Abstract: A method and device for reducing the amount of noise generated by a supersonic jet engine is provided. The method creates an envelope of air around the supersonic jet exhaust. The temperature and velocity of this envelope is controlled to eliminate or reduce the formation of noise making Mach waves.

Abstract: A rocket engine thrust nozzle assembly is disclosed wherein combustion products from a plurality of rocket engines having hexagonally-shaped exhaust areas flow to the inlet of a main nozzle that transitions the flow of the combustion products from a non-circular flow area to a circular flow area.

Abstract: A combined turbofan and ramjet aircraft engine includes a forward bypass duct which allows the engine to operate more efficiently during the turbofan mode of operation. By mounting a ramjet preburner in the forward duct and isolating this duct from the turbofan bypass air, a transition from turbofan operation to ramjet operation can take place at lower flight Mach numbers without incurring pressure losses or blockage in the turbofan bypass air.

Abstract: The rotor of a fan disposed in the annular noncombusted air flow duct of a urbojet propulsion engine of the bypass type is carried by a rotary body which includes the rotor assembly of a multi-stage turbine driven by noncombusted air supplied thereto by an axisymmetrical auxiliary duct which diverts noncombusted air from the primary duct of the engine just downstream of the compressor.

Abstract: A combination turbojet-ramjet engine for powering a hypersonic aircraft is disclosed wherein the intake duct for the ramjet portion of the engine is centrally located within the turbojet portion of the engine in order to improve the access from outside the engine to the turbojet engine portion, as well as to maximize the use of common components and to simplify the control systems for changing from turbojet operation to ramjet operation and vice versa. The ramjet intake duct extends generally parallel to the longitudinal axis of the hypersonic aircraft engine and is located radially inwardly of the compressor, the combustion chamber and the turbine wheel of the turbojet engine portion. Flow control devices are located on the turbojet engine air intake, as well as the ramjet intake duct to control the supply of air to the respective engine portions.

Abstract: The weight of a start cartridge in a cruise missile (10) of the type having a booster rocket (60) for initiating flight and a thrust producing gas turbine (12) for sustaining flight may be eliminated by placing the gas outlet (62) for the rocket engine (60) in ducting (64, 82) downstream of the gas turbine engine exhaust (20) so that operation of the rocket (60) to initiate flight of the missile (10), through eduction, draws sufficient air through the engine (12) to accelerate the same to starting speed.

Abstract: The invention comprises an expansion nozzle for a rocket engine destined to function successively in atmospheric and space conditions. The nozzle is comprised of a main wall defining a surface of revolution that splays outwardly and further comprises a complementary annular convergent divergent nozzle having an exhaust port and also having its axis coinciding with the principal axis of the main wall. The complementary annular nozzle is located outside the main wall and has a gas distribution torus cooperating with gas flow supply means to produce an annular flow of gas surrounding a main flow of gas exhausted at the down stream extremity of said main wall of the expansion nozzle. The annular flow of gas forms a source of static pressure P.sub.s that is substantially lower than atmospheric pressure P.sub.a. P.sub.s has a value of between 0.3 and 0.8 times P.sub.a.

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: Self-contained hybrid propulsion systems have long been recognized as a class of propulsion systems that combine a liquid propellant and a solid propellant into a single system. The propellants are stored separately and the liquid propellant is delivered to the motor casing that holds the solid propellant. The present invention contemplates gasifying the liquid propellant prior to introduction into the motor casing in order to enhance system performance. The solid propellant grain is ignited and partially burned generating heat to evaporate the remaining solid propellant grain at a controlled rate. The resulting mixture is then passed to a secondary combustion region where it is mixed with additional gasified liquid propellant to complete the combustion. An integrated turbopump assembly including a pump portion, a preburner portion and a turbine portion is provided to pressurize and gasify the liquid propellant.

Abstract: A turbocharged, compounded cycle ducted fan engine system includes a conventional internal combustion engine drivingly connected to a fan enclosed in a duct. The fan provides propulsive thrust by accelerating air through the duct and out an exhaust nozzle. A turbocharger is disposed in the duct and receives a portion of the air compressed by the fan. The turbocharger compressor further pressurizes the air and directs it to the internal combustion engine where it is burned and exits as exhaust gas to drive the turbine. A power turbine also driven by exhaust gas is also drivingly connected through the engine to the fan to provide additional power. The size and weight of the turbocharger are reduced since the compressor's work is partially achieved by the compression effect of the fan. The total propulsive thrust includes the fan generated thrust which bypasses the turbocharger and the thrust of exhaust gases exiting the turbine.

Abstract: The invention relates to a multistage launch vehicle and method of multistage launch vehicle operation in which at least one subsequent motor stage is ignited following lift-off in partially overlapping relationship with its immediately preceding motor stage prior to shut-down and staging of the preceding motor stage in order to effect positive thrust control over the launch vehicle throughout shut-down and staging of the preceding motor stage. Preferably, each of the motors comprising the subsequently fired motor stage is provided with a thrust profile in which maximum thrust is attained prior to staging of the preceding stage.

Abstract: The high cost of fabrication of a turbojet propulsion system 12 for a missile having an engine 18, mounted within a casing 14, including an inlet 20 to a compressor 22 and an exhaust gas outlet 24 is avoided in a structure having a propulsion system 12 with an inlet 26, 28; 37 located forward of the engine 18 for directing free stream air to the inlet 20 of the compressor 22, an outlet 30, 33; 38, 40 located aft of the engine 18 for directing exhaust gas into the free stream air, an engine starter 42 aft of the engine outlet 24, and a fuel storing bladder 46 forward of the engine 18.

Abstract: A propulsion system for an aircraft capable of hypersonic flight combines a turbo-jet with a ramjet which are so arranged that the ramjet or jets are located directly adjacent to the aircraft body while the turbo-jet or jets are located away from the aircraft body so that the ramjet or jets is placed between the turbo-jets and the body. This arrangement makes the turbo-jets much more accessible for maintenance and exchange work. Both types of engines have a common air intake which is provided with flaps capable of guiding the body boundary layer through the ramjet when the latter is not in operation. When the ramjet is in operation, the air intake flaps form the air intake in such a way that boundary layer turbulent air is mixed with undisturbed incoming air for the ramjet while the air inflow to the turbo-jet is closed off altogether.

Abstract: A thin membrane of high strength but ductile material is placed over a perforated structural support, thus allowing high pressure on the membrane side of the support to separate one chamber from another. Upon application of pressure through the perforated structural supporting structure, the membrane will collapse allowing communication between chambers. The arrangement has utility in rocket motors and gas generators where discrete impulse segments are desired to be available on command, and may be fabricated in concentric or tandem (series) volumes.

Abstract: A turbocharged, compounded cycle ducted fan engine system includes a conventional internal combustion engine drivingly connected to a fan enclosed in a duct. The fan provides propulsive thrust by accelerating air through the duct and out an exhaust nozzle. A turbocharger is disposed in the duct and out an exhaust nozzle. A turbocharger is disposed in the duct and receives a portion of the air compressed by the fan. The turbocharger compressor further pressurizes the air and directs it to the internal combustion engine where it is burned and exits as exhaust gas to drive the turbine. A power turbine also driven by exhaust gas is also drivingly connected through the engine to the fan to provide additional power. The size and weight of the turbocharger are reduced since the compressor's work is partially achieved by the compression effect of the fan. The total propulsive thrust includes the fan generated thrust which bypasses the turbocharger and the thrust of exhaust gases exiting the turbine.

Abstract: A gas turbine includes a change-over valve having a rotatable part and a static part. The valve is disposed between upstream and downstream compressors of the engine for selectively directing flows from the upstream compressor and an auxiliary inlet to the downstream compressor or an auxiliary outlet, the downstream static part of the valve being formed as a structural element of the engine casing from which extends forwardly a support member which carries bearings for supporting the forward rotating part of the valve and the casing of the compressor.

Abstract: The invention relates to fluid outlet ducts particularly cooling air outlet ducts of a gas turbine engine. A fluid outlet duct for a gas turbine engine is divided across its width into a plurality of second plenum chambers, each of the second plenum chambers has an associated exit nozzle located in the casing of the gas turbine engine for discharging cooling air. The cooling air from a heat exchanger is supplied through a first plenum chamber to each of the second plenum chambers. Each of the second plenum chambers has a valve to vary the flow of cooling into the second plenum chambers so that the total flow area of the fluid outlet duct may be varied to obtain relatively high exit velocities for the cooling air to give good thrust recovery.

Abstract: One major problem associated with the design of high-speed vertical take-off or landing aircraft is the requirement to have the front vectorable nozzles of the aircraft deployed in the airstream which passes over the fuselage of the aircraft when they are in use. The nozzles tend to act as air brakes and thus seriously effect the forward speed and flight characteristics of the aircraft. This invention attempts to solve this problem by providing a vectorable nozzle which is rotatable about one axis between a first position in which it is stowed inside a cavity within the aircraft fuselage when not required, and a second position in which it is deployed into the airstream when required.

Abstract: A valve for selectively changing the direction of flow of working fluid through a variable cycle engine which comprises a first and second compressor 14,16 spaced along a flow duct 26. The duct 26 having air intakes 42 leading to the second compressor 16 and outlets leading to nozzles 44. The valve comprising a sleeve 30 axially movable along the duct. The sleeve 30 having openings 31 in it in which are located doors 32. Links 36 are connected to each of the doors 32 so that as the sleeve 30 is moved axially the doors 32 are pulled open to open the air inlets 42 and the outlets and simultaneously obturate the flow duct 26. In a second position of the sleeve 30 the doors 32 and the sleeve 30 close off the air inlets and outlets and open the duct to allow the first compressor to supercharge the second compressor.

Abstract: The core engine of a mixed exhaust turbofan aeroengine is provided with a conical afterbody. The normal outlet guide vanes which remove residual swirl from the turbine exhaust gases are dispensed with and replaced by ducts within an afterbody having entries which receive the hot gases in the same manner and at the same angle as the passages between the guide vanes would have done. The hot ducts, whose aspect ratio is preferably transformed between their upstream and downstream ends, extend through the afterbody in substantially straight paths to avoid aerodynamic losses, and intersect its conical surface to form preferably slot-like nozzles from which the turbine exhaust gases issue as flattened jets. The exhaust jets partake of the original axial and tangential/swirl velocity components of the turbine gases, plus a radial component (consequential upon the orientation of the ducts) to achieve good penetration of the bypass stream.

Abstract: The exhaust mixer can be classified as being of the multi-lobed type, with troughs between the lobes. The lobes and troughs comprise confronting pairs of flow surfaces which are twisted between their upstream and downstream ends such that a vortex flow system is established downstream of the mixer to encourage mixing between the bypass airstream and the turbine exhaust stream. The downstream edges of the twisted flow surfaces are in the form of sinuous curves which destabilize the vortex system and cause the vortices to burst, hence increasing the mixing efficiency.

Abstract: A low aerodynamic drag structural link suitable for use within the housing of a turbofan jet engine; the link includes length adjustment capability, pivotal end mounting provision, maintained airstream orientation capability, low mass and jam nut length and orientation locking. Several variations in link construction including a single ball and socket arrangement, varying link cross-section along its longitudinal length and the use of fairing nose and tail inserts are disclosed.

Abstract: A turboprop engine having a variable pitch propeller driven through a speed reducing transmission. The engine has a compressor section with a low pressure component and a high pressure component. A flow control device is positioned between the low and high pressure components. During takeoff, air from the low pressure component is bypassed, and during cruise, the air from the low pressure component is passed directly through the high pressure component, thus obtaining improved matching of engine components for both takeoff and cruise.

Abstract: A high speed, high-altitude jet aircraft engine operable as a ram-jet and/or turbo-jet using a single combustion chamber in a main housing receiving compressed air and also a variable flow of ram air bypassing the compressor. Primary and secondary burners in different parts of the combustion chamber produce respectively a major high temperature portion of burnt gas exhausted direct through a main nozzle and a minor lower temperature portion ducted off to drive the turbine which is in a side housing, the minor portion being exhausted through a subordinate nozzle.

Abstract: The invention relates to a propulsion engine, particularly for supersonic aircraft. An internal propulsion system comprises an envelope displaying a tapered center portion, around which is coaxially mounted an external propulsion system which lacks a separate compressor. According to the invention, mechanisms are provided to divert toward the combustion chamber of the external propulsion system an adjustable fraction of the compressed air arriving from the last compressor stage of the internal propulsion system. A blower which is driven by the external turbine exhausts in parallel with the latter into an annular conduit which emerges between the internal nozzle and external nozzle and an exhaust nozzle, all of these nozzles being of variable geometry. The purpose of the invention is to improve variable cycle propulsion engines for second-generation supersonic aircraft.

Abstract: Each duct of a plurality of generally cylindrical exhaust ducts disposed in approximately uniform annular arrangement about an axis is a tube, or a lobe of a radially-branched passage, having a discharge end canted so that the radially outer portion of such duct end projects rearward farther than any other portion of the duct end. The canted duct ends direct the major portion of the exhaust of a jet engine inward toward the axis of the annular exhaust duct arrangement in a plurality of exhaust jets, effecting a rearwardly convergent relationship of the jets.

Abstract: A nozzle for mounting to a jet engine at the exhaust end of the existing exhaust nozzle. The nozzle comprises a jet having an inlet for the exhaust gas jet which is exhausted from the exhaust exit of the engine. The jet inlet has means to divide the exhaust jet into two or more separate exhaust jet streams. The external surface of the body has means which create an airflow to combine with the separate jet streams which issue from the outlet end of the body.

Abstract: A composite cycle rocket engine having an inner engine disposed to discharge directly into the nozzle of an outer engine, is described herein.

Abstract: Aircraft engine exhaust is mixed with air and fuel and recombusted. Air is drawn into the secondary combustion chamber from suction surfaces on wings. Exhaust of the secondary combustion chamber is blown over wing and fuselage surfaces.

Abstract: A gas turbine engine power plant comprises a main gas turbine engine terminating in a jet pipe in which a free turbine is mounted. The free turbine is adapted to drive an auxiliary compressor. Means are provided such that a portion of the gas turbine engine efflux may be directed through the auxiliary turbine such that the auxiliary compressor augments the efflux from the main engine. Alternatively the entire gas turbine engine efflux may pass directly through the jet pipe.

Abstract: A turbofan-ramjet engine is provided with optional duct burning in the turbofan mode. A core engine is provided with a first annular passage therearound for directing airflow around the inlet of the core engine. A second annular passage having a fan directs air flow into the inlet of the core engine and said first annular passage. A third annular passage extends around the first and second annular passage with the first and third annular passage each directing their flow into a common annular combustion chamber. Variable vanes are provided in the first and third annular passages to provide for blocking off the passage or acting as swirl vanes depending on mode of operation. Blocking vanes are also provided in the annular exhaust passage of the core engine. Coannular nozzles are provided with one located at the exhaust of the core engine and an outer one located aft of the common annular combustion chamber.

Abstract: A helicopter rotor is formed with spars and ribs. The spars form parts of the surfaces, and long slots are constructed in the spars to provide suction and blower slots. Air is withdrawn in slots nearest the leading edge, and engine exhaust is conducted to a suction/blowing device which in turn blows air through slots near the trailing edge. Helicopter engine exhaust is mixed with air and fuel and is recombusted. Air is drawn into a recombustion chamber of a suction/blowing device from suction surfaces on the helicopter rotor blades. This suctioned air is then re-routed to the slots near the trailing edge and blown over the upper and lower surfaces of the blade.

Abstract: An improved rocket nozzle system having a closure insulator located adjacent the combustion chamber of a rocket engine. The closure insulator has formed as an integral part thereof a plurality of nozzles. Each of the nozzles has incorporated therein a throat insert of pyrolized graphite cloth laminate and a consumable washer. The entire closure insulator assembly including nozzles is transfer molded as one piece into a steel housing. The housing is mounted on the combustion chamber thereby providing an effective nozzle system which is reliably operable under high temperature operation.

Abstract: A nozzle assembly for use with a jet engine for controlling the direction of the discharged fluid and hence the direction of thrust is disclosed. The nozzle housing and an internally mounted bifurcation structure are arranged to divide the exhaust fluid flowing into a circular entrance orifice into two separate fluid streams that flow out of rectangular discharge orifices on the rear face of the nozzle. Flow deflection doors, stored within a rearwardly opening cavity formed by the bifurcation structure between the two rectangular discharge orifices, are selectively extendable into each fluid stream to modify the direction of the discharged fluid. The two flow ducts formed by the bifurcation structure and nozzle housing are configured to discharge the two fluid streams toward one another such that the fluid streams converge into a single, thrust-producing fluid stream aft of the exit orifices when the flow deflection doors are not deployed.

Abstract: A method and apparatus applicable to jet engines for improving operating efficiency over broad ranges of flight conditions and for reducing engine noise output in take-off and landing by controlling the airflow entering and exiting the engines. A turbojet engine apparatus is described which operates efficiently at both subsonic and supersonic speeds and a method is described which enables a turbofan with an associated satellite turbojet or turbofan to operate more efficiently at both subsonic and supersonic speeds. In both cases, take-off and landing noise is reduced substantially. The apparatus consists essentially of arranging for two separate portions of an engine to act upon one airstream or, alternately, to operate on independent airstreams.

Abstract: Apparatus to reduce the degree of contamination in front of a solid propellant rocket motor following actuation of thrust termination devices. By erecting a debris catcher over the ports through which the thrust termination contamimates pass, a few large particles are permitted to flow outwardly from the catcher into the atmosphere immediately after thrust termination, while smaller particles such as might cause interference with radar observation of the rocket motor are trapped into close confinement with the casing of the spent rocket motor.

Abstract: An internal combustion engine comprises a housing defining an air inlet opening, a first set of rotatably mounted interengaging air intake and compression rotor wheels, and an air passageway connecting said air intake and compression rotor wheels to a second set of rotatably mounted interengaging compression and metering rotor wheels. The compression and metering rotor wheels deliver air to a combustion chamber defined by the housing. Fuel is injected into the combustion chamber and ignited. The combustion chamber communicates with a third set of rotatably mounted interengaging output rotor wheels which are driven by the exhaust gases and are connected to drive the aforementioned rotor wheels. The working area of the output rotor wheels is greater than the working area of the compression and metering rotor wheels, wherein the combustion gases always drive the output rotor wheels.