Abstract: A seat assembly (10) with a plurality of individual segments (16, 18, 20). Each segment includes a rigid flapper (26) hinged at one end to one movable member (14) and slidably contacting a planar surface (12') in another member (12).

Abstract: A system (38) and method for directing control actions to recover from the occurrence of a surge condition within a turbocompressor includes a high/low engine power output discriminator (40). Control action is taken by the system (38) in response to both the detection of a surge condition and the current engine power output level. The system (38) selects among a plurality of corrective control actions (56, 62, 68) to best achieve a return to normal engine operation.

Abstract: A convergent/divergent gas turbine engine exhaust nozzle having reverse thrust capability includes a concentric flap (22) and arc valve (36) pivotable about a common axis (34). A method and linkage (48) schedule the opening of the valve (36) in response to rotation of the flap (22) into a blocking position with respect to the normally aftward flow of the exhaust gas (12). The arc valve (36) is accelerated by the linkage (48) to full opening speed prior to unsealing an alternate, thrust reversing flow passage (38).

Abstract: A two-dimensional thrust vectoring nozzle (10) includes aftward divergent flaps (26, 28) positioned by a linkage assembly (38). The assembly (38) includes a linear track (42) pivotable about a transverse axis (44) for positioning a scissors linkage pivot (52) to selectively vary the divergence angle and vertical thrust vector of the nozzle (10).

Abstract: A linkage for area controlled thrust vectoring comprising of vane cascade (26) having a plurality of vanes (28, 29, 30) pivotable in unison and a flow regulating vane (36) oriented to provide a constant collection gas flow area is provided with a mechanical linkage for orienting the vanes (28, 29, 30) and the flow regulating vane (36) responsive to a single linear actuator (59).

Abstract: A linkage (36) is provided for manipulating an internal component (32) of a turbofan gas turbine engine (10) in response to an externally mounted actuator (50). The linkage includes a radial torque shaft (38) supported between two spherical bearings (40, 42). The shaft (38) is slidable within the radially outer bearing (42) for accommodating relative radial movement in addition to relative axial and circumferential movement between the inner compressor case (26) and the outer fan duct (28).

Abstract: A fuel spraybar (16) having an elongated, closed housing (26) discharges fuel from a plurality of openings (32) into an exhaust gas stream (12). A first fuel conduit (34) conducts liquid fuel through the housing (26) and into an annular volume (38) formed therebetween. The fuel, heated by contact with the housing (26) is vaporized in the annulus (38) for preferentially discharging liquid fuel (18) in an axial pilot zone (40) within the gas stream (12).

Abstract: A fastener (24) secures a thin sheet liner (16) to a supportive substrate (10) for protection against an adjacent high temperature environment (12). Cooling air (20) flows through a plurality of openings (34) and impinges (36) on the fastener (24). The fastener (24) is received within a dimple (32) in the liner (16) and includes an enlarged head portion (38).

Abstract: A family of airfoil cross sections, termed SC21xx, for use in a helicopter blade is disclosed. The airfoil (36) achieves maximum lift performance equivalent to prior art airfoil configurations without incurring increased aerodynamic drag at high velocities. The airfoil (36) was developed by thickening and drooping the leading edge region (38) of the prior art airfoil (30) improving lift in the leading edge region (38) and delaying the formation of sonic shock waves at high velocities.

Abstract: A sheet metal seal (42) provides both radial and axial sealing of the gap (34) formed between adjacent blade platforms (30, 32) in a rotor assembly (10) of a turbomachine.

Abstract: A heat management system is provided for a gas turbine engine (10) having first and second oil cooling loops (14, 16). The system distributes excess fuel flow from a main fuel pump (44) among a plurality of upstream locations (58, 60, 68) for managing the transfer of heat between the oil loops (14, 16) and the flowing fuel. A diverter valve (62) regulates the distribution of the bypass fuel responsive to engine heat generation, oil temperature, and/or fuel temperature. A passive fuel distribution configuration using one or more fuel flow restrictors (72, 74, 76) is also disclosed.

Abstract: A stator vane (4) for a gas turbine engine having a varying chordal length (52, 54) is disclosed. The chordal length reduces a maximum (52) at a point intermediate the ends (40, 44) of the vane for defining a radially varying nozzle throat (58, 56) between adjacent vanes (4, 4a) in a stator vane stage.

Abstract: A transverse pivotable flap is cantilevered between movable sidewall members (10) which include a stub flap (14). A removable central flap portion (18) is secured between the stub flats (10) by a pair of joints (26) having alternating lug sets 28, 30, secured by respective elongated pins (32, 34). Flow area (40) is provided to conduct internal cooling air (16) into the central flap portion (18).

Abstract: A manifold (20) includes a plurality of separate, identical flow channels (28) separated by flow dividers (46). The flow dividers (46) include a thickened boss section (54) for receiving a mounting bolt (50), and are skewed adjacent the manifold outlet for forming a plurality of tangentially directed nozzles (18).

Abstract: A swept leading edge blade (10) rotates in a stream of air (26) and experiences supersonic relative air velocities over at least a portion of the blade surface. The maximum camber points (28, 40) of adjacent blade segments (20, 18) are located to prevent intensification of the compression wave generated by each segment (20, 18).

Abstract: A nacelle structure (14) for a gas turbine engine (4) extends upstream from the engine air inlet (10) and includes an external surface (18) and an internal surface (16). The external surface (18) is concentric about a linear centerline (22) which is disposed parallel to the free airflow stream (6) encountered by the nacelle (14) in flight. The internal surface (16) is concentric about a curved centerline (30), the curved centerline (30) being parallel to the free airflow stream (6) at the nacelle inlet (12) and parallel to the engine centerline (26) at the engine air inlet (10) for smoothly directing the internal airflow stream (32) into the engine air inlet (10).

Abstract: A system (38) and method for directing control actions to recover from the occurrence of a surge condition within a turbocompressor includes a high/low engine power output discriminator (40). Control action is taken by the system (38) in response to both the detection of a surge condition and the current engine power output level. The system (38) selects among a plurality of corrective control actions (56, 62, 68) to best achieve a return to normal engine operation.

Abstract: A method for distributing a flow of fuel over the gas flow area of a gas turbine engine includes the steps of dividing the gas flow area into a plurality of sector shaped sub-areas (48), (50), (52), (54) disposed within a surrounding, annular sub-area (46) which is in turn disposed adjacent the augmentor wall (12). During operation, the method according to the present invention distributes the fuel over one or more of the sub-areas (46-54) in response to the current rate of fuel flow. At the lowest and intermediate fuel flow rates, fuel is distributed only over one or more of the sector shaped sub-areas (48-54). Only at the very highest rates of augmentor fuel flow is fuel distributed finally in the outer annular sub-area (46).

Abstract: An actuator assembly for imparting non-proportional tangential displacement (22b) to a plurality of unison rings (16b, 16c) disposed about the exterior of a compressor case (10b) is provided. The assembly includes a linear drive component (26b) mounted on trunnions (74, 76) and supported by a frame 48. The drive component (26b) imparts a rotating motion to a crankshaft 70 which in turn drives the unison rings (16b, 16c) via the respective crank arms (42b, 42c) and pushrod (30b, 30c) linkages. Radial loading of the compressor case (10b) is avoided by aligning both the pushrod (30b) and the elongated frame first end (50) tangential to the compressor case (10b).

Abstract: A two-dimensional nozzle for directing the exhaust of a gas turbine engine includes two spaced apart vertical sidewalls (16), an upper flap assembly (18) and a lower flap assembly (20). The upper flap assembly (20) includes at least a dorsal flap (22) pivotable about a first fixed axis (24). The lower flap assembly (20) includes a ventral flap (28) centrally pivotable about a second fixed axis (30) for creating a dual throat exhaust flow path.