Abstract: A gas turbine engine blade includes an airfoil having first and second sides and an internal passage extending longitudinally therebetween for channeling air to cool the airfoil. Respective pluralities of first and second turbulator ribs extend inwardly from at least one of the first and second airfoil sides. Adjacent ones of second ends of the first and second ribs are staggered apart longitudinally and overlap chordally for defining a gap therebetween facing transversely for turning cooling air upon flow through the gap. The gap increases convective heat transfer enhancement for reducing cooling air requirements and improving blade creep life.

Abstract: A wall adapted for use in a gas turbine engine between a first and a hotter second fluid includes a first side over which is flowable the first fluid, and an opposite second side over which is flowable the second fluid. An elongate slot extends inwardly from the second side and is disposed in flow communication with a plurality of longitudinally spaced apart holes extending inwardly from the first side. The holes are disposed at a compound angle relative to the second side for discharging the first fluid obliquely into the slot and at a shallow discharge angle from the slot along the second side. The holes are also disposed in converging pairs to impinge together in the slot the first fluid channeled therethrough. In a preferred embodiment, the slot has an aft surface including a plurality of longitudinally spaced apart grooves extending from the holes to the wall second side.

Abstract: A jet engine component, such as an aircraft gas turbine engine rotor blade or a scramjet engine fuel injector. The component has a wall portion including a first surface exposable to a cooler, higher static pressure fluid and a second surface exposable to a hotter, lower static pressure gas flow flowing across the second surface. The component further includes a generally straight film coolant passageway having an inlet on the first surface and an outlet on the second surface. The second surface has a seamless groove which is open substantially entirely along its longitudinal dimension extending from the outlet along the lower static pressure gas flow for improved film cooling of the second surface.

Abstract: A wall adapted for use in a gas turbine engine between a first and a hotter second fluid includes a first side over which is flowable the first fluid, and an opposite second side over which is flowable the second fluid. An elongate slot extends inwardly from the second side and is disposed in flow communication with a plurality of longitudinally spaced apart holes extending inwardly from the first side. The holes are disposed at a compound angle relative to the second side for discharging the first fluid obliquely into the slot and at a shallow discharge angle from the slot along the second side. In a preferred embodiment, the slot has an aft surface including a plurality of longitudinally spaced apart grooves extending from the holes to the wall second side.

Abstract: A turbine blade includes a plurality of cooling passages each having a turbulated section of the passage preferentially located along the portion of the turbine blade subjected to the highest temperature. Thus, turbulent air flow is provided in intermediate sections of the blade to enhance the heat exchange relation with the metal of the blade. The bores of the cooling passages adjacent the tip and root portions are smooth and provide adequate cooling in those sections at a lower heat exchange relationship. The cooling passage bores are formed by an electrochemical machining process using an elongated electrode with a chemical electrolyte for forming enlarged cavities within the blade.

Abstract: A gas turbine engine hot section component such as a turbine blade or vane having an airfoil is provided a non-film cooled portion of a heat transfer wall having a hot surface and a plurality of longitudinally extending micro-grooves disposed in the portion of the wall along the hot surface in a direction parallel to the direction of the hot gas flow. The depth of the micro-grooves is very small and on the order of magnitude of a predetermined laminar sublayer of a turbulent boundary layer. The grooves are sized so as to alter the boundary layer thickness near the leading edge of the airfoil to reduce heat transfer from the hot gas flow to the airfoil near the leading edge. In one embodiment the micro-grooves are about 0.001 inches deep and have a preferred depth range of from about 0.001 inches to 0.005 inches and which are square, rectangular, or triangular in cross-section and the micro-grooves are spaced about one width apart.

Abstract: An air admission insert particularly adapted for air admission holes in a combustor liner in a gas turbine combustion system comprises a pair of sleeve members, one of which is inserted in the other in axial but eccentric relationship so that the sides of the sleeves contact each other with the intervening space between the sleeves being crescent shaped. The insert is positioned concentrically in a combustor liner air admission hole. Combustion air flows through the insert and into the combustor liner and axially through the crescent space to cool the insert as well as the liner hold perimeter.

Abstract: Insulating panels are formed from compressed particulate material impregnated with insulating gases, the combination being held in a gas tight pouch. These pouches can be placed in the insulation space of a structure where insulation is to be provided and, if desired, encapsulated in a foamed insulating material. Among the possible particulate materials is precipitated silica, while various Freon gases can be employed as the insulating gas.