Abstract: This invention is directed to a wavy air-foil structure effective to pre-mix liquid fuel and air prior to ingestion into a gas turbine engine. The structure is a surface or boundary generally in the form of a linear combination of sinusoidals, and over which the incoming air is caused to flow at very near sonic speed. This invention is based on a new phenomenon observed in a study of small-perturbation theory of gas dynamics.

Abstract: A cospander for extending the high efficiency range of a turbine engine is described. A cospander, in one aspect, is a turbomachinery device including at least one stage wherein stator vane and rotor blade rows are continuously variable to an extent that each stage can be operated in compression, expansion, or null modes depending on the blade and vane angles at any specific operating point. Such a device placed ahead or within a typical compressor has the effect of reducing, increasing, or maintaining the through flow of the primary working fluid. Such flow variation control can be used to achieve the desired overall turbomachinery thermal efficiency improvements over a wider operating range than is otherwise possible. In one specific embodiment, for example, the cospander includes a variable stator vane and a variable pitch rotor blade in series flow relationship.

Abstract: A modified turbine engine design in which control of the forward load on a thrust bearing is improved, while reducing or eliminating the need for the use of a balance piston is disclosed. The turbine engine has a longitudinal axis and comprises, a compressor for pressurizing gases in the engine. The impeller has an impeller rear face and a tip. The engine further comprises a shaft mounted to a thrust bearing for rotation about the axis. The impeller is fixed to the shaft for rotation therewith. A cavity fixed within the engine casing is defined at least by the impeller rear face and is in communication with the pressurized gases at the impeller tip. The engine further comprises a bleed in communication with both the cavity and the pressurized gases, so that the cavity is pressurized by the pressurized gases from the impeller tip and the bleed, thereby loading the impeller rear face to load the thrust bearing through the shaft.

Abstract: A simple cycle gas turbine engine which includes a variable flow booster, a small high pressure compressor, and a power turbine is described. The variable flow booster, which includes a low pressure compressor, has at least two flow paths. In one embodiment, variable bleed valves, or doors, are used to control air flow through respective booster flow paths. In one embodiment, a first flow path directs air flow around booster compressor to the core engine, and a second flow path directs air flow at least partially through the booster compressor.

Abstract: A turbocharged ram tornado engine coupled with a jet tornado turbine heat recovery system, a variable speed fluid transmission, and a muffler/water recovery assembly. A combustor provides a source of hot gases which are forced into a tornado form inside an annular chamber housing a first turbine wheel, water is injected in the gas stream, the swirling gases are then passed to a second turbine wheel, via a turbocharger connected to the first turbine exhaust which provides air to the combustor, and a muffler/water recovery system for condensing and recovering water injected into the engine.

Abstract: A system for pressurizing a liquid oxygen tank (20) in a pressure-fed hybrid rocket (100) provides pressure at a first end of the tank (20) so that liquid oxygen can exit the second end. A hybrid heater (50) creates the pressure by mixing gaseous oxygen with the solid fuel (51) in the hybrid heater (50) to create an extremely high temperature. Helium is then mixed with the combustion products via an inlet (60) and this hot, inert mixture enters the liquid oxygen tank (20) via a diffuser (70). A second hybrid heater (90) pressurizes the gaseous helium sphere (40) to reduce the pressure decay caused by the withdrawal of the helium. Two of the hybrid heaters (50 and 90) are fed gaseous oxygen from a sphere (30). A novel ignition system is used to ignite the hybrid heaters (50, 90, 250). One of the hybrid heaters (250) is used to ignite the hybrid motor (100) in the rocket (11).

Abstract: A gas turbine engine with first and second axial flow compressors and an intercooler therebetween is provided with a diffuser to diffuse the air flow leaving the downstream end of the first axial flow compressor at a first radial distance from the axis to the upstream end of the intercooler at a second radial distance from the axis. The diffuser comprises a first radially extending wall, a second radially extending wall and a plurality of diffuser vanes extending therebetween which define a plurality of generally radially extending diffusing passages. The vanes are wedge shaped. This arrangement provides diffusion in a short axial length.

Abstract: In a gas turbine including a compressor having a bore and a rotor comprised of multiple stages extending between a first stage at a forward end of the compressor and a last stage at an aft end of the compressor, each stage including a rotor disk having a peripheral rim and multiple blades secured to the peripheral rim, a combustion system comprising a plurality of combustors utilizing discharge air from the compressor for combustion, and multiple turbine stages driven by combustion gases from the combustion system, the improvement comprising means for supplying cooling air at least to a peripheral rim of the last stage of the compressor.

Abstract: A method of adjusting capacity of air flow to a pressurized combustor in a gas turbine plant having a first compressor driven by a gas turbine, which is in turn driven by gases generated during combustion of a fuel in the combustor is disclosed comprising providing a second compressor located downstream of the first compressor, connecting the second compressor downstream of the first compressor, supplying air to the combustor, in addition to that provided by the first compressor, from the second compressor into an air path downstream of the first compressor, the additional air from the second compressor compensating for an insufficient air flow from the first compressor to achieve optimal air flow at full load operation of the gas turbine plant.

Abstract: A gas turbine engine with first and second axial flow compressors and an intercooler therebetween is provided with a diffuser to diffuse the air flow leaving the downstream end of the first axial flow compressor at a first radial distance from the axis to the upstream end of the intercooler at a second radial distance from the axis. The diffuser comprises a first radially extending wall, a second radially extending wall and a plurality of diffuser vanes extending therebetween which define a plurality of generally radially extending diffusing passages. The vanes are wedge shaped. This arrangement provides diffusion in a short axial length.

Abstract: The power produced by a gas turbine system is augmented by a direct contact heat exchanger for contacting and cooling humid ambient air with cooler water for producing cooled ambient air and warmed water, and a precompressor device for compressing said cooled ambient air to produce pressurized air that is warmer than ambient air and has a lower relative humidity. An evaporative cooler, which is supplied with the warmed water, is provided for cooling said pressurized air to produce cooled pressurized air at about ambient air temperature and relative humidity, which is supplied to the main compressor.

Abstract: An electrical power generation plant has a gas turbine subsystem free of a compressor. A compressor subsystem remote from the gas turbine subsystem has an inlet receiving air and an outlet furnishing compressed air. A compressed air line interlinks the outlets and gas turbine subsystem. An aspect or the invention includes combusting sufficient fuel having sulphur content with sufficient limestone to establish a ratio of Ca/S greater than 3.4.

Abstract: A method and apparatus provide cooling air from a compressor to hot components in a gas turbine engine. Compressed air discharged from an exit of the compressor is stratified into a main flow channeled to a combustor, and a separate bleed flow channeled to a heat exchanger. The bleed flow has an above average pressure whereas the main flow has a below average pressure. The bleed flow is cooled in the heat exchanger and channeled to hot components of the engine for cooling thereof.

Abstract: A method and apparatus for maintaining or holding the rotor of a high pressure compressor of a very rigid turbojet during turbojet operation. The interior of the rotor of the high pressure compressor comprises upstream-inclined balancing disks (13) which are free of fly-weights. The centrifugal force applied to such balancing disks creates a load or stress on the disks having a radial component rending to compensate for the stresses towards the outside due to the blades, in front of which are positioned the balancing disks. The process for the production of such an exhaust cone is made easier, particularly with respect to the blank to be supplied for obtaining such a part.

Abstract: An inlet air heating system for a gas turbine comprising a main air inlet; ducting leading from the main inlet to an inlet of the gas turbine; a manifold located adjacent the inlet for feeding intermediate temperature air into the ducting; a conduit connected to the inlet air manifold, the conduit having an ambient air inlet and connected to a compressor bleed air conduit, downstream of the ambient air inlet; and one or more eductors in the conduit through which the bleed air flows to entrain the ambient air into the conduit through the ambient air inlet and to mix the bleed air and ambient air for introduction into the ducting through the manifold.

Abstract: The operational capacity and efficiency of a combustion turbine is increased by treating the turbine inlet air to increase its density. The inlet air treatment is achieved in a columnar tower with the inlet air being drawn in at the top of the tower. Treatment water is injected into the tower at a downward velocity greater than that of the inlet air to establish a vertically descending, two-phase co-current flow of inlet air and treatment water. The system utilizes the natural phenomenon of drag-induced pressure boost to the air achieved by the difference in the relative velocities of the inlet air and the treatment water in the co-current flow. By utilizing a cold treatment water, for example at 32.degree.-40.degree. F., the tower creates a direct contact heat transfer situation that cools the inlet air, thereby further increasing its density.

Abstract: A method and system for supplying compressed air to a process plant using a combustor-turbine unit directly coupled to a bull gear meshing with pinions on which are mounted gas compression stages and expansion stages. Some stages compress a stream of air supplied to the combustor-turbine unit for combustion and to the process plant. Other stages expand or compress other gas streams directed to the combustor-turbine unit or to external applications. Direct energy transfers and intercooling and aftercooling after compression stages enhance the efficiency of the system.

Abstract: An inert gas turbine engine contains a working fluid closed loop system leading from an expansion turbine which is in fluid flow relationship with an expansion chamber. Working fluid within the expansion chamber is heated by the combustion of compressed air and fuel within the combustion chamber. The heated working fluid within the expansion chamber rotates an expansion turbine which in turn rotates the compressor. Combusted gases do not contact the turbine blades. The rotation of the expansion turbine is controlled by the temperature and pressure of the working fluid within the expansion chamber. The temperature of the expansion turbine and working fluid within the expansion chamber may be controlled by the use of various cooling means. Additional turbines may be incorporated within the engine to be turned by the working fluid within the expansion chamber.

Abstract: A method and apparatus for eliminating vortex whistle noise in a radial-to-axial compressor intake uses a plurality of vortex-disturbing tabs mounted to the trailing edges of inlet guide vanes and extending into the flow path between adjacent vanes. The tabs are mounted to the guide vane so that the tab is always perpendicular to the guide vane, regardless of the vanes angular position.

Abstract: In a method for increasing the compressor-related pressure drop of the gas turbine (8) of a gas turbine group with a waste-heat steam generator (13) connected downstream, an injector (5), through the mixing nozzle of which the precompressed air (4) from the compressor (2) flows, acts directly upstream of the gas turbine (8). At least part of the steam (10) produced in the waste-heat steam generator (13) is admitted to the driving nozzle of the injector (5), as a result of which the compressor air (4) is further compressed without consumption of power from the yield of the gas turbine (8).

Abstract: In order to reduce noise transmission through the air inlet duct (16) of a turbine engine (10), the air inlet duct (16) provides a non-linear path defined by a bend (24) for air from a source, and the turbine engine (10) includes a sound attenuation assembly (26) in the form of a splitter dissipative silencer disposed near the bend (24) in the air inlet duct (16). More specifically, the turbine engine (10) is such that the splitter dissipative silencer (26) can be formed as a modular assembly (42) removably securable near the bend (24) in the air inlet duct (16) upstream of a radial compressor (12) for removal and/or replacement or availability as an optional component of the turbine engine (10).

Abstract: An annular rotator is mounted between inner and outer stators and has a plurality of combustion chambers therein, each having an inlet port to receive a pressurized explosive fuel mixture from a source at one point in the rotation of the rotator and an exhaust port to later register with an exhaust opening in the outer stator when the fuel mixture is ignited. The intake ports later register with a source of presurized air and the exhaust ports register with additional exhaust openings in the outer stator to purge and cool the combustion chambers and to subsequently pressurize the chambers.

Abstract: Undesirably low efficiencies in a gas turbine engine of the sort having a radial compressor may be increased by forming the compressor with radially inner, high pressure ratio blades (16) and radially outer, low pressure ratio blades (18). The low pressure ratio blades (18) may discharge to a receiver (38, 40) for relatively low pressure air which in turn may be utilized as bleed air (FIG. 3) or as thrust (FIG. 2) as desired or even multi-staged with the high pressure blades (FIG. 4). The high pressure ratio blades (16) discharge to a receiver (48) which is in fluid communication with a plenum (54) to provide combustion air to a combustor (56) which in turn produces gases of combustion to drive a turbine wheel (66). By utilizing high pressure air from the high pressure ratio blades (16) for this purpose, engine efficiency is maximized.

Abstract: Hydraulic air compressor and turbine apparatus includes a hydraulic air compressor having an air intake disposed above the water intake for entraining air in falling water, and the output of the hydraulic air compressor is used as a compressed air source for a turbine. The turbine apparatus includes a regenerator for heating the incoming compressed air output from the hydraulic air compressor. The hydraulic air compressor may include a pump for providing a flow of water for entraining and compressing air to provide a closed cycle system, and it may include an air pump or fan to increase the amount of air entrained in a given flow of water. Several different structures for negating a loss of water velocity are also disclosed, and several different turbine systems are disclosed.

Abstract: The high cost of the fuel injection system in a gas turbine engine can be reduced by utilizing a low cost jet or nozzle (70) to spray fuel into the inlet (24) of a compressor (18). The resulting air-fuel mixture flows as a semi-discrete sector to a combustor (58) to be swirled therein and combusted to drive a turbine wheel (50).

Abstract: A narrow efficient operating island for a centrifugal compressor may be increased in a construction including a compressor section 18 mounting blades 22 surrounded by an annular shroud 38 by placing bleed passages 88 in the shroud 38 and angling them in the direction of flow, both axially (FIG. 1) and radially (FIG. 3).

Abstract: The present invention relates to air bleed arrangements on a compressor of a gas turbine engine. A rotor of the compressor has a plurality of stages of rotor blades, and the rotor is formed from a plurality of axially adjacent rotor discs. Two axially adjacent rotor discs define a chamber therebetween for a radially inward flow of bleed air. The rotor is provided with a plurality of apertures in rotor disc which bleed air from the compressor and supply the bleed air to the chamber. The adjacent rotor discs have opposing radially extending surfaces, and one of the opposing surfaces has a plurality of circumferentially spaced radially extending vanes and radially extending grooves which direct the bleed air radially inwardly to minimize pressure losses in the bleed air flowing through the chamber. This arrangement is not complex, is relatively easy and inexpensive to manufacture, has reduced weight and allows compressor to be manufactured by welding the rotor discs together.

Abstract: A system for cooling a turbine assembly in a gas turbine engine. The system has a bleed means for bleeding a portion of a first core of gas exited from a compressor section of the engine to thereby form a second core of gas. The second core of gas is additionally pressurized by a second compressor means to increase the pressure of the second core of gas. The second core of gas is then delivered to the turbine section for cooling of the engine such that the second core of coolant gas has a higher pressure than the first core of gas exiting the combustor section.

Abstract: Difficulties in starting turbine engines 10 at high altitudes are avoided in a starting system which includes a combustor 34 having a gas outlet 32 adapted to be in communication with a turbine engine nozzle 28, a fuel inlet 36, and an oxidant inlet 38 so that fuel may be oxidized within the combustor 34 by an oxidant to provide hot gases at the outlet 32. The system includes a first pressure vessel 58 adapted to receive liquid oxygen and a second pressure vessel 60 adapted to receive liquid nitrogen. Pumps 62, 64, and conduits 56, 54, 68 establish a flow path between the vessels 58, 60 and the oxidant inlet 38 while a mixing valve 66 is disposed in the flow path to mix the oxygen and nitrogen flowing therein in a desired proportion.

Abstract: The installation comprises a compression unit and a power unit. The compression unit comprises at least one engine operation according to a Diesel cycle and supercharged by a turbine having an expansion turbine supplied with the exhaust gases of the engine and rotary compressor taking air from the atmosphere through a filter. The compressor delivers a primary supercharging flow to the Diesel cylinders via a cooler. The Diesel engine drives two compressors supplied with atmospheric air through the filter. One of the compressors provides a secondary flow to the power unit. The power unit comprises a combustion chamber which is supplied through the exchanger with part at least of the secondary airflow from the compression unit. The combustion chamber feeds a turbine producing on its shaft the whole of the output power. The Diesel type engine may be a conventional crank type engine or a free piston engine.

Abstract: A one-piece housing casting comprising a hub adapted to receive a turbine shaft therethrough and a radially extending flange on the outer periphery of the hub. The flange includes radially aligned annular surface portion and means for spacing the surface a predetermined distance from a radially aligned housing wall close to the outlet of the centrifugal compressor. The flange also includes an axially extended portion having an axially extended surface and means for spacing the surface away from an axially extending surface on a turbine housing wall to define a passage therebetween. Preferably, the means for spacing the walls apart from the turbine housing wall surfaces comprise diffuser vanes integrally formed on the radially and axially aligned surfaces of the flange of the diffuser member.

Abstract: An axial flow turbine having one or more turbine wheels functioning both as the compressor and the turbine sections. Fluid flow is redirected over the compressor section to effect multiple stages of compression.

Abstract: Apparatus is disclosed of transferring energy into a flow of air without use of rotating parts, in which an inlet flow of air is made to swirl into a first circular path moving longitudinally of a housing, heat is added to the inlet flow of air at the end of the first path to produce hot gases, and the hot gases are directed into a second swirling circular path coaxially of and within the first path and having an open interface therewith, whereby energy is transferred from the hot gases in the second path into the inlet flow of air in the first path. An embodiment of the apparatus is disclosed in the form of a fluid mass reaction engine having an annular housing (11) including an inlet (10) thereinto and an outlet (38) therefrom, a combustion chamber (25) in the housing spaced from the inlet, spaced coaxial openings into and out of the combustion chamber, and a means for heating air flowing through said combustion chamber.

Abstract: An apparatus for optimizing blade and sealing slots of a compressor of a gas turbine comprising a rotor including a plurality of axially spaced discs and rotor blades on the discs and a stator casing surrounding the blades and forming clearance slots therewith. The blades and stator casing form a path for flow of air in a main stream in which the air is compressed and discharged at a discharge end of the compressor. A seal at the last of the discs opposes outflow of compressed air thereat and forms sealing slots through which leakage air can flow from the main stream. The leakage air is conveyed from the compressor and a portion thereof is diverted as a secondary stream in a direction opposite the main stream of air flow in the compressor, and discharged towards the last rotor disc and at least the next adjacent rotor disc to flow thereon.

Abstract: A gas generator having a double entry low pressure centrifugal compressor, a single entry high pressure centrifugal compressor and a radial inflow turbine positioned coaxially and coupled for interdependent rotation by a single shaft assembly for the double entry compressor mounted to the generator frame using a plurality of rigid hollow cross-over ducts which also carry diffused compressed air to the high pressure compressor, where the single shaft assembly uses two shaft portions connected by an expansible coupling having a helical spline torque transmitting device to allow the double entry compressor rotor to "follow" compressor housing axial position changer and to maintain a close running clearance despite thermal expansion in the structure between the low pressure (LP) and high pressure (HP) compressor portions of the gas generator.

Abstract: A compact, highly efficient single spool gas turbine gas generator uses a double-entry centrifugal first stage compressor, a single-entry centrifugal second stage compressor, and a radial inflow turbine to achieve an overall pressure ratio of greater than 15:1. The first stage pressure ratio is more than twice the second stage pressure ratio, and the first stage entrance Mach numbers are greater than about 1.4. The specific speed of each the compressors ranges from 0.65 to 0.85, and of the turbine from 0.50 to 0.75.

Abstract: A compressor disc assembly with an integral centripetal accelerator for the exhaustion of air for use in a cooling device of a gas turbine, the bleeding of air at the surface of the compressor drum being effected by orifices situated in the platforms of the blades and in the median plane of the compressor disc. The assembly is formed from two half discs. The centripetal accelerator includes several sectors detachably mounted in the space defined between the two half discs of the compressor, each sector being formed from two half shells having on their internal face ribs in the form of centripetal blades by which the half shells are assembled together. The invention is utilized for cooling a gas turbine by bleeding air from the compressor.

Abstract: The bore of the compressor for a gas turbine engine is heated by selectively bleeding compressor air from downstream stages so that heating only occurs at discreet times during the engine operating envelope. The bled air is admitted into the bore at the mid-stage station of the compressor wherein the compressor disks are scrubbed so as to expand and close the gap between the outer air seal and tips of the compressor blades during cruise of the aircraft and prevented from heating the disks during the high powered operations of the engine.

Abstract: An integrated coal-fired gas turbine power plant includes a hydraulic compressor for providing pressurized air to a coal combustion system and cooling air to the gas turbine. The power plant may further include a regenerator between the compressor and the combustion system, and a gas cleanup segment between the combustion system and the gas turbine.

Abstract: A diametrically opposed pair of radially inwardly directed vortex-disturbing tabs are secured to a compressor intake assembly to eliminate vortex whistle therein.

Abstract: An integrated coal-fired gas turbine power plant includes a hydraulic compressor for providing pressurized air to a coal combustion system and cooling air to the gas turbine.

Abstract: A system for cooling a gas turbine by bleeding air from the compressor and channeling the air at the center of the engine towards the turbine in order to cool it. The air is bled from the plane of the disc bearing the blades of the compressor, said disc having radial ducts for the flow of the air. The invention is utilized in gas turbine engines to cool the turbine.

Abstract: A gas turbine engine which includes a combustion chamber adapted to be supplied with fuel by at least one atomizer nozzle. The gas turbine engine includes an air compressor for supplying compressed air to the combustion chamber, with the compressor including a secondary guide vane set having a plurality of ducts. At least one of the ducts of the secondary guide vane set is connected to the atomizer nozzle by means of a conduit. The at least one duct is equipped with a boundary layer bleed-off.

Abstract: An air blast nozzle system for a gas turbine engine including a fuel nozzle supplied with fuel, and blast air taken from the discharge side of the engine compressor. The blast air is boosted in pressure by a boost pump acting as a supercharger during selected modes of operation for improved fuel atomization.

Abstract: An air proportioner is provided for a liquid hydrocarbon fueled gas turbine of the type which is convertible to oil gas fuel and to coal gas fuel. The turbine includes a shell for enclosing the turbine, an air duct for venting air in said shell to a gasifier, and a fuel injector for injecting gasified fuel into the turbine. The air proportioner comprises a second air duct for venting air from the air duct for mixing with fuel from the gasifier. The air can be directly injected into the gas combustion basket along with the fuel from the injector or premixed with fuel from the gasifier prior to injection by the fuel injector.

Abstract: An auxiliary air system contained within the nacelle of a gas turbine engine having the functions of cooling engine components and measuring air inlet temperature. The system comprises an air ducting pipe, temperature sensors extending into the pipe, a structure to be cooled, such as an electronics module, preferably with a heat transfer chamber, an air flow inducer, and, if desired, means for directing air onto other engine components.

Abstract: A gas turbine engine is disclosed which has an integral steam power system consisting of heat absorbing boilers which convert an unpressurized liquid into an expanded and heated steam by utilizing heat normally lost through component cooling systems and the exhaust system. Upon completion of the steam power cycle, the steam is condensed back to a liquid state through a condensing system located within the compressor and other functional components of the gas turbine engine. A system of high pressure air and friction seals restrict steam or liquid condensate within designed flow bounds. The gas turbine engine disclosed is designed to give improved fuel efficiency and economy for aircraft and land use applications.

Abstract: An axial compressor is provided with a cooling air manifold surrounding a portion of the shroud, and means for bleeding air from the compressor to the manifold for selectively flowing it in a modulating manner axially along the outer side of the stator/shroud to cool and shrink it during steady-state operating conditions so as to obtain minimum shroud/rotor clearance conditions. Provision is also made to selectively divert the flow of cooling air from the manifold during transient periods of operation so as to alter the thermal growth or shrink rate of the stator/shroud and result in adequate clearance with the compressor rotor.

Abstract: A gas turbine engine comprises a compressor, a combustion system and a turbine, the compressor and the turbine being interconnected by a shaft. The combustion system has an initial diffusing section and duct means are provided to convey cooling air from at least part way down the diffusing section, to the turbine. At least one aperture is provided in the duct means and adapted to allow part of the cooling air to flow forwardly of the shaft on its outer surface to cool it.

Abstract: In a gas turbine engine, fuel and air is forced into the primary zone of the engine combustion chambers during the start-up sequence by means of compressed air which is ducted through the engine fuel burners and impinges on fuel issuing from the burners, forcing the fuel to flow through tubes into the combustion chamber primary zone.