Abstract: A thrust reverser system for fan jet type aircraft gas turbine engines using a rack and pinion arrangement for moving blocker doors between the stowed and deployed positions. Blocker doors fill a plurality of circumferentially arranged openings.

Abstract: A turbopump system for pumping two separate fluids is disclosed having a r of concentric turbine rotors located within a turbine casing, each directly connected to a pump impeller. The centrifugal pumps, which may be located on opposite sides of the turbine casing are separately connected to their respective fluids such that the two fluids may be pumped simultaneously. A propellant fluid is directed through a generally annular space defined between the concentric turbine rotors to rotate them in opposite directions. Turbine blades attached to each of the turbine rotors extend into this annular space in order to drive the turbine rotors. The pitches of the turbine blades are such that the rotors are driven in opposite directions and one of the turbine rotors may be driven at a different speed than the other.

Abstract: A mechanism is disclosed, comprising: a cylinder-block; a cam; an internal fluid distributor slide valve presenting a recess whose shape is complementary of a central core; grooves made between internal distributor valve and central core; and seal elements disposed on either side of each groove. According to the invention, each seal element comprises a ring-type joint made of a plastics material, the maximum value of the pressures in the grooves being equal to 450 bars and that of the speed of rotation of the internal fluid distributor valve with respect to the central core, being equal to 200 rpm. One application of the invention is the production of a motor having a satisfactory volumetric efficiency.

Abstract: A gas turbine engine is provided with a thrust reverser. In order to avoid inadvertent operation of the thrust reverser when the gas turbine engine is not running, a safety switch is provided on the engine at a position remote from the thrust reverser. The thrust reverser control system is arranged so that when the engine is not running the thrust reverser will only operate when the safety switch is being manually operated.

Abstract: A method is disclosed for discharging combustion gases from an engine powering an aircraft at subsonic and supersonic velocities. The engine includes a variable area exhaust nozzle having primary and secondary flaps defining a generally converging channel and exhaust channel, respectively, for the combustion gases. The method comprises the steps of contouring the secondary flaps to a first contour at a first position at a first value of a nozzle pressure ratio NPR across the converging and exhaust channels. The method further includes the step of contouring the secondary flaps to a second contour at a second position at a second value of NPR. In a preferred embodiment, the first contour is straight at low values of NPR and the second contour is generally concave for high values of NPR.

Abstract: To eliminate the fish-mouth geometry at the downstream portion of a thrust reverser fitted jet engine which results from the formation of substantially arcuate notches when the thrust reverser doors are stowed, moveable flaps are positioned in close proximity to the doors such that when the doors are in their stowed position, the flaps would fill, at least partially, the arcuate notches to effect a substantially planar exit opening, thereby substantially eliminating the fish-mouth geometry. Moreover, the flaps can be used to effect a uniform thickness for the exit opening of the jet engine by each having a shape corresponding to that of the corresponding arcuate notch; and the principal outer surfaces and inner surfaces of the flaps provide continuity with the outer surfaces and the inner surfaces, respectively, of the doors, with the ends of the flaps have trailing edge which have thicknesses similar to those of the trailing edges of the doors.

Abstract: An apparatus for influencing a flow of air through a passageway defined between an inner wall and a double outer wall that includes first and second walls each formed with an opening, comprises a first door, a second door, vortex generation means positioned between the first and second walls and adjacent the upstream end of the opening in the first wall, and an actuator pivotally attached to the first and second doors. The actuator has a first condition wherein the first and second doors close the openings in the first and second walls respectively and has a second condition wherein the first door is at the downstream end of the opening in the first wall and extends across the passageway and directs the flow of air through the opening in the first wall, and the second door is at the down-stream end of the opening in the second wall and deflects the flow of air through that opening.

Abstract: An afterburner flameholder for a gas turbine engine having a central diffuser cone, a generally cylindrical outer shell and fuel spray bars between the shell and cone defining an afterburner region is provided, which comprise an annular member with an asymmetric V-shaped cross section. The annular member is adapted to be secured to the engine in the afterburner region with the apex of the V facing upstream in the axial direction toward the fuel spray bars. The annular member has a first and second circular sidewall member each joined at one end forming an apex, the annular member in cross section has unequal length sidewall members. The difference in length between the distal ends of the first and second circular sidewall members is sufficient to cause spanwise vortices shed at the distal ends of the sidewall members to be out of phase when they meet downstream from the flameholder.

Abstract: A reheat system for a gas turbine engine includes a plurality of radial flameholders constructed of non-metallic material, such as a ceramic composite or carbon/carbon reinforced material, mounted on a metallic support structure in the engine jet pipe. The metallic support structure may be cooled by air from the engine bypass duct, and the non-metallic flameholders are capable of withstanding higher exhaust gas temperatures in the jet exhaust beyond the capability of currently used metal alloy materials. However, the mounting assembly must be capable of absorbing the considerable differential thermal expansion which will take place.

Abstract: A scramjet engine is disclosed which is effective for use in a hypersonic aircraft as an aircraft-integrated scamjet engine. The engine includes a first surface having an aft facing step, and a cowl upper surface spaced from the first surface to define an integrated inlet-combustor therebetween. Means for injecting fuel into the inlet-combustor at the step are provided for mixing fuel with supersonic airflow for generating supersonic combustion gases in the inlet-combustor. In the preferred embodiment of the invention, the fuel injecting means is effective for injecting fuel to create a fluid boundary defining a subsonic fuel zone and a supersonic fluid zone. The fluid boundary is variable and eliminates start and unstart problems requiring variable inlet geometry in a conventional scramjet engine.

Abstract: A shaft seal arrangement of a turbocharger for an internal combustion engine having a throttle valve disposed upstream of a compressor of the turbocharger. The shaft seal arrangement is comprised of first and second seal rings fitted on a cylindrical thrust spacer mounted on a rotatable shaft on which turbine and compresser wheels are mounted. The first and second rings are located spaced from each other and positioned between the compressor and bearings for the rotatable shaft. An annular air chamber is formed between the first and second seal rings and supplied with atmospheric air through an air supply passage.

Abstract: In an engine cooling system wherein supply pipes 40, 50 supply cooling fluid (e.g.: air) to the turbine 20 from a compressor bleed 10, upon failure of one of the supply pipes 50 the backflow of cooling fluid from the turbine to the point of failure may be stemmed by use of a vortex amplifier 140, 160 connected into each supply pipe 40, 50, the tangential ports of the vortex amplifiers being interconnected by a balance pipe 180. This allows the total minimum bleed flow rate of cooling fluid supplied to the turbine from the compressor bleed to be reduced.

Abstract: A rocket motor or other gas generator the case of which is ventable at a predetermined temperature substantially greater than the ambient temperature range for storage and operation thereof and substantially less than the auto-ignition temperature of gas generant material contained therein so as to prevent catastrophic propulsion or burst in case of fire. The venting is provided by memory metal members which are interleaved between plies of matrix-impregnated fibrous material in the case and which have a transformation temperature which is substantially higher than the ambient temperature range and substantially lower than the auto-ignition temperature and characterized by a change in shape when heated to a temperature above the transformation temperature.

Abstract: A noise reduction system for a jet engine including a tubular exhaust shroud affixed to the engine through which engine exhaust gases pass, porous acoustical lining affixed to a substantial portion of the internal circumferential surface of the shroud, a chamber in contact with the acoustical lining outer surface and a controllable source of air pressure affixed to the chamber to control the passage of air through the acoustical lining to thereby vary the acoustical absorption properties of the lining.

Abstract: This disclosure is for a liner for the exit nozzle of a gas turbine engine. The liner is double walled liner with insulation material between the walls, and with a plurality of tubes extending through the walls and the insulation. Primary cooling is accomplished by film cooling through the tubes. During periods of low coolant medium flow, the liner provides a barrier against radiation and convection. The use of the insulating material allows both walls of the liner to carry the pressure loads.

Abstract: A turbine ramjet engine is provided which has a ram pressure air duct shrouding the basic turbo-engine in an annular manner. When the ramjet operation is disconnected, this ram pressure air duct, while at the same time intake air is released into the basic engine, is blocked off with respect to the supply of ram pressure air. It is provided to unblock or block an intake-side air inflow cross-section of the basic engine by the axial adjustment of a drop-shaped body with respect to a flow divider. A ring slide is provided which is axially adjustable to always be opposite this body which blocks the ram pressure air duct when the inflow cross-section of the basic engine is unblocked and unblocks it when the inflow cross-section is blocked. In addition, the ram pressure air duct is to be widened in a multicornered manner in the upstream direction at the level of a blocking zone caused by a local overtravelling of countersurface of the mentioned body and the ring slide.

Abstract: A change-over arrangement for an integrated turboramjet engine has a guide ring which can be displaced in the axial direction and an annular cone which can be displaced against the guide ring. In the turbo-operation, the annular cone is in its front position and the guide ring is in its rear position closing off a ram air inlet duct, and in the ramjet operation, the annular cone, in its rear position, closes off the turbo-inlet duct, and the guide ring opens up the ram air inlet duct. As a result, a low-loss introduction of the air current can take place into the turbo-inlet duct or into the ram air inlet duct, without, at the same time, requiring disadvantageous enlargements of the diameter of the outer casing of the engine.

Abstract: A turbojet bypass engine includes an upstream fan of a conventional type also a downstream fan arranged to rotate in the opposite direction to the upstream fan and driven by a free turbine interleaved with the low pressure turbine which drives the upstream fan. The two fans are supplied by separpate overlapping and inverleaved air flow paths which are formed largely within the nacelle structure of the engine, giving the nacelle an ovoide shape which allows such engines with a high bypass ratio to be mounted below the wings of an aircraft.

Abstract: This disclosure describes an arrangement for cooling the exhaust nozzle of a gas turbine engine by selectively diverting the unheated total engine air flow directly to the exhaust nozzle of the engine. After the total air flow impinges on the back surface of the exhaust nozzle to rapidly cool the nozzle, the air is redirected back to the fan and the core engine for supporting combustion and to provide cooling in a generally conventional manner.

Abstract: A system for selectively venting cool bypass duct air into selected locations between a gas turbine engine and the engine shroud. The venting system is useful in a turbine engine having a core engine surrounded by an engine casing and nacelle, with a fan at the inlet directing air flow into the bypass duct between core engine and engine casing. The venting system basically comprises tubes inserted in holes in the engine shroud which separates the bypass duct from the engine compartment. Often sensors, electronic components or the like are advantageously located in the engine compartment. However, elevated temperatures in the engine compartment often damages or degrades those components. The vent tubes are positioned so as to direct cool air onto those components. A large number of vents may be provided in the shroud as manufactured, each plugged at manufacture, with only those opened which are needed for a specific engine with specifically located heat sensitive components.