Abstract: A new and improved gas turbine engine including a gas generator and a power turbine is disclosed. The power turbine includes a first rotor having a plurality of first turbine blade rows extending radially outwardly therefrom, and a second rotor having a plurality of second turbine blade rows extending radially inwardly therefrom. The power turbine is supported aft of the gas generator and is effective for receiving combustion gases therefrom and expanding the gases through the first and second turbine blade rows for extracting substantially all output power therefrom for driving the first and second rotors in counterrotating directions.

Abstract: A cooling air outlet duct for a heat exchanger of a gas turbine engine is provided with a rotatable wall portion. The rotatable wall portion is rotatable about a pivot positioned at its upstream end so that the total flow area between the downstream end of the rotatable wall portion and a first wall of the fluid outlet duct and the fluid flow rate may be varied to obtain relatively high exit velocities for the cooling air discharging from an exit nozzle in a fan casing to give good thrust recovery. This produces a lightweight and uncomplicated valve for the fluid outlet duct.

Abstract: A propulsion engine (12) which combusts propellant received from the storage tank (24) in which a portion (28) of the tank contains propellant in the liquid state and in which an ullage (26) in a remaining portion of the tank contains the propellant in a gaseous state including a first propellant circuit (34) coupling liquid propellant stored in the tank to an evaporator (30); a second propellant circuit (36), coupling the gaseous propellant from the evaporator to the propulsion engine combustor and to the ullage; at least one power generating device (14 and 16) disposed in the second propellant circuit between the evaporator and the ullage, for providing a power output from energy of the gaseous propellant flowing in the second propellant circuit controlled by at least one control valve controlling a flow of gaseous propellant to the at least one device under the control of at least one valve control signal; a bypass circuit (40) coupled in parallel with the at least one power generating device containing a

Abstract: A power augmented gas turbine starter system which has a starter including a casing rotatably mounting a turbine wheel therewithin. A low pressure air supply communicates with the casing. A combustor is associated with the low pressure air supply for receiving air therefrom and supplying a flow of hot gas to the starter for augmenting the power of the starter. The air supply may be a bleed air conduit from an auxiliary power unit or from a main propulsion engine, and the combustor may be removably mounted directly in-line within the air supply conduit. The invention contemplates the use of a single auxiliary power unit and a single combustor for supplying a flow of hot gas to two or more gas turbine starters associated with their respective main propulsion engines. A cross-bleed circuit may be provided so that cross-bleed from one main engine can be used to start another main engine by providing bleed air for the other engine's gas turbine starter.

Abstract: A gas turbine engine (10) is provided with a bypass conduit (56) for venting air (52) leaking past a rotating compressor seal (46). The leakage air (58) is returned to the working fluid flow path (64) upstream of this first turbine stage (32).

Abstract: A propulsion system (10) having a propulsion engine (18) which combusts propellant received from a storage tank (12) in which a portion (14) of the tank contains propellant in a liquid state and in which an ullage (16) in a remaining portion of the tank contains the propellant in a gaseous state is disclosed. A first propellant circuit couples liquid propellant stored in the portion of the tank storing the propellant in a liquid state to an evaporator (22) for gasifying the liquid propellant. A second propellant circuit couples the gaseous propellant from the evaporator to the propulsion engine for combustion by the engine and to the ullage. A turbine (24) is disposed in the second propellant circuit coupling the gaseous propellant from the evaporator to the propulsion engine and to the ullage between the evaporator and the ullage for providing a power takeoff (26) powered by energy of the gaseous propellant flowing in the second circuit.

Abstract: A thrust ring assembly for use in a thrust reverser system for a gas turbine engine. The ring assembly 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 nacelle. The thrust reverser basically comprises at least one opening through the engine casing and nacelle, a turning vane cascade in the opening, a cover movable by an actuator between positions covering and uncovering the opening and at least one blocker door movable between a position allowing flow through the bypass duct and a position diverting flow to the turning vane cascade. The actuator and cascade are severely aerodynamically loaded during thrust reversal. The improved thrust ring assembly carriers those loads and transfers them to the engine casing. The forward ends of the cascade and actuator are secured to a bulkhead ring which is secured to the base of a frustoconical ring.

Abstract: A thrust engine is disclosed in the form of a hybrid staged combustion-expander topping cycle engine. The engine comprises a mixed cycle, one cycle being an expander cycle (14) operating at low temperatures and the other being a staged combustion cycle (12) operating at a higher temperature. A portion of liquid hydrogen from a fuel pump (16) is passed in heat exchange relation with the engine nozzle (22) for cooling same and heating the hydrogen, and the heated hydrogen is passed to an oxidizer turbine (28) for driving a liquid oxygen pump (81). Another portion of the liquid hydrogen from the fuel pump is passed in heat exchange relation with the engine thrust chamber (21) and combustor (20), and a portion of the resulting heated hydrogen is introduced, together with a portion of the pressurized liquid oxygen from the liquid oxygen pump, into a preburner (20) for combustion therein, and passage of the combustion gasses to a fuel turbine (26) for driving the fuel pump.

Abstract: High temperature deterioration of an injector 34 for a gas turbine including a compressor and a turbine wheel 10, 14, 30 is avoided through the use of swirler blades 60 that form a forced vortex adjacent the ends 52, 58 of fuel and oxidant injection tubes 38, 40 to prevent separation of the mixed stream thereby eliminating backflows or eddies of burning fuel.

Abstract: A control system and method of controlling a gas turbine engine. The control system receives an error signal and processes the error signal to form a primary fuel control signal. The control system also receives at least one anticipatory demand signal and processes the signal to form an anticipatory fuel control signal. The control system adjusts the value of the anticipatory fuel control signal based on the value of the error signal to form an adjusted anticipatory signal and then the adjusted anticipatory fuel control signal and the primary fuel control signal are combined to form a fuel command signal.

Abstract: A high performance turbine engine includes refrigeration apparatus powered by the engine and effective upon a fractional air flow received from the engine compressor to provide chilled air for cooling of internal engine components. The refrigeration process facilitates rejection of unwanted heat to ambient air despite high stagnation temperatures at high flight speeds. The invention allows increased turbine inlet temperature and improved fuel efficiency or thrust per pound of engine weight with present-day metals and cooling technology being used for engine construction.

Abstract: The cooling and mounting of vanes 50 in a turbine nozzle 24 between shrouds 28, 29 is facilitated by impaling the vanes 50 with threaded fasteners 54, 70 and conducting relatively cool air from the compressor 15 through the grooves 72 between adjacent threads 70 of the threaded fastener 54.

Abstract: A turbine assembly includes an annular fairing and strut subassembly which provides mechanical strength to the turbine assembly and support for the rotor aft bearing. A readily demountable annular outlet guide vane assembly is mounted to the aft of the fairing and strut assembly for deswirling the exhaust gasses which exit the turbine assembly. The aft mounting of the outlet guide vanes allows for an axially shorter and lighter turbine engine design which removes greater inlet swirl than prior designs, with equal or less pressure loss.

Abstract: A hypersonic scramjet engine fuel injector and a hypersonic scramjet engine having such a fuel injector. The engine has a serially connected inlet, combustor, and exhaust nozzle. Multiple, horizontally-spaced-apart fuel injectors are positioned in and connected to the combustor's top portion. Each fuel injector has a hollow wedge shape housing. To improve fuel-air mixing for better combustion, the housing's end wall has multiple, discrete, convergent-divergent fuel outlet nozzles and its side walls have horizontally-extending exterior grooves. Fuel is used to cool the fuel injector housing with the housing's bottom and side walls having fuel-exit holes and the housing having an interior serpentine fuel passageway.

Abstract: The components of a fuel injection system for a turbine engine 10 and provided with a purging system are reduced by utilizing a construction including a main control valve 24, at least one start fuel line 46 connected to a start injector 50, a plurality of main fuel lines 51 connected to respective main injectors 52 and a two-way valve 44 connected to the main fuel lines. The system includes a purge fuel line 54 and a three-way valve 28 connected to allow the flow of fuel to the start fuel line 46 and to the two-way valve 44 or for connecting the start fuel line 46 and the valve 44 to the purge line 34.

Abstract: A discharge system for a hermetic compressor of the type that comprises a shell in whose interior a cylinder is housed defining a compression chamber in conjunction with a rolling eccentric piston driven by an electric motor, such cylinder having a discharge orifice in communication with a discharge valve to communicate the compression chamber with a discharge muffler chamber. The discharge orifice has a surrounding wall defining a valve seat on the internal face of the discharge muffler chamber. The discharge valve in the muffler chamber comprises a disc; disc guide means fastened in an inner wall of the discharge muffler chamber; and an elastic biasing element forcing the sealing disc against the valve seat and dimensioned to allow the sealing disc to be displaced off the when a certain pressure in the compression chamber is reached.

Abstract: A hypersonic scramjet engine fuel injector and a hypersonic scramjet engine having such a fuel injector. The engine has a serially connected inlet, combustor, and exhaust nozzle. Multiple, horizontally-spaced-apart fuel injectors are positioned in and connected to the combustor's top portion. Each fuel injector has a hollow wedge shape housing. To improve fuel-air mixing for better combustion, the housing's end wall has multiple, discrete, convergent-divergent fuel outlet nozzles and its side walls have horizontally-extending exterior grooves. Fuel is used to cool the fuel injector housing with the housing's bottom and side walls having fuel-exit holes and the housing having an interior serpentine fuel passageway.

Abstract: An axial flow compressor for a bypass type jet engine has a low pressure stage which includes a one-piece, annular hoop shroud mounted to the airfoil of each of the rotor blades in the low pressure stage to divide the airflow entering the compressor into an outer, bypass airflow path and an inner airflow path directed to the higher pressure stages of the compressor. The hoop shroud is formed of circumferentially wound, parallel fibers such as silicone carbide fibers embedded in a metal matrix material such as titanium, wherein the volume ratio of the hoop shroud is approximately one-third fibers and two-thirds metal matrix material.

Abstract: A high bypass ratio turbofan engine having a fan section, a booster compressor disposed aft of the fan section relative to the flow of combustion gases through the engine, and a core section disposed aft of the booster compressor. A low pressure counterrotating turbine, disposed aft of the core section, is used for driving the fan section and the booster compressor. The counterrotating turbine includes at least one set of rotating turbine blades and at least one set of oppositely rotating counterrotating turbine blades. A twin spool shaft is provided for coupling the turbine blades to the booster compressor and for coupling the counterrotating turbine blades to the fan section. A reduction gear is disposed in the drive shaft for coupling the turbine blades to the fan section and for reducing the rotational speed of the turbine output power to match the rotational speed of the fan section thereby splitting the usable work of the turbine blades between the fan section and the booster.

Abstract: A two-stage turbine includes a first stage comprising a first stator having blades configured for accelerating the drive fluid to supersonic velocities and directing the fluid to the impulse blades of the first rotor. The turbine's second stage includes a second stator having blades configured for directing the fluid from the first stage to the reaction blades of a second rotor at a sonic velocity. Efficient operation of the turbine over a broad range of drive fluid velocities is achieved by providing the stators and rotors with elliptical leading edges to form high angles of attack and providing extended stator tail sections to control fluid flow to the rotors.