Abstract: An external fairing flap (26) is provided with an hydraulic spring (42, 44) for causing the external flap (26) to follow the motion of an exhaust nozzle control flap (18). The external flap (26) is hinged (32) at the leading edge thereof, and contacts the control flap (18) via corresponding rubstrips (40, 36) at the flap trailing edges (28, 38).

Abstract: A method and fixture (10) for assembling compliant seals (30) includes an aligning member (12) having a grooved surface (14) for receiving a plurality of loose bristle elements (16). The bristle elements (16) are received within the grooves (26) and aligned thereby. First and second backing rings (20, 22) sandwich the aligned bristles (16) and are clamped by first and second clamping members (18, 24). The clamping rings (20, 22) and bristle elements (16) are then removed from the aligning member (12) and permanently secured together, forming the finished seal (30).

Abstract: A segmented seal (36) is provided for sealing between two adjacent rotor disks (18, 20). The individual segments (36) includes sideplate members (42, 44) and integral hook members (50, 52) which engage the facing sides of the respective disks (18, 20). Strut members (60, 62) support the central portion (64) of the axially and circumferentially extending wall members (40) which collectively establish the radially inner gas boundary for an annular working fluid stream (30).

Abstract: A clamp assembly (32) for securing an elongated seal (30) between adjacent hinged flaps (12) includes a wearplate (60) and a bridge clamp (54). The wear plate includes a pair of integral tabs (64) for engaging a threaded mounting plate (35) for providing antirotation between the plate (60) and the seal (30). The wear plate (60) also includes at least two ears (68) disposed on opposite sides of one of the arms (70, 72) of the bridge clamp (54) to prevent rotation between the clamp (54) and plate (60).

Abstract: A variable area, convergent-divergent (10) nozzle for directing the flow of exhaust gas from an aircraft turbine engine to achieve thrust vectoring about yaw, pitch and roll axes includes a convergent section (16) and a divergent section (18) each comprised of hingedly connected planar members (201-214, 301-322) selectively positioned by actuators (101-114) for defining deformable gas directing flow ducts.

Abstract: A nozzle guide structure (10) is provided for receiving a fuel nozzle (34) in an opening (30) in a planar bulkhead (28) of a combustion chamber (12). The guide structure (10) includes an inner bushing (44), a transverse heat shield (46), and an annular retainer (52) disposed opposite the heat shield (46) with respect to the bulkhead (28). Flow openings (62) disposed in the retainer (52) admit cooling air (64) into an annular gap (50) formed between the bushing (44) and the bulkhead (28). the cooling air (64) subsequently flow between the bulkhead (28) and the heat shield (46) to provide cooling thereto.

Abstract: A visual deficiency screening system (8) includes a display (14) and a means (10, 18) for generating a series of randomly differing images (12). A hood (24) and eyepiece (26) are provided for properly positioning a test subject relative to the display (14).

Abstract: A temperature controlled rotating seal (46) includes an annular runner (48) having knife edges (50) and a surrounding static shroud (44). Baffles (64, 66) extend radially from the adjacent rotor (10) and a static stator vane assembly (36), forming a plurality of annular mixing volumes (60, 62) disposed upstream of the seal (46). Cool air (70) is provided to the innermost mixing volume (62) thereby controlling the local gas temperatures at the rotor periphery (14).

Abstract: A method for periodically restoring the blade tip to annular seal clearance in a gas turbine engine (10) includes incrementing (42a, 42b) the flow of external cooling air responsive only to an accumulated engine use parameter (58).

Abstract: A yaw and pitch thrust vectoring exhaust nozzle (10) includes upper and lower divergent flap assemblies (22, 24) each divided by oppositely skewed diagonal hinges (58, 60). Vertical sidewalls (27, 28) are also hinged to open the nozzle outlet laterally during yaw thrust.

Abstract: A rotor assembly (34) includes a disk (36) having an axially extending, unbroken annular spigot (38). The spigot (38) supports a plurality of radially extending lugs (42) which are juxtaposed with a corresponding plurality of dogs (14) integral with an annular sideplate (12). A smooth key (54) fits within the opening (44) formed between adjacent engaged dogs and lugs (14, 42) and is retained axially by a sheet metal retainer (46) received through a slot (52) disposed in the key (54).

Abstract: A method and apparatus for eliminating vortex whistle noise in a radial-to-axial compressor intake uses a plurality of flow fences (36) disposed along the radially inner gas passage boundary (10). The fences (36) disrupt a portion of the swirling gas flow (32) resulting from the position of a plurality of pivotal inlet guide vanes (28) disposed about the radially outward facing inlet (18).

Abstract: A gas turbine engine rotor (18) and shaft (28) are secured against radial loosening between the shaft (28) and axially spaced rotor spigots (24, 26) by annular spacer rings (42, 43) which are fabricated from a material having a thermal coefficient of expansion significantly higher or lower than the thermal coefficients of the shaft (28) or rotor (18). The ring (42, 43) is pre-assembled during production of the engine in an interference fit with either the rotor hub (22) or the shaft (28), being sized to remain tight with the pre-assembled member over a range of operating temperatures, and becoming more tightly fit with the other member due to differential thermal expansion therebetween.

Abstract: A yaw and pitch thrust vectoring exhaust nozzle includes a spherical collar, and two surrounding clam shells hinged to a gimbal ring and pivotal about an axis perpendicular to the clam shell hinge axis. The clam shells define a variable area nozzle throat (42) which can be selectively directed vertically by rotation of the clam shells about the hinge axis (38), or horizontally by rotation of the clam shells and gimbal ring about the pivot axis (30), thereby biasing the nozzle exhaust flow vector. Divergent flaps (44, 46) further direct the exhaust downstream of the nozzle throat (42).

Abstract: A seal assembly (28) is provided for sealing between a static sidewall (12), a rotating first member having a flat and annular surface (22) disposed therein, and a concentric rim (34) disposed between the sidewall (12) and first member (22). The rim (34) is supportably secured to an arcuate hood (18) which extends axially from the rim (34).

Abstract: A method and system for providing training mode operation of a multiple engine aircraft simulates the failure of one engine and the simultaneous operation of at least one other engine in an elevated, emergency power condition. The engines (20, 22) are mechanically linked (24) and provide a single power output (26). A training power schedule (56a, 56b) is selected (60, 62) to limit the combined engine power output to a level equivalent to the combined output with one engine inoperative and another in an elevated, emergency power operating condition.

Abstract: A plurality of cascade vanes (28, 29, 30) pivot in unison for selectably directing a flow of gas (20) therethrough. A flow regulating vane (36) disposed adjacent one end vane of the plurality of cascade vanes (28, 29, 30) pivots (46) counterrotationally with respect to the cascade vanes (28, 29, 30) for maintaining a constant collective gas flow area.

Abstract: A plurality of cascade vanes (28, 29, 30) for directing a flow of gas (20) is provided with at least one flow regulating vane (36) which counterrotates with respect to the cascade vanes (28, 29, 30) over a gas flow vectoring range of motion. The cascade vanes (28, 29, 30) and flow regulating vane (36) corotate when moving through an adjacent stowing range of motion for achieving a stowed, gas flow blocking arangement.

Abstract: A ducted fan gas turbine powerplant is mounted to aircraft structure via a single plane mounting which comprises a single peripheral array of struts which connect the core gas generator, the fan cowl and the aircraft structure. The structure which connects the struts to the core gas generator is extended downstream of the array of struts and terminates in a plain bearing. A downstream portion of the core gas generator includes a cylindrical portion which is in sliding engagement with the plain bearing, so that radial forces generated by the movement arm, the fulcrum of which coincides with the strut connection with the core gas generator are transmitting via the struts to the aircraft structure and relative axial expansion between the extension and the core gas generator is enabled.

Abstract: A distributed data acquisition system for a first engine controller determines a final, validated value of a critical flight condition variable by first determining: a first tentative value based on local environmental conditions, a second tentative value based on the performance of a first engine, and a third tentative value transferred directly from a similar data acquisition system linked to a second engine on the same aircraft. The data acquisition system compares the three tentative values and selects a final, validated value for use by the controller.