Abstract: The present disclosure generally pertains to rocket based combined cycle (RBCC) propulsion units. In one exemplary embodiment, at least one rocket thruster is integrated with a jet engine but is external to the flow path of the jet engine, forming an altitude compensating plug nozzle. Since the rocket thruster is external to such flow path, the rocket flow from the rocket thruster interacts with the jet flow from the jet engine aft of the nozzle of the jet engine. Such interaction occurs without a significant performance penalty in the operation of the jet engine. In fact, it is possible that the interaction of the rocket flow with the jet flow may actually improve the efficiency of the jet engine under some conditions. Moreover, having the rocket thrusters positioned external to the flow path of the jet engine helps to avoid many of the problems plaguing conventional RBCC propulsion units.

Abstract: There is disclosed a vehicle and methods for maneuvering the vehicle. The vehicle may include a plurality of multiple-impulse rocket motors, each of which comprises a plurality of independently ignitable solid fuel propellant charges, and a processor that generates at least one command to ignite at least one solid fuel propellant charge of at least one of the plurality of multiple-impulse rocket motors.

Abstract: A core engine of a variable cycle gas turbine engine includes a low pressure spool for generating streams of bypass air and pressurized air, and a high pressure spool for further pressurizing the stream of pressurized air to generate streams of combustion air and supercharged auxiliary air. A peripheral case surrounds the engine case to form a peripheral duct. An auxiliary combustor and propulsor are positioned within the peripheral duct. A bleed duct extends from the high pressure spool to the auxiliary combustor. Variable ductwork directs airflow through the bleed duct and peripheral duct in two modes. A first mode comprises directing the stream of auxiliary air to the auxiliary combustor, and directing stream of inlet air through the peripheral duct. A second mode comprises directing the stream of auxiliary air into the stream of bypass air, and preventing inlet air from entering the peripheral duct.

Abstract: A method and system for operating a gas turbine engine is provided. The system includes in serial flow arrangement, at least one compressor, a combustor, and at least one turbine. The method includes compressing atmospheric air in the gas turbine engine, channeling at least a portion of the compressed air from the turbine engine flowpath to an oxidizer supply system, and channeling an oxygen-depleted flow of the compressed air from the oxidizer supply system to the gas turbine engine flowpath.

Abstract: A gas turbine engine assembly includes at least one propelling gas turbine engine and an auxiliary engine used for generating power. The propelling gas turbine engine includes a fan assembly and a core engine downstream from said fan assembly. The core engine includes a compressor, a high pressure turbine, a low pressure turbine, and a booster turbine coupled together in serial-flow arrangement such that the booster turbine is rotatably coupled between the high and low pressure turbines. The auxiliary engine includes at least one turbine and an inlet. The inlet is upstream from the high pressure turbine and is in flow communication with the propelling gas turbine engine core engine, such that a portion of airflow entering the propelling engine is extracted for use by the auxiliary engine.

Abstract: Provided is a turbofan engine capable of effectively using the air flowing through the inner diameter side area of a fan that is disposed at the front end side. The turbofan engine in which the fan is disposed at the front end side is equipped with a first low-pressure compressor at the upstream side and on the inner diameter side of the fan. Accordingly, the first low-pressure compressor can be operated by effectively using the air flowing through the rotation center portion of the fan, so that the air can be efficiently used and the output of the engine can be increased.

Abstract: An engine for an aircraft flying in a cruise flight direction comprises at least one rotor having a plurality of rotor blades. The at least one rotor is driven about a rotor axis at a rotational frequency. The rotor axis is oriented in the cruise flight direction of the aircraft, and at least one geometric parameter of the rotor blades that influences the propulsion characteristics of the rotor is variable. The engine further comprises a controller which periodically varies the geometric parameter of the rotor blades at at least one frequency that is at least as high as the rotational frequency of the rotor.

Abstract: An engine combination for generating forces with a gas turbine engine generating force, and an internal combustion engine provided in the combination as an intermittent combustion engine generating force having an air intake, there being a plurality of air transfer ducts each extending from a different location in the gas turbine engine so as to be capable to provide air in each of those air transfer ducts at one end thereof at pressures differing from one another and connected at the other end of each to the air intake to transfer air thereto.

Abstract: The invention relates to an aircraft propeller (1) comprising a turbomachine (8) housed in a nacelle (10) and a cooler (45) capable of being traversed by a hot fluid, which is to be cooled by thermal exchange with cold air external to the cooler. The propeller (1) comprises an air stream (13) (13b) capable of directing pressurized air towards an air duct (20) realized between an outer wall (6) and an inner wall (60) of the nacelle (10). The cooler comprises volumetric cooling means (14) located in the air duct (20) and surface cooling means (15), unrelated to said volumetric cooling means, located at an outside wall (6, 101) of the aircraft.

Abstract: A power extraction system is disclosed having a forward fan stage configured to pressurize an airflow, an aft fan stage having a tip-fan configured to pressurize a first portion of a pressurized air flow from the forward fan stage wherein the aft fan stage is driven by a second portion of the pressurized airflow, and a power drive system adapted to supply power from the aft fan stage.

Abstract: An engine that operates and produces the entire required vehicle thrust below Mach 4 is useful for a Hypersonic combined cycle vehicle by saving vehicle and engine development costs. One such engine is a combined cycle engine having both a booster and a dual mode ramjet (DMRJ). The booster and the DMRJ are integrated to provide effective thrust from Mach 0 to in excess of Mach 4. As the booster accelerates the vehicle from Mach 0 to in excess of Mach 4, from Mach 0 to about Mach 2 incoming air delivered to the DMRJ is accelerated by primary ejector thrusters that may receive oxidizer from either on-board oxidizer tanks or from turbine compressor discharge air. As the TBCC further accelerates the vehicle from about Mach 0 to in excess of Mach 4 exhaust from the turbine and exhaust from the DMRJ are combined in a common nozzle disposed downstream of a combustor portion of said DMRJ functioning as an aerodynamic choke.

Abstract: Units of engines are described herein. The units contain a unique combined-cycle (deflagration-detonation) “contra-rotation, anti-gyration, gyroscopic,” turbine fan-jet/free-piston engine configuration for induced air supercharging that boosts the performance of novel Ramjet engines or Ramjet engine configurations by improving internal air-stream dynamics. These dynamics are the result of co-operative air stream intermixing through convergent, supercharge-attenuated, inducted, compressed, tuned, pre-heated ambient air. Results are achieved through the varying of the geometric structural form and the utilization of unique engines and air induction and propulsion conformations, aided with supplemental air, fuel, oxygen and optional water and electrolyte charging.

Abstract: The invention relates to a nacelle (100) for the turbojet of an aircraft that comprises a downstream structure including: an outer structure (106), a concentric inner structure (109) surrounding a downstream portion (111) of the turbojet and including an upstream section having a relatively small diameter (113) and a downstream section having a relatively large diameter (114), wherein said inner structure (109) defines together with the outer structure (106) an annular flow channel (108); and a guiding system (140) for connecting the inner structure (109) and the downstream portion (111) of the turbojet engine or a portion of a suspension mast; characterised in that the guiding system (140) includes a means for combining a translation and rotation movement of at least a portion of said inner structure (109) between a working position, in which the inner structure (109) is used as a cowling for the turbojet downstream portion (111), and a maintenance position in which the inner structure (109) exposes said tur

Abstract: A gas turbine engine comprising an intermediate compressor and two shafts connecting respective high and low pressure turbines and compressors respectively is characterised by the intermediate pressure compressor connecting to either the high pressure shaft or the low pressure shaft via a gearbox.

Abstract: A gas turbine engine with a turbine with an exhaust manifold thereabout from which fluids can be transferred to the turbine and an air compressor having an air transfer duct extending therefrom so as to be capable to provide compressed air in that air transfer duct at one end thereof to the air intake of an internal combustion engine provided as an intermittent combustion engine. The air intake and an exhaust outlet are each coupled to combustion chambers therein and a rotatable output shaft is also coupled to those combustion chambers for generating force. The exhaust outlet has an exhaust transfer duct extending therefrom so as to have the intermittent combustion engine be capable to provide exhaust therefrom in that exhaust transfer duct at one end thereof, and the exhaust transfer duct being connected at an opposite end thereof to the exhaust manifold to be capable of transferring intermittent combustion engine exhaust thereto.

Abstract: Disclosed are propulsion systems 30 providing reduced fuel burn, weight and cost. A single gas generator core 38 drives multiple bladed propulsion elements 36 with a power train 48. The core 38 has a forward compressor 56 and a rearward turbine 40 and rotates about a longitudinal core axis 62. The bladed propulsion elements 36 rotate about bladed propulsion element axes 78 that are not coaxial with the core axis 62. The bladed propulsion elements 36 discharge an ambient air stream 50 rearward as a bypass stream 52 portion and a core stream 54 portion. The core stream 54 portion is directed to the compressor 56. The propulsion systems 30 mount inside a fuselage 22 of an airframe 20 or they are suspended beneath a wing 24 via pylons 32.

Abstract: An engine assembly includes a gas-turbine engine having a tailcone portion and a bypass duct, a rocket engine combustion assembly located at the tailcone portion of the gas-turbine engine, and a movable nozzle segment subassembly that is selectively engageable with the gas-turbine engine bypass duct in an open position and with the rocket engine combustion assembly in a closed position.

Abstract: One aspect relates to a hybrid propulsive technique, comprising providing at least some first thrust associated with a flow of a working fluid through at least a portion of an at least one jet engine. The hybrid propulsive technique includes extracting energy at least partially in the form of electrical power from the working fluid, and converting at least a portion of the electrical power to torque. The hybrid propulsive technique further includes rotating an at least one substantially axial-flow independently rotatable compressor rotor at least partially responsive to the converting the at least a portion of the electrical power to torque.

Abstract: One aspect relates to a hybrid propulsive technique, comprising providing a flow of a working fluid through at least a portion of an at least one jet engine. The at least one jet engine includes an at least one turbine section, wherein the at least one turbine section includes at least one turbine stage. The at least one turbine stage includes an at least one turbine rotor and an at least one independently rotatable turbine stator. The hybrid propulsive technique further involves extracting energy at least partially in the form of electrical power from the working fluid, and converting at least a portion of the electrical power to torque. The hybrid propulsive technique further comprises rotating an at least one at least one independently rotatable turbine stator at least partially responsive to the converting the at least a portion of the electrical power to torque.

Abstract: A power generation system for propelling, and generating electrical power in, an aircraft, having a gas turbine engine in an engine compartment in the aircraft with an air inlet in the aircraft that is curved along its extent in leading to an air compressor in the gas turbine engine having a compressor air transfer duct extending therefrom to an internal combustion engine provided as an intermittent combustion engine at an air intake thereof. An inlet duct manifold is positioned against the duct wall of the inlet duct to cover a perforated portion thereof on the air compressor side of a curve therein with the inlet duct manifold having an inlet air transfer duct extending therefrom that is coupled to the intermittent combustion engine air intake.

Abstract: An engine combination for generating forces with a gas turbine engine generating force, and an internal combustion engine provided in the combination as an intermittent combustion engine generating force having an air intake, there being an air transfer duct connected from a compressor in the gas turbine engine to the air intake to transfer compressed air thereto.

Abstract: A Brayton-cycle rotary ramjet engine (10) operated within the confines of a helically elongated pass-through duct formed between a preferably stationary radially outward surface (14) and an outer rotating flow channel (36). The flow channel (36) is contoured between its inlet (34) and outlet (38) to include a supersonic diffuser (40), a combustor (42) and an expansion nozzle (44). Gaseous fuel, or liquid fuel atomized by a fuel slinger (58) within a housing (46), or solid fuel in the form of fine particulates, is inter-mixed with an oxidizer prior to being directed to the flow channel inlets (34). The air and fuel are combusted in the flow channels (36) and exhausted through the rear of the housing (46). A generator (22) can be coupled to a power shaft (18) to convert net shaft power into electricity. Preferably, the rotor (24) and stator (12) are fabricated from a ceramic or other high-temperature material so that combustor exit temperatures (T3) can be operated at highly efficient levels.

Abstract: A method of transmitting power among a plurality of shafts in a turbine engine is disclosed herein as well as a turbine engine for practicing the method. The turbine engine includes a compressor section having a low pressure portion and a high pressure portion. The turbine engine also includes a turbine section spaced from the compressor section along a centerline axis. The turbine section includes a low pressure portion and a high pressure portion. The turbine engine also includes a low pressure shaft extending between the low pressure portion of the compressor section and the low pressure portion of the turbine section. The turbine engine also includes a high pressure shaft extending between the high pressure portion of the compressor section and the high pressure portion of the turbine section. The turbine engine also includes a tower shaft operably engaged with both of the low pressure shaft and the high pressure shaft.

Abstract: Hypersonic inlet systems and methods are disclosed. In one embodiment, an inlet for an airbreathing propulsion system includes an inboard surface at least partially shaped to conform to a plurality of streamline-traces of a design flowfield approaching an aperture, an outboard surface spaced apart from the inboard surface, an upper surface extending between the inboard and outboard surfaces, and a lower surface extending between the inboard and outboard surfaces, wherein leading edges of the inboard, outboard, upper, and lower surfaces cooperatively define the aperture.

Abstract: A power and cooling management system configured to flexibly couple various adaptive modules to an integrated power and cooling unit to suit any aircraft platform is provided. The integrated power and cooling unit has a compressor(s), power turbine(s), cooling turbine(s) and integral starter generator(s) mounted to the shaft of the power and cooling turbine. The integrated power and cooling unit may be pneumatically and/or pneumatically coupled to an adaptive module that comprises an additional compressor and an additional turbine or electrically coupled to a fuel cell which provides the main power after entering the full operation mode. When the engine includes an integral starter generator mounted thereto, the integral starter generator of the integrated power and cooling unit is operative to receive electric power from the engine mounted generator. Alternatively, a motor/generator may be mounted to the shaft of the additional turbine of the adaptive module.

Abstract: An engine assembly, an acoustical liner and an associated fabrication method are provided to address fan blade flutter and fan noise control simultaneously within the same liner area. Fan blade flutter is therefore controlled without necessarily increasing the weight of the engine, impairing the structural integrity of the engine, or increasing the noise generated by the engine. The acoustical liner may have additional acoustical degrees of freedom which permit these seemingly competing concerns to be addressed in a complementary manner. The acoustical liner may include inner and outer barrels with the inner barrel having a perforated face sheet, a perforated back skin and a core disposed between the perforated face sheet and the back skin. The fluid communication between the core and the space between the inner and outer barrels provides additional acoustical degrees of freedom which may be utilized to reduce fan blade flutter while concurrently limiting fan blade noise.

Abstract: The temperature rise of a wheel space between a high-pressure turbine and a low-pressure turbine is suppressed. Cooling air is led from outside a casing 17 to a wheel space via a low-pressure turbine initial stage stator blade 5 and a diaphragm 11. An upstream side space seal portion 41 is adapted to restrict and divide the upstream side space into an outer circumferential portion 25 and an inner circumferential portion 27 and to allow cooling air to the upstream side space outer circumferential portion 27 to blow out into the upstream side space outer circumferential portion 25. Also, a downstream side space seal portion 42 is adapted to restrict and divide the downstream side space into an outer circumferential portion 26 and an inner circumferential portion 28 and to allow cooling air to blow out into the downstream side space outer circumferential portion 26.

Abstract: A gas turbine engine includes a high towershaft driven by a high pressure spool and a low towershaft driven by a low pressure spool.

Abstract: To provide an aircraft propulsion system which can secure the optimum thrust and thrust vector for flight conditions, as well as the optimum sectional area for the engine, and which is highly compatible with the environment. An electrical generator is coupled to a turbofan engine, the electrical generator is driven by output power of the turbofan engine to output electric power, and an electromagnetic driving fan is driven by the electric power. On the other hand, after bringing each of coils in the electromagnetic driving fan to a superconductive state, liquid hydrogen is introduced to a heat exchanger, collects the energy of exhaust as heat, is then vaporized, and thereafter supplied to a combustor and to a fuel cell. Further, the electromagnetic driving fan is changed in its rotational phase by a rotating mechanism portion, is made movable in a width direction of a wing and a wing chord direction by a slide mechanism portion, and can be stored inside or outside the wing by a storage mechanism portion.

Abstract: A gas turbine engine assembly includes at least one propelling gas turbine engine and an auxiliary engine used for generating power. The propelling gas turbine engine includes a fan assembly and a core engine downstream from said fan assembly. The core engine includes a compressor, a high pressure turbine, a low pressure turbine, and a booster turbine coupled together in serial-flow arrangement such that the booster turbine is rotatably coupled between the high and low pressure turbines. The auxiliary engine includes at least one turbine and an inlet. The inlet is upstream from the high pressure turbine and is in flow communication with the propelling gas turbine engine core engine, such that a portion of airflow entering the propelling engine is extracted for use by the auxiliary engine.

Abstract: The present invention provides a pulse detonation insert that induces flow obstructions within the pulse detonation chambers, wherein the flow obstructions are operable to induce turbulence within a primary fluid flow passing over the obstructions. This turbulence may take the form of vortices that enhance the mixing of the oxidizer and fuel within the primary flow. Additionally, supports couple to the pulse detonation chamber walls and flow obstructions to hold the flow obstructions in place within the pulse detonation chamber. The combustion/Detonation of the mixed oxidizing fuel results in an increased velocity of the primary flow exiting the pulse detonation chamber and reduced the amount of unburnt fuel within the exhaust.

Abstract: An engine that operates and produces the entire required vehicle thrust below Mach 4 is useful for a Hypersonic combined cycle vehicle by saving vehicle and engine development costs. One such engine is a combined cycle engine having both a booster and a dual mode ramjet (DMRJ). The booster and the DMRJ are integrated to provide effective thrust from Mach 0 to in excess of Mach 4. As the booster accelerates the vehicle from Mach 0 to in excess of Mach 4, from Mach 0 to about Mach 2 incoming air delivered to the DMRJ is accelerated by primary ejector thrusters that may receive oxidizer from either on-board oxidizer tanks or from turbine compressor discharge air. As the TBCC further accelerates the vehicle from about Mach 0 to in excess of Mach 4 exhaust from the turbine and exhaust from the DMRJ are combined in a common nozzle disposed downstream of a combustor portion of said DMRJ functioning as an aerodynamic choke.

Abstract: The invention relates to a twin-spool gas turbine engine comprising a high-pressure rotor (and a low-pressure rotor, which rotors are mounted in bearings supported by an intermediate casing, at least one accessory gear box, a drive means, driving radial and coaxial shafts for transmitting movement to the accessory gear box, the drive means comprising a high-pressure drive gear driving one of said radial shafts and connected to the high-pressure rotor, a low-pressure drive gear connected to the low-pressure rotor upstream of the high-pressure rotor and driving the other of said radial shafts, the low-pressure drive gear being supported in the intermediate casing by an axial-positioning bearing.

Abstract: Gas turbines with multiple gas flow paths are provided. In this regard, a representative gas turbine includes: a spool; a compressor; a turbine mechanically coupled to the spool; the compressor having a first set of blades and a second set of blades, the second set of blades being located downstream of the first set of blades, the first set of blades and the second set of blades being driven by the spool; and means for enabling the first set of blades to rotate at a lower rotational speed than the second set of blades.

Abstract: A turbofan flow delivery system includes a fan case housing a fan. Flow exit guide vanes are arranged downstream from the fan and extend radially inwardly from the fan case toward a bypass flow path. A supply passage includes an opening provided in the fan case between the fan and flow exit guide vanes configured to selectively provide pressurized air to a component using air from the bypass flow path. Swirling air from the bypass flow path enters the supply passage and is converted to a static pressure.

Abstract: A representative gas turbine includes: a high pressure spool; a high pressure compressor; a high pressure turbine mechanically coupled to the high pressure spool; a lower pressure spool; a lower pressure turbine mechanically coupled to the lower pressure spool; a fan having a first set of fan blades and a second set of fan blades, the second set of fan blades being located downstream of the first set of fan blades and being operative to rotate at a rotational speed corresponding to a rotational speed of the lower pressure spool; and a gear assembly mechanically coupled to the lower pressure spool and engaging the first set of fan blades such that rotation of the lower pressure spool rotates the first set of fan blades at a lower rotational speed than the rotational speed of the second set of fan blades.

Abstract: A method facilitates assembling a gas turbine engine assembly. The method comprises providing at least one propelling gas turbine engine that includes a core engine including at least one turbine, coupling an auxiliary engine to the propelling gas turbine engine such that during operation of the propelling gas turbine engine, such that at least a portion of the airflow entering the propelling gas turbine engine is extracted from the propelling gas and channeled to the auxiliary engine for generating power, and coupling a modulating valve in flow communication to the propelling gas turbine engine to control the flow of airflow from the propelling gas turbine engine to the auxiliary engine, wherein the modulating valve is selectively operable to control an extraction point of airflow from the propelling gas turbine engine.

Abstract: An ultra-compact aerovortical swirl-enhanced combustion (ASC) system features an aerovortical swirl generator for use in rocket thrusters utilizing hypergolic or non-hypergolic propellants. The ACS thruster can be sized for diameters ranging from about 0.5 to about 2.0 inches, and producing thrust levels of approximately 5 lbf to about 250 lbf. A plurality of helicoid flow channels in the swirl generator introduces swirl into a flow stream of a first propellant within ultra-compact sized rocket thrusters. The ASC system also includes injectors for introducing a second liquid propellant into the swirling flowfield to promote rapid and efficient atomization, mixing and vigorous combustion, which, results in major improvements in combustion and propulsion performance over current rocket thrusters, but in much shorter combustor systems.

Abstract: A turbine overspeed control system for use with a gas turbine engine of a gas turbine electrical powerplant. The overspeed control system preferably comprises a first and second overspeed control subsystem. When an overspeed condition of the gas turbine engine is detected, the first overspeed control subsystem acts to remove air from a combustion section of the engine, while the second overspeed control subsystem operates to alter the angle at which an incoming flow of air is supplied to a compressor turbine located therein. The result of operating the first and/or second overspeed control subsystem is a reduction in the speed of the gas turbine engine that is much more rapid than can be accomplished by simply shutting off a fuel supply thereto.

Abstract: A rotary combustion engine (12) with a self-cooling system comprises a heat exchanging interface for discharging excess heat from the rotary combustion engine (12), and a direct drive fan (44) integrated on the rotary engine output shaft (26) for providing a flow of forced air over the heat exchanging interface.

Abstract: A gas turbine with contrarotating HP and LP turbines comprises an LP turbine having a plurality of moving wheels alternating with nozzles, the moving wheels of the LP turbine rotating in a direction opposite to the direction of rotation of the moving wheel of the HP turbine, and an inter-turbine casing having inner and outer casing walls defining a stream passage between the HP and LP turbines together with arms extending across the passage between the inner and outer casing walls. The gas turbine does not have a nozzle or device for forming the function of deflecting the stream between the outlet from the HP turbine and the first moving wheel of the LP turbine. The HP turbine is advantageously designed to deliver a stream that gyrates in the inter-turbine casing.

Abstract: A system for transferring torque between a pair of independently, concurrently rotating shafts of a turbofan engine includes a magnetic gearbox. The magnetic gearbox has a first ring structure, a second ring structure and an intermediate ring structure. Each ring structure has an annular aperture therethrough and a plurality of permanent magnets embedded therein. The intermediate ring structure is disposed between the first and the second ring structures. Each ring structure is coaxially concentric with, and independently rotatable with respect to the remaining ring structures. The first and second ring structures are each coupled to separate ones of the rotating engine shafts, and the intermediate ring is operable to transfer torque between the pair of shafts. Preferably, the intermediate ring structure is coupled to a rotating machine. The rotating machine has a controller, and is operable for adjusting a ratio of torque transferred between the pair of shafts.

Abstract: A method for exhausting gas from an aircraft engine assembly is provided. The method includes coupling a first exhaust duct in fluid communication only to a first engine. The first exhaust duct includes a primary outlet and a secondary outlet. The method further includes coupling a second exhaust duct in fluid communication only to a second engine. The second exhaust duct includes a primary outlet and a secondary outlet. The method also includes aligning a portion of the first engine secondary outlet concentrically within a portion of the second engine secondary outlet.

Abstract: Methods and apparatus for generating thrust from a gas turbine engine using a pulse detonation system is provided. The gas turbine engine includes a fan assembly, and a bypass duct channeling air to a core engine and the pulse detonation system. The method includes injecting a charge of fuel into a pulse detonation tube that is coupled radially outward from the core engine, and controlling a supply of air from the fan into the pulse detonation tube through an opening in the inlet end of the pulse detonation tube using a rotary valve that is positioned radially outward from the core engine.

Abstract: A method facilitates assembling a gas turbine engine assembly. The method comprises providing at least one propelling gas turbine engine that includes a core engine including at least one turbine. The method also comprises coupling an auxiliary engine to the propelling gas turbine engine such that during operation of the propelling gas turbine engine, at least a portion of the airflow entering the propelling gas turbine engine is extracted from the propelling gas turbine engine upstream from the core engine turbine, and channeled to the auxiliary engine for generating power.

Abstract: Some but not all engines of an aircraft are provided with thrust reversers. A safety device disables, at least from the position corresponding to idling, the action of the throttles corresponding to the engines devoid of thrust reversers when the thrust reversers are controlled to deployment.

Abstract: A bypass turbofan engine comprises a first propulsion system and a second propulsion system. The first propulsion system comprises a first fan rotor, a core engine, a first low pressure turbine and a first fan shaft drivingly connecting the first turbine and the first fan rotor. The second propulsion system comprises a second fan shaft drivingly connecting to a second fan rotor and the first propulsion system and arranged so that the first and second shafts are not coaxial with one another.

Abstract: An engine arrangement (14) comprises a gas turbine engine unit (16) and first and second support members (18, 20) extend from a structural component (24) of the engine casing (22) to support the engine unit in a position spaced from the body or airframe of the aircraft.

Abstract: A novel combined engine for a single-stage spacecraft is provided that combines a air-breathing engine utilizing oxygen in the atmosphere as oxidizer and rocket engines for obtaining thrust outside the atmosphere and that does not require a portion whose shape is variable in accordance with the flight speed. Rocket engines 15 are provided on struts 12 that form air introduction channels 10 in the air intake section 4. The rocket jets 18 from the rocket engines 15 control the flow of the airflows 16 introduced into the combustion chamber 20 in accordance with the flight speed. When the spacecraft 1 is stationary or in subsonic flight, the rocket jets 18 promote air intake into the combustion chamber 20 by lowering of static pressure due to expansion (ejector effect). In the subsonic flight condition, it performs the role of air compression, mixing with incoming air, fuel injection and ignition and during supersonic/ultra-supersonic flight it performs the role of a variable diffuser.

Abstract: A method facilitates operating a gas turbine engine assembly including a propelling engine and an auxiliary power engine. The method comprises channeling a portion of airflow entering an inlet to the propelling gas turbine engine through a fan assembly and a bypass duct such that the air bypasses the core engine, and channeling a portion of airflow from the bypass duct towards the auxiliary power engine.