A gas turbine engine comprising a core gas generator engine for generating combustion gases, an additional power turbine, an additional fan section, and a booster compressor is disclosed. The power turbine includes first and second counterrotatable turbine blade rows effective for rotating first and second drive shafts, respectively. The fan section includes a first fan blade row connected to the first drive shaft and a second fan blade row connected to the second drive shaft. The booster compressor includes a first compressor blade row connected to the first drive shaft and a second compressor blade row connected to the second drive shaft.

High bypass ratio counterrotating turbofan engine

Film coolant passages through the wall of a hollow airfoil for a gas turbine engine are oriented to direct coolant fluid therefrom at a substantial acute angle to the downstream direction and at a shallow angle to the surface of the airfoil. Each passage includes a metering portion at its inlet end and a four sided diffusing portion at its outlet end. The diffusing portion includes a pair of adjoining surfaces which are both parallel to a central axis of the passage, and another pair of adjoining surfaces which diverge from the central axis, the diverging pair of surfaces being located on the downstream side of the passage. With this configuration the coolant completely fills the diffusing section and exits as a wide film of coolant within the boundary layer downstream of the passage outlet.

Cross-flow film cooling passages

A film cooling passage through the external wall of a hollow airfoil for a gas turbine engine has a metering section in series flow relation with a diffusing section leading to the passage outlet at the external surface of the airfoil over which a hot gas flows during operation. The diffusing section is generally rectangular in cross section perpendicular to the axis of the passage. Upstream and downstream longitudinally extending, spaced apart, facing surfaces of the diffusing section are joined together by a pair of spaced apart side walls which face each other and diverge from each other in the longitudinal direction toward the outlet of the passage. These side walls blend, along their length, with the downstream surface as a smooth curve of large radius. The diameter of the corner curvature of the passage is on the same order of magnitude as the distance between the upstream and downstream surfaces of the diffusing section at the location in question. The large corner curvatures, as opposed to filet radii, allow the side walls to diverge from each other at a greater angle without separation of the coolant fluid therefrom, permitting the spreading of smaller amounts of coolant air over a larger area of the airfoil external surface.

Film cooling passages with curved corners

A generally annular combustor system has a chamber with an annular chamber portion and a plurality of can-like chamber portions protruding from the annular chamber portion. An external fuel/air premixer is associated with each can chamber portion to provide a mixture with a desired (generally lean) fuel/air ratio tangentially through the can sidewall. A first portion of the total compressed air flows to the premixers, a second portion enters the combustion chamber through ports which, for can chamber volumes sufficient for complete combustion, can be located in the can chamber portion proximate the connection between the can portion and the annular portion. The premixer associated with each can chamber portion preferably includes a venturi and a fuel nozzle for spraying liquid fuel or gas into the venturi inlet to provide a fully premixed fuel/air mixture. An auxiliary fuel nozzle can be positioned in the annular portion for selectively supplying fuel for reheat of the combustion gases, using the dilution air, to maintain combustor outlet temperature. The annular chamber portion advantageously can be nested between the compressor and turbine components of large axial flow-type gas turbines in retrofit applications.

Star combustor with dilution ports in can portions

An auxiliary fan is disposed upon the periphery of at least a portion of the fan of a conventional gas turbofan engine. Flow modulating components such as variable vanes and nozzles are utilized in conjunction therewith to produce a wide range of operating bypass ratios. Such an engine is one of the variable cycle variety and has the capability of operating efficiently at both subsonic and supersonic speeds.

Two-spool variable cycle engine

1,235,006. Turbines; axial flow compressors; jet propulsion plant. SOC NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION. Dec. 5, 1968 [Dec. 12, 1967], No. 57836/68. Headings F1C, F1J and F1T. A two-tier blade ring for a turbomachine such as a compressor ducted fan or the like comprises angularly spaced blades each formed by an inner section 2 having a tip portion 5 and a radially aligned outer section 3 having a base portion 6a mounted in a groove 6 in the tip 5. The blade sections 2, 3 are exposed to different working fluid flows, such as two air-flows or flows of air and combustion gases, and are separated by an annular member 1 having circumferentially spaced apertures engaging the tips 5 to strengthen the assembly. Insertion of each base 6a in its groove 6 is facilitated by a groove 7 in the member 1 in alignment with each aperture. As shown, the blades may be variable pitch having a spigot 4 secured in taper-roller bearings 8 in an annular member 9 fixed to the shaft r of a gas turbine engine compressor. Each spigot 4 carries Fig. 1, a gear 10 meshing with a rack 11 on an arm 12 secured to an axially slidable sleeve 13 mounted on shaft r. A hydraulic jack or other pneumatic, mechanical or electrical device can actuate the sleeve 13 to vary the blade pitch. In a modification, Fig. 6, the pitch varying mechanism comprises a toothed sector 21 secured to each spigot 4 and meshing with a gear 22 mounted on shaft r. Normally gear 22 rotates at the same speed as shaft r but can be rotated relative thereto to effect variation of the blade pitch. Fig. 5 (not shown) illustrates an arrangement in which two succesive two-tier compressor rotor stages have pitch varying mechanism as shown in Fig. 1, the respective control sleeves 13 being slidable in mutually opposite directions by a common hydraulic actuator (16) fixed to stator structure between the rotor stages. The two-tier stator blade rings (15) of the compressor may also be constructed similar to Fig. 1 with fixed or variable pitch blades.

Improvements in or relating to blading arrangement for turbomachines

1,233,718. Gas turbine ducted fan engines; driving compressors. SOC. NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION. Dec. 11, 1968 [Dec. 14, 1967], No.59031/68. Headings F1C, F1G and F1J. A gas turbine engine comprises an axial-flow compressor and a turbine, the compressor being divided into n co-axial units where n is two or more, each compressor unit comprising two contra-rotating rotors, and the turbine comprises (n+1) independent rotors each driving a transmission shaft, the turbine rotating in opposite directions considered from one to the next, each turbine rotor driving two rotors of two successive compressors, except for the first and last turbine rotors which each drive only one compressor rotor. In the gas turbine ducted fan engine shown in Fig. 1 the axial flow compressor assembly 7 comprises an L. P. contra-rotating compressor 11 and an H.P. contra-rotating compressor 12 and the turbine assembly 9 comprises three independent turbine rotors 13, 14, 15 each driving a shaft 26, 27, 28. The engine comprises also a ducted fan 2 upstream of the L.P. compressor. The H. P. turbine 13 is connected by shaft 26 to drive the inner rotors 22, 24 of the H.P. compressor. The L. P. turbine 15 is connected by shaft 28 to drive the outer rotor 19, 21 of the L. P. compressor, also the ducted fan 2. The I, P, turbine 14 is connected by shaft 27 and the shaft 27a which are inter-connected by spline coupling 31, to drive the inner rotor 18, 20 of the L. P. compressor also the outer rotor 23, 25 of the H, P, compressor. The bearings for the shafts are indicated diagramatically. The ducted fan engine shown in Fig. 2 again comprises three turbine stages 13, 14, 15 and two contra-rotating compressor stages 11, 12 but in this case the H. P. compressor does not comprise a rotatable outer drum or casing. The H. P. turbine 13 drives through shaft 26 an internal ring gear 58 which drives a gear 56 which in turn drives a shaft 49 carrying gears 48 which engage internal ring gears 49 carried by the compressor discs 42 which in turn carry blades 40. The L. P turbine 15 drives through shaft 28 the outer rotor 19, 21 of the L. P. compressor also the ducted fan 2. The I. P. turbine 14 drives through shaft 27 the inner rotor 18, 20 of the L. P. compressor, also through gearing 59, 57 shaft 53 and gearing 52, 47 the discs 43 and blades 41 of the I. P. compressor. In the assembly three gears 56 engage with the internal ring gear 58 each gear 56 driving a shaft 49 and gears 48; similarly three gears 57 engage the internal ring gear 59 each gear 57 driving a shaft 53 and gears 52.

Gas turbine power plants having axialflow compressors incorporating contrarotating rotors

Two-flow gas turbine jet engine

A burner for the reheat chamber of a dual-flow gas turbine jet engine, comprising a fuel-prevaporizing device having a feedpipe fed jointly from a liquid fuel source and a secondary air duct and having a discharge pipe opening into the reheat chamber to discharge thereinto a mixture of air from the secondary duct and of fuel preheated by heat exchange with a primary flow of hot gases from the turbine.

Burners for reheat combustion chambers

A gas turbine unit, in particular for an aircraft turbojet engine, comprises a fan rotating coaxially in front of an axial flow compressor. The fan has an outer blading ring fixed to a rim attached to a hub of the fan through an inner blading ring, and the inner blades are pivotally attached, at the root and tip ends respectively, to the hub and to the rim about pivot axes parallel to the axis of rotation of the hub. The inner blades are all inclined in the same direction and at the same angle in relation to the hub circumference. The rim can thus expand freely under the effect of the centrifugal forces applied to it in operation by the outer blading, without local distortion and without the application of any bending stresses to the inner blades.

Armand Jean-Baptiste Lacroix

Papers

Improvements in or relating to blading arrangement for turbomachines

Gas turbine power plants having axialflow compressors incorporating contrarotating rotors

Two-flow gas turbine jet engine

Burners for reheat combustion chambers

Fan for gas turbine unit