize exposure to hot combustion gases or flame radiation.

Abstract: A turbine blade assembly for a gas turbine engine that includes a turbine platform including contoured edges. The turbine blade assembly includes a platform, a turbine airfoil extending radially outward from the platform, a shank that extends radially inward from the platform, and a dovetail extending from the shank. The platform includes a leading edge and a trailing edge separated by a pressure edge and an opposite suction edge, the pressure edge including a plurality of arcs extending between the leading edge and the trailing edge, the suction edge including a plurality of arcs extending between the leading and trailing edges.

Abstract: A hollow cooled turbine airfoil is provided having pressure and suction side walls and a plurality of trailing edge cooling passages that feed cooling air bleed slots at the trailing edge. The suction side wall has a selectively reduced thickness, which allows shortened trailing edge slots, improving trailing edge cooling.

Abstract: A combustor for a gas turbine includes a dome assembly that facilitates extending a useful life of the combustor in a cost-effective and reliable manner. The dome assembly includes a dome plate and a heat shield coupled to the dome plate. The dome plate includes an impingement baffle and an opening extending therethrough and sized to receive a fuel injector. The impingement baffle also includes a plurality of cooling openings in flow communication with the heat shield that direct cooling airflow for impingement cooling and film cooling of the heat shield.

Abstract: A one-piece deflector-flare cone assembly for a gas turbine engine combustor that facilitates extending a useful life of the combustor in a cost-effective and reliable manner is described. The one-piece assembly includes a deflector portion and a flare cone portion. The deflector portion includes an integral opening that extends through the deflector portion for receiving cooling fluid therein. The cooling opening extends circumferentially within the deflector portion. Cooling fluid discharged from the cooling opening is used for film cooling a portion of the deflector portion to facilitate reducing an operating temperature and extending a useful life of the combustor.

Abstract: A turbine nozzle includes ceramic outer and inner bands and an integral ceramic vane aft segment. A metal vane forward segment adjoins the aft segment in a hybrid assembly.

Abstract: A turbine cooling circuit includes a flow sleeve having an aft end for adjoining the rim of a turbine disk. A seal rotor is spaced from the aft end adjacent a row of inlet holes. An inducer has an outlet disposed radially outwardly from the sleeve inlet holes. And, a seal stator surrounds the seal rotor to define a rotary seal disposed radially inwardly from the inducer outlet.

Abstract: A gas turbine engine component including a nozzle outer band, a plurality of nozzle vanes extending inward from the outer band, and an inner band extending circumferentially around inner ends of the vanes. Further, the component has a shroud integral with the outer band adapted for surrounding a plurality of blades mounted in the engine for rotation about a centerline thereof.

Abstract: Methods and apparatus for reducing temperature overshoot in an engine, such as in a gas turbine aircraft engine, are described. In an exemplary embodiment, an unmeasured temperature to be regulated in an engine is determined by measuring a temperature in the engine wherein the measured temperature being related to the unmeasured temperature, determining a bias of the measured temperature wherein the bias being an amount estimated to be a difference between the measured temperature and a steady state measured temperature and adding the bias to the measured temperature to restore the relationship between measured and unmeasured temperature so that the unmeasured temperature may be properly regulated. The bias is determined using a heat transfer model.

Abstract: An aerospike engine has at least one nozzle surface and a plurality of pulse detonation devices mounted to the nozzle surface in place of the more common deflagration-based combustors. Each pulse detonation device is oriented such that its combustion products are directed along the nozzle surface Incorporating pulse detonation devices into the aerospike engine produces the advantage of a more efficient thermodynamic cycle. The pulse detonation aerospike engine is also capable of operating on either air or oxidizer.

Abstract: A cooled turbine element including an airfoil and a flowpath boundary member extending laterally from either an inboard end or an outboard end of the airfoil. The member has a flowpath face and an outside face which is cooler than said flowpath face creating a tendency for the member to deflect in a direction away from the flowpath face and causing a thermally induced tensile radial stress in a region of the trailing edge of the airfoil. The element has an interior cooling passage and at least one cooling hole extending from the interior cooling passage to an opening located in an area upstream from the stressed region of the trailing edge to cool the area so the airfoil thermally deflects to a shape corresponding to that of the boundary member thereby lowering the thermally induced tensile radial stress in the airfoil at the trailing edge thereof.

Abstract: A multi-hole cooled combustor liner is provided which reduces hot streaks and associated material distress in the liner. Areas of increased cooling hole density are disposed upstream of the primary dilution holes and in circumferential alignment with fuel cup centers. Additional cooling holes are provided in between primary dilution holes are arranged in alternating pairs of circumferentially angled holes so as to provide a converging cooling air flow in the downstream direction.

Abstract: A gas turbine engine rotor assembly including a rotor having a radially outer rim with an outer surface shaped to reduce circumferential rim stress concentration between each blade and the rim. Additionally, the shape of the outer surface directs air flow away from an interface between a blade and the rim to reduce aerodynamic performance losses between the rim and blades. In an exemplary embodiment, the outer surface of the rim has a concave shape between adjacent blades with apexes located at interfaces between the blades and the rim.