Abstract: The invention is a valve system particularly useful as a bleed valve system in a gas turbine engine. The valve system includes a valve ring or partition (44) that bounds a plenum (52). The valve ring has a set of two or more fluid flow regulating apertures (46), and one or more chains (60L, 60R) of metering plates (56). Each metering plate is pivotably supported relative to one of the apertures (46) to regulate fluid flow through the aperture. A coupling system that includes a series of transfer links (76) extending between neighboring metering plates successively conveys pivotal motion and angular orientation from each metering plate to the successive metering plate in the chain. An actuation system (90) includes a drive system (94) responsive to an actuator (92) for governing the angular orientation of the metering plates.

Abstract: A method for preparing a protectively coated, apertured article to be recoated reduces the likelihood that the apertures will become constricted as a consequence of subsequent recoating. The invention, described in the context of a gas turbine engine blade (12) having transpiration cooling passages (34), includes the step of diffusing an auxiliary coating (52) into an existing coating (28) and into the exposed substrate material (26) at the periphery of the passages to form a diffusion zone (54). The blade is then subjected to a stripping agent so that the diffusion zone (54) any undiffused existing coating, and any undiffused auxiliary coating are removed. The method causes a compensatory enlargement of the mouth (44) of each passage so that excess coating that accumulates in the passage mouths during subsequent recoating does not restrict the flow capacity of the passages.

Abstract: The invention is a method and apparatus for combining two fluid streams, for example concentric inner (42) and outer (44) fluid streams in a turbine engine, thereby reducing noise emissions of the engine. A mixing system (12) according to the invention has a convoluted mixer body (18) and a fluid communication path (88) situated forwardly of a main mixing zone (54) for inducing a small quantity of fluid to flow from one of the streams and into the other of the streams in response to a differential pressure between the streams. In one specific embodiment, the fluid communication path (88) is a series of circumferentially distributed crossover apertures (92) forwardly of the leading edge (48) of the mixer body (18). In operation, the mixing system circumferentially interleaves the inner and outer streams of a turbine engine and introduces a quantity of fluid from one of the streams into the other of the streams, the fluid introduction occurring prior to or concurrently with the interleaving.

Abstract: The invention is a method for repairing a protectively coated, locally damaged turbomachinery vane without removing the vane from its parent turbomachinery module. The method includes applying a diffusible coating precursor to the damaged region of the vane and diffusing the coating precursor into the vane in the presence of a nonreactive atmosphere. The nonreactive atmosphere is confined to a chamber that contains the damaged vane, the chamber being bounded in part by the turbomachinery module itself. A conventional stress relief machine is used to heat only the damaged region of the vane to a temperature sufficient to promote diffusion of the precursor into the vane. The method uses equipment and supplies that are routinely available or easily fabricated from common, inexpensive materials. Use of the method avoids the expense and delay of removing the vane from its parent module.

Abstract: A turbine blade (18) for a gas turbine engine includes a machining datum (60) that extends radially from a pocket (54) in the outer shroud (36) of the blade. The datum is spaced from a sidewall (64) of the pocket so that the datum is peripherally continuous irrespective of whether the blade is in a prefinished state or in a completely finished state. Because the datum's peripheral continuity survives the original manufacturing process, the datum is available for use in post-manufacturing inspection and repair operations.

Abstract: The blades (38) of a gas turbine engine fan rotor (32) are uniquely distributed in the rotor hub (48) so that combination tone noise attributable to one or more physical nonuniformities in the blades is mitigated. The invention recognizes that the frequency spectrum of the combination tone noise includes both decay prone harmonics and decay resistant harmonics and that allocation of acoustic energy into the decay prone harmonics and out of the decay resistant harmonics is effective for mitigating combination tone noise. In particular, the blades are distributed so that the spatial or circumferential spectrum of the physical nonuniformity is dominated by circumferential harmonics whose order is no greater than the order of the highest decay prone harmonic of the combination tone noise spectrum.

Abstract: A blade or vane for a gas turbine engine includes a primary cooling system (42) with a series of medial passages (44, 46a, 46b, 46c, 48) and an auxiliary cooling system (92) with a series of cooling conduits (94). The conduits of the auxiliary cooling system are parallel to and radially coextensive with the medial passages and are disposed in the peripheral wall (16) of the airfoil between the medial passages and the airfoil external surface (28). The conduits are chordwisely situated in a zone of high heat load (104, 106) so that their effectiveness is optimized. The conduits may also be chordwisely coextensive with some of the medial passages so that coolant in the medial passages is protected from excessive temperature rise. The chordwise dimension C of the conduits is limited so that potentially damaging temperature gradients do not develop in the airfoil wall (16).

Abstract: A method for reliably starting a gas turbine engine includes commanding a rapid transient enrichment of the fuel quantity delivered to the engine combustion chamber, maintaining the fuel quantity at the enriched level while operating the engine's ignition system, and, upon the expiration of a fixed time interval or an indication of successful ignition, commanding a derichment to a fuel quantity equal to that which would be obtained during a conventional startup procedure. The degree of transient enrichment depends on the operating environment and operational state of the engine. In one embodiment of the invention the rate of enrichment corresponds to a virtual step change in the commanded fuel quantity and the rate of derichment is less than the rate of enrichment.

Abstract: A variable exhaust nozzle (26) for a turbine engine nacelle (16) includes a pair of semi-cowls (58, 60) disposed about a longitudinally extending central axis (14) and a pair of corresponding shells (36, 38) each of which is spaced radially inwardly from the corresponding semi-cowl and is pivotable about a pivot axis (42, 44) between an extended position and a retracted position. The semi-cowls and shells have curved, geometrically similar inner and outer surfaces (86, 100) respectively so that gap (108) between the shells and the cowl is uniform irrespective of the angular orientation of the shells. A seal (128) which includes longitudinally extending legs (128a, 128b) and an optional circumferential leg (128c) prevents detrimental air leakage. Pivotal movement of the shells effects large changes in the discharge area A.sub.D of the nozzle without compromising the external aerodynamic characteristics of the nacelle.

Abstract: A containment case (130) for a turbine engine circumscribes an array of rotatable blades and comprises a penetration resistant containment ring (152) and an impact isolator (138a, 138b) extending from the containment ring. The containment case has an impact zone (136) which is the region where any blade fragment which separates from a blade is anticipated to strike the containment case. The containment ring is axially coincident with the impact zone, and the impact isolator is axially spaced from the impact zone. In one embodiment, The impact isolator is a circumferentially extending convolution. During operation, the impact isolator protects the integrity of a connection (139, 141) between the containment case (130) and an adjacent case (62, 64) and reduces the severity of forces transmitted to components (108) mounted on the case.

Abstract: A cooled rotor assembly for a gas turbine engine includes rotor blades (48) and a plurality of cover plates (86a) retained in the blade attachment slots (34) of a rotor disk (30). Each blade has a platform (46) with a trailing edge (58), and each cover plate includes at least one conduit (120, 126) for drawing leakage air from a plenum (102) in front of the disk and impinging that air against the platform trailing edges (58) to cool the tailing edges. In the preferred embodiment, each conduit (120) is defined by the rear face of a blade root (40) and a groove (122) cast into the cover plate. The invention makes productive use of leakage air by using that air to cool the platform trailing edges rather than allowing it to flow unbeneficially into the engine gaspath (24) in the vicinity of the platform leading edges (56).

Abstract: A fixture (90) for use in selectively applying a protective coating to a gas turbine engine blade (10) includes a base (92) having a blade retention slot (102), and a disposable shield (110). The shield (110) includes a sheet metal sleeve (112) and a cap (114) having a window (116) the cap being affixed to one end of the sleeve. When fully assembled with the blade properly positioned in the fixture only those portions of the blade selected to be coated are exposed. Additional features such as a locator dowel (104) extending from the floor of the slot and mutually cooperative track (121) and nub (122) assist in locating the blade correctly in the fixture and orienting the window with respect to the blade platform (16). A lock exemplified by a groove (128) and the nub (122) resists separation of the shield from the base during coating operations.

Abstract: A support arrangement for a rotor shaft (10) includes a bearing (28) radially interposed between the shaft and a support frame (26). The bearing has a durable support (100) and a destructible support (104). The destructible support is nested within a circumferentially extending channel (78) having three-dimensionally curved surfaces (82, 84). An adjacent oil seal (44) has similar three-dimensionally curved surfaces (52, 54) In the event that the rotor becomes unbalanced during operation, the destructible support disengages from its housing (76) and glides radially within the channel to restrain the rotor longitudinally while precluding the transmission of imbalance forces from the rotor to the frame. The curved character of the seal surfaces preserves the integrity of the seal so that loss of oil will not prevent continued, low speed operation of the rotor. An optional, temporarily operable damper (148) attenuates especially severe vibrations immediately following the onset of the imbalance.

Abstract: A gas turbine engine (10) includes a translating sleeve (38) disposed within a downstream portion of an outer nacelle (20). A variable area fan exhaust nozzle (30) is defined between the trailing edge (32) of the translating sleeve (38) and a conical core cowl (26) disposed radially inwardly of the outer nacelle (20) and spaced apart therefrom. The translating sleeve (38) translates downstream to cooperate with the decreasing diameter of the core cowl (26) to increase the area of the fan exhaust nozzle (30).

Abstract: A method for suppressing noise in a turbofan engine includes segmenting coaxial inner and outer flow streams into circumferentially interleaved inner and outer segments beginning at a common location along the engine axis, diverting additional portions of the outer stream radially inwardly from an axial location downstream of the common location and combining the inner segments, the outer segments and the additional portions into a common stream so that the additional portions are introduced into discrete radial locations in the common stream. The diverted additional portions of the outer stream eliminate localized regions of hot, high velocity gases in the combined stream to achieve improved noise suppression.

Abstract: A gas turbine engine having coannular components 22, 30 includes means for receiving and retaining a replaceable brush seal 60 so that the seal bristles impede the leakage of air through an intercomponent gap 41. In the preferred embodiment one of the components is a hollow, cylindrical barrel 30 comprising upper and lower barrel segments and the receiving and retaining means includes a retainer 52. The retainer cooperates with the barrel to define a groove 56 within which the brush seal is trapped radially and axially. The brush seal slides circumferentially within the groove to facilitate seal installation and removal.

Abstract: A repair kit for replacing an unserviceable vane assembly in a turbine engine includes a serviceable vane subassembly 120 having a serviceable airfoil 124 with a base 134 attached to the root end thereof and an opposing base 156 which is unattached to but slidably engagable with the serviceable airfoil. The opposing base is slipped over the tip end of the airfoil and slid toward the opposing base. The vane subassembly and opposing base are pivoted into position between inner and outer engine cases in place of an unserviceable vane assembly so that receptor sockets 148, 150 in the attached base engage support pins 72 extending from the inner case. An adhesive 182 is applied to the tip end of the serviceable airfoil and the opposing base is translated into its installed position near the tip end of the airfoil. Upon curing of the adhesive, the opposing base becomes attached to the airfoil. The base is secured to the outer case by studs 196 and nuts 198.

Abstract: An afterburner 20 for a gas turbine engine 10 has a fuel spray ring 24a for injecting fuel into the afterburner, a flameholder gutter 34a for stabilizing combustion of a fuel-air mixture flowing through the afterburner, and an ignitor 35 for initiating combustion and includes an enclosure 50 attached to the gutter. The enclosure has radially inner and outer walls 52, 54 and circumferentially spaced apart webs 60, 62 extending between the walls to define a radially and circumferentially bounded chamber 64. Each web has a forward opening 72 and an aft opening 74 so that a portion of the spray ring and a portion of the gutter are embraced by the enclosure. The enclosure is ideally circumferentially aligned with the ignitor and regulates the fuel-air ratio within and in the vicinity of the chamber to ensure reliable lighting of the afterburner and flawless advancement to full afterburning operation.

Abstract: An array (32) of flow directing elements such as fan blades (20) for a turbine engine (10) is disclosed. Each blade in the array has a set of circumferentially shrouds (42, 44) whose contact edges (46, 48) abut each other during engine operation. A shroud angle .sigma. establishes the running chord angle .alpha. of the airfoil (30) portion of the array. Slight differences in the shroud angles of adjacent blades cause corresponding differences in the running chord angles so that the array is detuned and airfoil flutter is precluded.

Abstract: A swept turbomachinery blade for use in a cascade of such blades is disclosed. The blade (12) has an airfoil (22) uniquely swept so that an endwall shock (64) of limited radial extent and a passage shock (66) are coincident and a working medium (48) flowing through interblade passages (50) is subjected to a single coincident shock rather than the individual shocks. In one embodiment of the invention the forwardmost extremity of the airfoil defines an inner transition point (40) located at an inner transition radius r.sub.t -inner. The sweep angle of the airfoil is nondecreasing with increasing radius from the inner transition radius to an outer transition radius r.sub.t-outer, radially inward of the airfoil tip (26), and is nonincreasing with increasing radius between the outer transition radius and the airfoil tip.