Abstract: A power plant for an aircraft such as a UAV with a gas turbine engine that drives an electric generator to produce electrical power. The electric generator is rotatably supported by two foil bearings. A centrifugal compressor is secured to a forward side of the generator rotor shaft. The centrifugal compressor draws in cooling air that flows through the two foil bearings and between a space formed between the rotor coil and the stator coil of the electric generator to provide for cooling of both foil bearings and the coils of the generator.

Abstract: A gas turbine shroud ring segment assembly (36) with an outer structural block (32A) having cooling channels (72) between inlets (60) on a front face and outlets (74, 76) on a front hanger rail (48). The outlets may be positioned on a radially inner surface (77) of the front rail for impingement and forced convection cooling of backsides of radially inner front lips (44) of adjacent shroud ring segments (30A, 30B) mounted on front and rear rails (48, 50) of the block. The outlets may enter a pocket (86) on the inner surface configured to allow coolant flow in all positions of the ring segments (32A, 32B). The cooling channel may form a main channel (72A) and tributary channels (72B, 72C). These channels may be drilled upward from the rail to the inlet. The tributaries may have offset intersections (72E, 72D) with the main channel.

Abstract: A seal strip (60A, 60B, 60C, 60F, 60G, 60H, 60J) with an imperforate width-spanning portion (64) first and second rounded edges (65, 66) and one or more strip-thickening elements (67, 68, 74, 76, 82, 84, 90, 91) between the rounded edges. The strip-thickening elements may have transverse slots (80, 86, 88, 93L, 93R, 72, 78) for increased flexibility of the strip. The strip-thickening elements may also have perforations (92) or gaps (69) to admit coolant and/or to reduce weight. Folded embodiments (60A, 60B) may have dimples (70) on the width-spanning portion (64) to limit bending of the folded portions (67, 68, 74, 76). The seal may be slidably mounted in opposed slots (58A, 58B) in respective adjacent turbine components (54A, 54B), filling a width (W) of the slots, and a side of the seal may be cooled by compressed air (48).

Abstract: A turbine blade tip shroud (22C, 22D) with a seal rail (32C, 32D) oriented in a circumferential direction of rotation (29) relative to a turbine axis, and extending radially outward from the tip relative to the turbine axis. A first tooth (48, 68) and a second tooth (50, 70) form respective downstream and upstream lateral departures or bumps on the seal rail. Each tooth has a sharp top leading edge (48, 50, 68, 70) and a smoothly curved side surface (49, 51, 69, 71). A back portion (56, 76) of the seal rail may span linearly from a lateral peak (66, 78) of the second tooth to a back end (62) of the seal rail that is centered on an extended centerline (60) of a front portion (54, 74) of the seal rail. The teeth may be disposed proximate or over a stacking axis (52) of the blade.

Abstract: Respective seals (54, 78) for the upper and lower spans (48A, 48B) of an exit frame (48) of a turbine combustion system transition piece (28). Each seal has a first strip (55, 79) and a second strip (66, 88) of a sealing material. The two strips of each seal are attached together along a common edge. The second strip is flexible, generally parallel to the first strip, and has a bead (72, 90) along its free edge. This forms a spring clamp that clamps a rail (68, 86) of the exit frame between the bead and the first strip of each seal. A tab extends axially aft from the first strip of each seal for insertion into a circumferential slot (58, 82) in a turbine inlet support structure (52, 76), thus sealing the transition piece (46) to the turbine inlet for efficient turbine operation.

Abstract: A turbine blade tip shroud (22C, 22D) with a seal rail (32C, 32D) oriented in a circumferential direction of rotation (29) relative to a turbine axis, and extending radially outward from the tip relative to the turbine axis. A first tooth (48, 68) and a second tooth (50, 70) form respective downstream and upstream lateral departures or bumps on the seal rail. Each tooth has a sharp top leading edge (48, 50, 68, 70) and a smoothly curved side surface (49, 51, 69, 71). A back portion (56, 76) of the seal rail may span linearly from a lateral peak (66, 78) of the second tooth to a back end (62) of the seal rail that is centered on an extended centerline (60) of a front portion (54, 74) of the seal rail. The teeth may be disposed proximate or over a stacking axis (52) of the blade.

Abstract: A gas turbine engine blade (10), having: an airfoil (12) having a pressure side (22) and a suction side (24); an inner platform (30), having: a gas path surface (40); a coolant surface (42); and at least one row (114) of film cooling holes (112) spanning there between and disposed on the pressure side of the airfoil. The film cooling holes are angled to include a directional component substantially parallel with the pressure side at a trailing end of the airfoil, and the at least one row is oriented substantially parallel to a mateface wall (122) of the inner platform. An array (130) of turbulators (132) is disposed on the coolant surface.

Abstract: A gas turbine shroud ring segment assembly (36) with an outer structural block (32A) having cooling channels (72) between inlets (60) on a front face and outlets (74, 76) on a front hanger rail (48). The outlets may be positioned on a radially inner surface (77) of the front rail for impingement and forced convection cooling of backsides of radially inner front lips (44) of adjacent shroud ring segments (30A, 30B) mounted on front and rear rails (48, 50) of the block. The outlets may enter a pocket (86) on the inner surface configured to allow coolant flow in all positions of the ring segments (32A, 32B). The cooling channel may form a main channel (72A) and tributary channels (72B, 72C). These channels may be drilled upward from the rail to the inlet. The tributaries may have offset intersections (72E, 72D) with the main channel.

Abstract: A gas turbine engine blade (10) having: an airfoil (12) having a leading edge (14), a trailing edge (16), a pressure side (22), and a suction side (24); an inner platform (30) associated with a base (18) of the airfoil, wherein the inner platform includes a platform aft face (34) spanning between an platform pressure side face (36) and a platform suction side face (38); and a trailing edge undercut (70) across the aft face. The trailing edge undercut tapers to zero penetration before reaching the platform suction side face. In a side cross section, a leading portion (100) of the trailing edge undercut is defined by a leading radius (102), an upper portion (104) is defined by an upper radius (106), and a lower portion (108) is defined by a lower radius (110). For at least one location along the aft face, values for all three radii are distinct.

Abstract: A scoop (54) over a coolant inlet hole (48) in an outer wall (40B) of a double-walled tubular structure (40A, 40B) of a gas turbine engine component (26, 28). The scoop redirects a coolant flow (37) into the hole. The leading edge (56, 58) of the scoop has a central projection (56) or tongue that overhangs the coolant inlet hole, and a curved undercut (58) on each side of the tongue between the tongue and a generally C-shaped or generally U-shaped attachment base (53) of the scoop. A partial scoop (62) may be cooperatively positioned with the scoop (54).

Abstract: A combustion chamber liner (41) with a forward section (44) and an aft section (46). The aft section has an array of aft axial cooling fins (62) covered by a tubular support ring (52), thus forming an array of aft axial grooves (66) between the aft axial fins. Inlet holes (54) in the front end of the support ring may admit coolant (37) into an upstream end of the aft axial cooling fins. An impingement plenum (61) may receive the coolant just before the aft axial cooling fins. Each aft axial fin may include a plurality of axially spaced bumpers (64) that contact the support ring. Spaces or grooves (68) between the bumpers provide circumferential cross flow of coolant between the grooves. The aft axial grooves may discharge the coolant as film cooling along the inner wall (76) of a transition duct (28).

Abstract: A seal strip (54) with a central relatively thin portion (68) and first and second thicker side portions (70, 72) that may be wedge-shaped adjacent the central portion. Each side portion may be formed of a linear array of base-in prisms (56), where each prism includes a base adjacent and normal to the central portion, and a thickness tapering distally toward an adjacent edge of the seal strip. The base-in prisms of each side portion may be separated by transverse slots (55) along the length of the strip. The transverse slots of the first side portion may be unaligned with the transverse slots of the second side portion along the length of the strip. A retention pin (58) may extend normally from an end of the seal strip. The seal strip may be mounted in tapered slots (49) of a gas turbine transition exit frame (48).

Abstract: A process for redesigning a distressed component in which the distressed component is under thermal and structural loads, for improving the life of the component. The process includes obtaining the operating conditions of the machine in which the distressed component is used, finding the boundary conditions under which the distressed component operates, producing a 3-dimensional model of the distressed component with such detail that the distress levels are accurately represented on the model, subjecting the model to a series of technical analysis to predict a life for the component, reiterating the technical analysis until the levels of distress on the model accurately represent the distress that appears on the actual component, and then predicting a remaining life of the component based on the analysis, or redesigning the model and reanalyzing the model until a maximum life for the component has been found.

Abstract: Respective seals (54, 78) for the upper and lower spans (48A, 48B) of an exit frame (48) of a turbine combustion system transition piece (28). Each seal has a first strip (55, 79) and a second strip (66, 88) of a sealing material. The two strips of each seal are attached together along a common edge. The second strip is flexible, generally parallel to the first strip, and has a bead (72, 90) along its free edge. This forms a spring clamp that clamps a rail (68, 86) of the exit frame between the bead and the first strip of each seal. A tab extends axially aft from the first strip of each seal for insertion into a circumferential slot (58, 82) in a turbine inlet support structure (52, 76), thus sealing the transition piece (46) to the turbine inlet for efficient turbine operation.

Abstract: A seal strip (60A, 60B, 60C, 60F, 60G, 60H, 60J) with an imperforate width-spanning portion (64) first and second rounded edges (65, 66) and one or more strip-thickening elements (67, 68, 74, 76, 82, 84, 90, 91) between the rounded edges. The strip-thickening elements may have transverse slots (80, 86, 88, 93, 93R, 72, 78) for increased flexibility of the strip. The strip-thickening elements may also have perforations (92) or gaps (69) to admit coolant and/or to reduce weight. Folded embodiments (60A, 60B) may have dimples (70) on the width-spanning portion (64) to limit bending of the folded portions (67, 68, 74, 76). The seal may be slidably mounted in opposed slots (58A, 58B) in respective adjacent turbine components (54A, 54B), filling a width (W) of the slots, and a side of the seal may be cooled by compressed air (48).

Abstract: A seal strip (54) with a central relatively thin portion (68) and first and second thicker side portions (70, 72) that may be wedge-shaped adjacent the central portion. Each side portion may be formed of a linear array of base-in prisms (56), where each prism includes a base adjacent and normal to the central portion, and a thickness tapering distally toward an adjacent edge of the seal strip. The base-in prisms of each side portion may be separated by transverse slots (55) along the length of the strip. The transverse slots of the first side portion may be unaligned with the transverse slots of the second side portion along the length of the strip. A retention pin (58) may extend normally from an end of the seal strip. The seal strip may be mounted in tapered slots (49) of a gas turbine transition exit frame (48).

Abstract: A combustion chamber liner (41) with a forward section (44) and an aft section (46). The aft section has an array of aft axial cooling fins (62) covered by a tubular support ring (52), thus forming an array of aft axial grooves (66) between the aft axial fins. Inlet holes (54) in the front end of the support ring may admit coolant (37) into an upstream end of the aft axial cooling fins. An impingement plenum (61) may receive the coolant just before the aft axial cooling fins. Each aft axial fin may include a plurality of axially spaced bumpers (64) that contact the support ring. Spaces or grooves (68) between the bumpers provide circumferential cross flow of coolant between the grooves. The aft axial grooves may discharge the coolant as film cooling along the inner wall (76) of a transition duct (28).

Abstract: A scoop (54) over a coolant inlet hole (48) in an outer wall (40B) of a double-walled tubular structure (40A, 40B) of a gas turbine engine component (26, 28). The scoop redirects a coolant flow (37) into the hole. The leading edge (56, 58) of the scoop has a central projection (56) or tongue that overhangs the coolant inlet hole, and a curved undercut (58) on each side of the tongue between the tongue and a generally C-shaped or generally U-shaped attachment base (53) of the scoop. A partial scoop (62) may be cooperatively positioned with the scoop (54).

Abstract: A process for redesigning a distressed component, such as a turbine blade in a gas turbine engine, in which the distressed component is under thermal and structural loads, for improving the life of the component. The process includes obtaining the operating conditions of the machine in which the distressed component is used, finding the boundary conditions under which the distressed component operates, producing a 3-dimensional model of the distressed component with such detail that the distress levels are accurately represented on the model, subjecting the model to a series of technical analysis to predict a life for the component, reiterating the technical analysis until the levels of distress on the model accurately represent the distress that appears on the actual component, and then predicting a remaining life of the component based on the analysis, or redesigning the model and reanalyzing the model until a maximum life for the component has been found.

Abstract: An apparatus and process for testing airflow through a turbine vane under actual engine airflow conditions. The apparatus includes first a second static pressure chamber each with an opening on which a cover plate is to be secured to enclose the chamber and hold the vane in place for testing. Each chamber includes a pressurized air supply tube that extends into the chamber to deliver the volume of air necessary for testing the airflow. The chamber is of such a size that the high volume of pressurized air delivered to the chamber will enter the vane as static air for the testing. The cover plates are formed of two halves in which the vane endwall can be secured between the halves. A stand supports both cover plates so that the plates and the vane can be moved into position where the cover plates are secured to the opening of the chamber for testing. The pressurized air supplied to both chambers flows into the vane and through the film cooling holes so that proper testing of the airflow can be done.