Abstract: An apparatus and method for cooling a wall for use in a gas turbine engine is provided that includes a cooling air passage having a plurality of segments connected in series by one or more chambers, an inlet aperture, and an exit aperture. The inlet aperture connects the cooling air passage to one side of the wall. The exit aperture connects the cooling air passage to the opposite side of the wall. Cooling air on the inlet aperture side of the wall enters the cooling air passage through the inlet aperture and exits through the exit aperture.

Abstract: A method and apparatus for cooling a wall within a gas turbine engine is provided which comprises the steps of: (1) providing a wall having an internal surface and an external surface; (2) providing a cooling microcircuit within the wall that has a passage for cooling air that extends between the internal surface and the external surface; and (3) increasing heat transfer from the wall to a fluid flow within the passage by increasing the average heat transfer coefficient per unit flow within the microcircuit. According to one aspect, the present invention method and apparatus can be tuned to substantially match the thermal profile of the wall at hand.

Abstract: A hollow airfoil is provided which includes a body, a trench, and a plurality of cooling apertures disposed within the trench. The body extends chordwise between a leading edge and a trailing edge, and spanwise between an outer radial surface and an inner radial surface, and includes an external wall surrounding a cavity. The trench is disposed in the external wall along the leading edge, extends in a spanwise direction, and is aligned with a stagnation line extending along the leading edge.

Abstract: A hollow airfoil is provided which includes a body having an external wall and an internal cavity. The external wall includes a suction side portion and a pressure side portion. The portions extend chordwise between a leading edge and a trailing edge and spanwise between an inner radial surface and an outer radial surface. A stagnation line extends along the leading edge. A plurality of cooling apertures, disposed spanwise along the leading edge. According to one aspect of the present invention, the apertures extend through the external wall along a helical path. According to another aspect of the present invention, the apertures are alternately directed towards the suction side portion and the pressure side portions of the airfoil.

Abstract: Film cooling effectiveness for internally air cooled turbine blades of gas turbine engines is enhanced by including an interconnecting passage to the film cooling hole that is in communication with a lower internal pressure in the blade to create a suction at the exit end of the film hole to limit penetration into the gas stream.

Abstract: A hollow airfoil is provided which includes a body, a trench, and a plurality of cooling apertures disposed within the trench. The body extends chordwise between a leading edge and a trailing edge, and spanwise between an outer radial surface and an inner radial surface, and includes an external wall surrounding a cavity. The trench is disposed in the external wall along the leading edge, extends in a spanwise direction, and is aligned with a stagnation line extending along the leading edge.

Abstract: A bypass air valve for a gas turbine engine is provided which includes a liner, a strap, and means for selectively actuating the valve. The liner includes an inner radial surface, an outer radial surface, a plurality of first regions, and a plurality of second regions. Each first region includes a plurality of first apertures. Each second region includes a plurality of second apertures. The strap includes a plurality of openings and a plurality of third regions. Each third region includes a plurality of third apertures. The valve may be selectively actuated into an open or a closed position. In the open position, the first regions are substantially aligned with the openings and the third regions are substantially aligned with the second regions. In the closed position, the first regions are substantially aligned with the third regions and the second regions are substantially aligned with the openings.

Abstract: An internally air cooled turbine blade for a gas turbine engine of the type including a trailing edge section, leading edge section and mid chord section wherein each section includes a straight through radial passage communicating cooling air from the root to the tip of the blade, and the radial passages (feed channels) on the pressure side and suction side supply the cooling air to the film cooling holes in the airfoil surface and the radial passage (feed chamber) in the mid chord section replenishes the feed channels with cooling air through replenishment cooling holes interconnecting the feed channels and feed chamber. The rotation of the blade imparts a centrifugal pumping action to the air within the feed chamber for maximizing the cooling effectiveness of the cooling air. The air discharging at the tip cools the tip of the blade and provides tip aerodynamic sealing and the leading edge and trailing edge are similarly fed cooling air.

Abstract: An internally cooled turbine blade for a gas turbine engine is modified at the leading and trailing edges to include a dynamic cool air flowing radial passageway with an inlet at the root and a discharge at the tip feeding a plurality of radially spaced film cooling holes in the airfoil surface. Replenishment holes communicating with the serpentine passages radially spaced in the inner wall of the radial passage replenish the cooling air lost to the film cooling holes. The discharge orifice is sized to match the backflow margin to achieve a constant film hole coverage throughout the radial length.

Abstract: A bulkhead in the form of a rib extending spanwise adjacent the leading edge of the airfoil of a turbine blade of a gas turbine engine serves to stiffen the airfoil to attain protection from impact by foreign objects, but doesn't materially affect cooling effectiveness in the event of a puncture of the skin. Holes extending spanwise in the rib reduce the mass, allow radial flow, and restrict flow into the leading edge. This construct allows a reduction in the thickness of the skin of the airfoil, and hence a weight reduction.

Abstract: The leading edge of the turbine airfoil of a turbine blade for a gas turbine engine is cooled by utilizing helix shaped holes that are rectangular in cross-section and curved to produce film holes egressing at low angles adjacent the outer surface of the leading edge. The helix holes extend from the root to the tip of the blade and are parallelly and longitudinally spaced.

Abstract: The method of manufacturing an air cooled stator vane for a gas turbine engine by casting an inner shell with suction side wall and pressure side wall of the airfoil, stamp forming from a blank sheet metal a pair of sheaths and attaching one sheath of the pair of sheaths to the suction side wall and the other sheath of the pair of sheaths to the pressure side wall. Forming in the step of stamp forming pockets with a slot at the end of each pocket for flowing a film of cooling air over the exterior of each of the pair of sheaths and drilling holes in the end of the pocket remote from the slot prior to the step of joining to fluidly connect the pocket with cooling air.

Abstract: An air cooled stator vane or the like for a gas turbine engine is constructed to include a plurality of axially and radially spaced pockets formed on the outer surfaces of the airfoil sections. Air is directed into the pocket on one end of a passage defined by the pocket to impinge on the back wall of the pocket, change directions and flow through the passage and discharge into the gas path as a film of cooling air coalesced by slots formed at the end of the pocket. The outer wall can be cast in its entirety or fabricated from sheet metal sheaths supported to a cast inner shell. The pockets are fed cooling air through a central passage in the stator vane or through impingement inserts disposed in the central passage.

Abstract: Passive clearance control is attained for rotor blades for gas turbine engines in environments where the airfoil of the rotor is solid or where coolant is unavailable or is available but not at sufficient quantity to provide passive clearance control. In this situation a judiciously located curved slot extending from the tip of the blade adjacent the pressure side and terminating at the pressure side at a location where there is sufficient static pressure so that the pressure drop across the slot induces a pumping action. In embodiments where there is sufficient coolant for tip cooling, the curved slot is judiciously located to be in communication with the film cooling hole discharging on the pressure side of the blade and the tip of the blade adjacent the pressure side of the blade.

Abstract: An air cooled stator vane for a gas turbine engine is fabricated by casting the suction side wall and pressure side wall of the airfoil into separate halves and joining both halves at complementary sides and ribs, and forming pockets with a slot at the end of each pocket for flowing a film of cooling air over the exterior wall and drilling holes in the end of the pocket remote from the slot to fluidly connect the pocket with cooling air. The method includes in another embodiment a sheet metal sheath with pockets and slots identical to the pockets and slots in the shell configuration formed over the shell. Another embodiment includes perforated inserts extending in the cavity formed by the shell.

Abstract: This disclosure describes a cooling and manufacturing enhancement for gas turbine engine outer air seals. The invention provides convective cooling in a counter-flow array, internal heat transfer augmentation through the use of boundary layer trip-strips, coolant Mach number control through the use of tapered convective channels, inter-segment purge without additional secondary cooling air, and is castable.

Abstract: A cooling enhancement for a gas turbine engine blade outer air seal is made up of a plurality of arcuate seal segments, and each one has one or more internal cooling passages. Each segment has a supply surface opposite and in spaced relation to a sealing surface, and first and second end surfaces extending from the sealing surface to the supply surface. Each of the end surfaces includes a substantially planar perpendicular surface and a substantially planar canted surface intersecting at an included angle of between 100 degrees and 140 degrees, and the supply surface has a plurality of cooling orifices adjacent the second end surface. At least one cooling passage extends from the first end surface towards the second end surface, and each cooling passage intersects the first end surface and terminates short of the second end surface. Each cooling passage communicates with the supply surface through at least one of the cooling orifices.

Abstract: An air cooled gas engine turbine blade that includes a plurality of longitudinally spaced cavities adjacent the leading edge of the blade is designed to include angularly disposed impingement passages flowing cooling air into each of the cavities in a direction extending from the root to the tip of the blade and including an annular projection upstream of the impingement passage but adjacent thereto for directing air into the respective cavities with total instead of static pressure. The impingement holes are oriented to align with the film cool holes in the blade surface at the leading edge. Ribs formed between cavities are also oriented to be parallel to the impingement holes.

Abstract: Internal cooling of the trailing edge of the airfoil of the turbine blade for a gas turbine engine includes a cascade formed from juxtaposed rows of longitudinally extending spaced airfoil shaped vanes or ribs leading cooling air from a supply source through the space between adjacent vanes and discharging out of the blade.

Abstract: Flow rate of coolant air is increased upon elevated nozzle gas temperatures to increase nozzle liner cooling effect. A hot side curved plate is superposed over a cold side curved plate which overlays a pressurized cooling air cavity. First and second sets of apertures are formed through the first and second plates respectively. At low temperature operation, the gap between the plates is small and the staggering of the apertures restricts the flow rate of cooling gas through the plates. As temperatures rise, the hot side plate thermally expands and the gap between the plates increases to reduce the mechanical impedance effect of the staggering and to thus increase coolant flow rates.