Abstract: A telemetry carrier ring for use in a gas turbine engine includes an annular support ring connected to the engine and an annular carrier ring coupled to the support ring, each ring exhibiting different growth characteristics in response to thermal and mechanical loading. The carrier ring is coupled to the support ring by a plurality of circumferentially spaced web members which are relatively thin in an engine radial direction to provide a predetermined degree of radial flexibility. the web members have a circumferential width and straight axial line of action selected to transfer torque and thrust between the support ring and the carrier ring without substantial deflection. The use of the web members with radial flexibility provides compensation between the support ring and the carrier ring since the carrier ring grows at a different rate than the supporting ring.

Abstract: A gel/solid bipropellant propulsion system employs fuel-rich combustion gs from a solid gas generator and an oxidizer gel in a highly efficient combustion chamber wherein the fuel-rich combustion gases and the oxidizer gel are each metered through a vortex injector into a combustion chamber to produce a hypergolic reaction. The solid gas generator (SSG) has a preferred composition of glycidly azide polymer (GAP). The GAP SSG is composed of GAP diol and/or triol polymerized with a di-or tri-function isocyanate, such as isophorone diisocyanate. The gel/solid bipropellant propulsion system comprises the SSG in combination with an oxidizer storage/extrusion vessel system, a combustion chamber system, and a system controller which controls initial ignition of the SSG to produce fuel rich combustion gas which pressurize the system. The system controller monitors pressures and flow rates of fuel and gel oxidizer.

Abstract: A system and method of separating the gaseous components of air to provide a compressed oxygen-enriched gas stream and a nitrogen-enriched gas stream are disclosed. The separation system and method generate an oxygen-enriched stream including substantially pure oxygen, such as about 99% or greater oxygen, and having a pressure of up to about 10,000 pounds per square inch, thereby eliminating the need for a separate compressor to compress the oxygen-enriched gas stream. The oxygen separation system and method are used in aircraft to supply substantially pure oxygen that facilitates in-flight operation of an aircraft emergency power unit, especially at high altitudes.

Abstract: A plunger pump for supplying solution in accordance with an extrusion action of a plunger within a pump chamber. The plunger pump includes a cleaning chamber sealed by a seal member around and relative to the plunger, with a cleaning liquid inlet-side flow conduit and a cleaning liquid outlet-side flow conduit for communicating with the cleaning chamber, respectively. A mechanism isolates one of the inlet- and outlet-side flow conduits from the atmosphere and releases the other to an atmosphere under a condition where the cleaning liquid inlet- or outlet-side flow conduits are filled with the cleaning liquid. If the seal member in sliding contact with the plunger is degraded or aged, it is possible to observe a leakage of the liquid from the pump chamber along a sealing face between the seal member and the plunger.

Abstract: A removable section of duct is sealingly coupled to a relatively fixed length of duct by means of a flexible, preferably elastic, sleeve which is attached to one of the lengths of duct towards one end thereof and at its other end to a relatively stiff collar which is shaped to fit closely over the adjacent end of the other duct. The elasticity of the sleeve tends to retain the collar in position and to maintain a tight seal while permitting a degree of relative movement including longitudinal thermal growth.

Abstract: A combustor effective for reducing exhaust emissions is disclosed. The combustor includes first and second liners defining a primary combustion zone, and a plurality of circumferentially spaced carburetors. Each of the carburetors includes a fuel injector providing fuel into an air swirler for mixing the air and fuel and providing a fuel/air mixture into the combustor primary zone. The combustor also includes means for obtaining a fuel distribution from each of the carburetors extending radially from a first liner upstream end through a center region of the combustor to a second liner upstream end with values of the fuel distribution at the center region being generally no greater than about values of the fuel distribution adjacent to at least one of the first and second liners. The first and second liners have upstream portions characterized by the absence of film cooling thereof.

Abstract: A gas turbine combustor and a method of operating the gas turbine combustor, with the gas turbine combustor comprising a combustion sleeve, a combustion sub-chamber disposed at an upstream side of the combustion sleeve and having an air and fuel supply system, and a combustion main chamber disposed on a downstream side of the sub-chamber and having an air and fuel supply system. The combustion sub-chamber and combustion main chamber are formed in such a manner that the start-up of the gas turbine is effected by the hot combustion gas generated in the sub-chamber. The combustion sub-chamber is provided with an inner sleeve protruding from a side wall of the sub-chamber such that a tip of the inner sleeve is located in the vicinity of a downstream end of the sub-chamber, and a fuel nozzle is provided in a tip of the inner sleeve for increasing the fuel-to-air ratio during an initiating period of an increase in the gas turbine rotation speed.

Abstract: A method and system (10) for rerouting the pressurized oil flow of an engine (not shown) from a primary filter (14) to a secondary filter (18) in the event of a first sensed pressure differential across the primary filter (14) which exceeds an allowable pressure differential, the excessive differential indicative of a primary filter (14) obstruction. In the event of a subsequent second sensed pressure differential which exceeds a second and higher pressure differential across the secondary filter (18), the oil flow is then diverted through a pump relief valve (19) toward the oil reservoir (26), away from an engine bearing compartment (50). The system (10) also provides sequenced first and second level warnings (40, 42) to the controller/flight crew, said warnings responsive to the operating status of the system (10).

Abstract: A blow-off device for a gas turbine bypass engine having a primary air flow path separated from an outer secondary air flow path by a partition includes a plurality of discharge passages passing through the partition from the primary air flow path to the secondary air flow path, each of the discharge passages having a first hinged flap in the outer wall of the primary air flow path, a second hinged flap in the inner wall of the secondary air flow path, and a connecting rod mechanically interlinking the first and second flaps, the connecting rod being hinged at one of its end to the first flap and at its other end to the second flap.

Abstract: A solid propellant component grain is supported in a combustion chamber. A liquid propellant component container is mounted forward of the combustion chamber. The liquid propellant component is supplied through conduits from the container into the combustion chamber, and ignited to form combustion gas which is discharged out the rear of the combustion chamber to generate thrust for propelling a rocket. A high pressure tank containing a non-flammable gas such as helium is at least partially embedded in the grain. A conduit leads from the tank into the container, such that the high pressure gas pressurizes the container and urges the liquid propellant component to flow from the container into the combustion chamber. The tank provides internal structural support for the grain, with the wall of the combustion chamber constituting a safety barrier in the event of structural failure of the tank after the tank is filled with high pressure gas.

Abstract: An axial flow compressor with an inner casing and an outer casing. The inner casing has bleed apertures. The outer casing is provided with loading devices which are positioned adjacent to the bleed apertures such that the loading devices apply loads on the inner casing in the vicinity of the bleed apertures. Components of the loads applied by the loading devices are arranged to oppose and preferably balance any loads applied on the inner casing due to the bleed apertures' releasing the pressure differential across the inner casing. The loading devices comprise cylinders secured to the inner casing and pistons secured to the outer casing to form chambers. The chambers are supplied with fluid at a predetermined pressure. The loading devices oppose undesirable local reductions in clearance between static shrouds and the rotor blades.

Abstract: An apparatus and method are provided for cooling the rotating blades in the turbine section of a gas turbine. Cooling of the leading edge is accomplished by forming a radial passageway in the leading edge portion of the blade airfoil. Rows of holes along the leading edge of the blade airfoil allow cooling air to flow from the radial passageway to the surface of the leading edge, thereby cooling the leading edge portion of the airfoil. The flow area of the radial passageway reduces as it extends radially outward so that the velocity of the cooling air flowing through the passageway is maintained. A plenum formed in the root portion of the blade distributes cooling air to small diameter holes in the center and trailing edge portion of the airfoil.

Abstract: A simplex air blast fuel injection system and method for the atomization of fuel for ignition to drive a gas turbine includes a simplex nozzle which receives a fuel from a fuel pump powered by the turbine over a range of pressures between maximum and minimum pressures during the operation of the turbine, and also received fuel from the fuel pump at a substantially lower pressure than the minimum fuel pressure when the turbine is cranked during startup. The fuel is discharged from the nozzle orifice as a swirling stream of atomized fuel during turbine operation, and as a film which is insufficiently atomized to initiate ignition during the turbine startup. An air compressor is also powered by the turbine to supply air at a low pressure and high volume to the fuel issuing from the nozzle orifice during both turbine operation and startup.

Abstract: A waste heat recovery boiler system has a plurality of waste heat recovery boilers each having a main steam generator and a reheater capable of generating main steam and reheat steam by heat derived from a hot waste gas introduced into the waste heat recovery boiler. The flows of main steam from the main steam generators are joined to form a single flow of main steam to be used in an external installation, and the flows of reheat steam from the reheaters are joined to form a single flow of reheat steam to be used in the external installation. The steam discharged from the external installation is distributed to the reheaters. Flow rate controllers are provided for controlling the flow rates of the steam through the reheaters in relation to the hot waste gas supplied to the respective waste heat recovery boilers.

Abstract: A turbofan gas turbine engine (10) comprises a fan rotor (32) driven by a core engine (11) via a shaft (36) and a reduction gear (38). A lubricant source (63) is arranged to supply lubricant to the reduction gear (38) to lubricate gear teeth and bearings (52) in the reduction gear (38). An annular collector (76) is positioned coaxially around the reduction gear (38) and is secured to the fan rotor (32) and is driven by the reduction gear (38). The annular collector (76) is arranged to collect lubricant leaving the reduction gear (38) on its inner surface. A number of static scoops (82) collect lubricant from adjacent the inner surface of the annular collector (76) to return the lubricant to valves (86). The valves (86) allow the lubricant to return to the lubricant source (63) in normal operation.

Abstract: A gas turbine engine with counterrotating fans placed upstream from the engine so that gas circulates from upstream to downstream at the level of a LP compressor (1), then from downstream to upstream in a HP compressor (5), chamber an (6) and HP and LP turbines (7 and 8), then again downstream in first and second turbines (10, 11) before being expelled into the cold airflow coming from first and second fans (14, 15).

Abstract: A differential gear assembly which is suitable for providing a power off-take from a gas turbine engine, comprises first and second power input gears which drive a plurality of idler gears rotatably mounted on a common cage structure. The idler gears are grouped in meshing pairs so that one idler gear of each pair is additionally in meshing engagement with one of the power input gears and the other idler gear of each pair is additionally in meshing engagement with the other of the power input gears. Together the first and second power input gears drive the common cage structure at the mean speed of the power input gears and the common cage structure in turn drives a single power output shaft.

Abstract: A rocket engine combustion chamber is defined by a generally cylindrical member comprising a metal matrix and heat insulating particles wherein the particles are distributed in the matrix such that the amount of particles gradually decreases in a radial direction from the inner surface to the outer surface of the member. Such a member is manufactured by controlled composite electroplating.

Abstract: A gas turbine engine intended for use as an aircraft engine which includes a gas generator which supplies interdigitated, contra-rotating radially inner and outer low-speed power turbines driving forward mounted contra-rotating front and rear fans, the gases flowing through the gas generator from the rear of the engine towards the front, and then flowing in a rearward direction through the power turbines. The front fan is driven by the radially inner power turbine via a central shaft extending towards the front of the engine and supporting a disc carrying the front fan, and the rear fan is driven by the radially outer power turbine via a forwardly extending drum which is rotatably mounted between the central shaft and a stationary structure of the engine and which carries the rear fan. This structure permits the pod mounting of a high by-pass ratio engine under wing of an aircraft.

Abstract: A hinge for pivotably connecting upstream and downstream wall sections of an exhaust nozzle of a gas turbine engine. Each wall section includes a liner spaced from the interior surface of the wall section to define respective first and second cooling air flow passages therebetween. One of the wall sections includes a curved end portion and the hinge includes a leaf seal which extends from the other wall section and biases against the curved end portion to form an air tight seal and define, at least in part, a plenum at the hinged connection of the wall sections. The plenum provides an air flow communication path between the first and second cooling air flow passages to transfer cooling air from the air flow passage of the upstream wall section to the air flow passage of the downstream wall section to thereby increase the efficiency of cooling of each of the wall sections.