486 results about "Shock cooling" patented technology

A reheat combustion system for a gas turbine comprises a mixing tube adapted to be fed by products of a primary combustion zone of the gas turbine and by fuel injected by a lance; a combustion chamber fed by the said mixing tube; and at least one perforated acoustic screen. The or each said acoustic screen is provided inside the mixing tube or the combustion chamber, at a position where it faces, but is spaced from, a perforated wall thereof. In use, the perforated wall experiences impingement cooling as it admits air into the combustion system for onward passage through the perforations of the said acoustic screen, and the acoustic screen damps acoustic pulsations in the mixing tube and combustion chamber.

An apparatus and method for impingement cooling of a turbine vane, in which the vane includes an insert having a plurality of impingement cavities formed therein and in series such that cooling air flows from a first impingement cavity onto the wall for impingement cooling, and is then directed into the second impingement cavity and redirected for impingement cooling on another part of the wall. Supports for the insert form seals that direct cooling air from one impingement cavity into the next impingement cavity in the series. A trailing edge impingement cavity directs cooling air through holes to provide impingement cooling to the trailing edge region, the cooling air passing through a trailing edge discharge passage to cool the trailing edge. The insert is formed as a single piece, and has from 3 to 5 impingement cavities separated by ribs.

A combustor for a gas turbine engine includes an inner liner and an outer liner circumscribing the inner liner and forming a combustion chamber with the inner liner. The outer liner is a dual walled liner with a first wall and a second wall. The combustor includes a fuel igniter comprising a tip portion configured to ignite an air and fuel mixture in the combustion chamber and an igniter tube positioning the fuel igniter relative to the combustion chamber. The igniter tube includes a plurality of holes configured to direct cooling air toward the tip portion of the fuel igniter.

A method enables a gas turbine engine including a combustor to be operated. The combustor includes a mixer assembly including an air swirler, a premixer, a dome plate, and a heat shield. The method comprises discharging fluids from the air swirler into the premixer, and directing cooling fluids through a cooling opening defined in at least one of the swirler and the dome plate towards an upstream side of the heat shield for impingement cooling of the heat shield.

A gas turbine engine combustor having a dome heat shield includes a cooling scheme having a plurality of impingement cooling holes extending through the combustor and a plurality of adjacent ejector holes for directing cooling air past the heat shield lips of the dome heat shields. The impingement and ejector holes are preferably staggered to reduce interaction therebetween.

An extended impingement cooling structure to cool outside an air supply plenum comprises an inner wall; an impingement sheet; a series of supports to maintain the inner wall in spaced relation to the impingement sheet, and a baffle supported between the inner wall and the impingement sheet. The baffle has a collector plenum area that receives impingement cooling air from the air supply plenum and a channel in fluid communication with the collector plenum and extending outside the air supply plenum with openings to allow impingement cooling air to pass therethrough and having a series of lands extending into the channel wherein the lands are located in proximity to impingement cooling air outlets in the inner wall.

A cooling hole having an inlet passage forming an inward spiral flow path and an outlet passage forming an outward spiral flow path in which the two paths are counter flowing in order to improve the heat transfer coefficient. The spiral cooling hole is used in a blade outer air seal (BOAS) for a turbine in which the edges of the shroud segments include a counter flowing micro serpentine flow cooling circuit with thin diffusion discharge cooling slots for the BOAS edges. The total BOAS cooling air is impingement from the BOAS cooling air manifold and metered through the impingement cooling holes to produce impingement cooling onto the backside of the BOAS. The spent cooling air is then channels into the multiple micro serpentine cooling flow circuits located around the four edges of the shroud segments. This cooling air then flows in a serpentine path through the horizontal serpentine flow channels and then discharged through the thin diffusion cooling slots as peripheral purge air for the mate faces as well as the spacing around the BOAS or shroud segments. Trip strips are used in the serpentine flow channels for the augmentation of internal heat transfer cooling capability. The micro serpentine flow cooling air circuits spaced around the four edges of the shroud segments are formed into the shroud segments during the casting process of the shroud segments.

A shaft seal assembly is disclosed having a stator including a main body and axial and radial projections therefrom. The rotor is radially extended and encompasses the axial and radial projections from said stator. A passageway formed between the radial projection of stator and rotor results in an axial passageway having its opening facing rearwardly from the rotor and away from the source of impinging coolant and/or contaminant. A concentric circumferential receptor groove in the stator facing the housing allows insertion of conductive means for transmission of electrostatic charge away from the shaft through the shaft seal assembly to the housing and ground. The receptor groove is opposite the axial passageway and provides for both a substantially lower contaminant environment and improved engagement with conductive means. The dimension of interface gap between the rotor and the radial projection from the stator, which the access to the shaft of any impinging material is fixed at a predetermined value and does not vary with the relative movement between the rotor and the stator. The shaft seal assembly provides improved rejection or warding off of contaminants from ingress into the labyrinths and ultimately restrains attack of the bearing environment as well as substantial elimination of bearing current and attendant bearing fluting or frosting.

The invention relates to a side by side composite fiber spinneret plate, a method of preparing a three-dimensional crimp antibacterial fiber filament, and an application of the three-dimensional crimp antibacterial fiber filament. The method comprises the steps that the side by side composite fiber spinneret plate is used; every two groups of rectangular spinneret orifices in unequal dimensions at 120 degrees form every equilateral three-blade A and B; an ingredient A and an ingredient B with different thermal shrinkage rates are selected; an antimicrobial agent ingredient C is added into the ingredient A; the ingredient A and the ingredient C are subjected to melt extrusion by two twin screws, enter a parallel composite component and the spinneret plate after metered, and are subjected to parallel composite distribution, cross air blow, low-temperature shock cooling, oiling and winding to form a winding yarn; and the winding yarn is subjected to drawing, false twisting and thermoforming to form a DTY (Draw Textured Yarn) or subjected to the drawing and thermoforming to form the DTY. A three-dimensional crimp antibacterial fiber prepared by the method is soft and elastic, has the characteristics of good breathability, wicking, quick drying, antibacterium and the like since various grooves are formed in the cross section, and is applied to the development of shell fabrics of undergarments, sports suits, socks, bedclothes and the like and medical dressings.

A gas turbine engine airfoil has a platform cooling scheme including an impingement hole for directing cooling air against an undersurface of the airfoil platform.

A fuel lance is disclosed for injecting a gaseous and/or liquid fuel mixed with air into an axial hot gas flow flowing through a sequential combustor of a gas turbine, the fuel lance having at least one finger extending in a longitudinal direction into the axial hot gas flow of the gas turbine essentially perpendicular to the hot gas flow. The finger is configured as a streamlined body which has a streamlined cross-sectional profile. The body has two lateral surfaces essentially parallel to the axial hot gas flow. The body includes an enclosing outer wall having a longitudinally extending air plenum for the distributed introduction of air into the at least one finger. The air plenum is provided with a plurality of distributed impingement cooling holes, such that air exiting through the impingement cooling holes impinges on the inner side of the leading edge region of the body.

A turbine airfoil such as a turbine stator vane used in an industrial gas turbine engine, the vane including a trailing edge region having a thin wall cooling channel arrangement of mini cooling channels formed by a series of rows of elongated flow blockers that form the mini cooling channels between adjacent flow blockers. The adjacent row of flow blockers are offset from the each other such that the inlet and the outlet flow of cooling air is discharged directly onto the flow blocker in order to produce impingement cooling. The mini cooling channels have a spacing to hydraulic diameter ratio of less than 4.0 and a mini channel length to hydraulic diameter ratio of 5.0 of less in order to maintain a high flow velocity within the mini channels. In one embodiment, the flow blockers have a progressively decreasing length in the flow direction of the cooling air.

A turbine assembly with a hollow aerofoil having a main cavity with an impingement tube, insertable inside the main cavity for impingement cooling of an inner surface of the main cavity, and a platform at a radial end of the hollow aerofoil, and a cooling chamber for cooling the platform arranged relative to the hollow aerofoil on an opposed site of the platform. The cooling chamber is limited at a first radial end by a wall segment of the platform and at an opposed radial second end from a cover plate. The impingement tube extends in span wise direction through the cooling chamber from the platform to the cover plate and restricts a sub-cavity of the main cavity. The wall segment includes an entry aperture for a cooling medium to enter from the cooling chamber of the platform into the sub-cavity of the hollow aerofoil.