Abstract: A cooling fluid reuse system for a turbine engine for capturing spent cooling fluids from row one turbine vanes and directing those fluids into a mixing chamber to be mixed with liner cooling fluids and used in a combustor. The cooling fluid reuse system may be formed from a fluid channel extending from the a cooling fluid collection channel in the first row of turbine vanes, through a rotor cooling fluid collection chamber in the rotor assembly, through a cooling fluid manifold, and into the at least one mixing chamber.

Abstract: The invention relates to a turbine 1, in particular a gas turbine, which along a swivel-mounted rotationally symmetrical rotor 2 has a compressor 3, a combustion chamber 5 and a turbine section 7 formed of a plurality of turbine stages 15, in which each turbine stage 15 contains mutually interacting blades 17 and vanes 16 which a hot working fluid 13 can flow around, with a coolant provided by the compressor 3 for cooling the blades 16, 17, which can flow in a channel along the rotor 2 from the compressor 3 to the turbine section 7 and into which a liquid can be introduced for cooling. In order to specify a turbine in which less wear occurs and in which the lifetime of the components is increased, it is proposed that the channel extends outside the rotor 2 and that the liquid can be introduced at the end of the channel which faces the compressor 2.

Abstract: A method for testing a cooling system for use in a gas turbine engine control system is provided. The method includes connecting an inlet of the cooling system to a differential pressure sensor, connecting an outlet of the cooling system to the differential pressure sensor, and determining whether or not a difference in pressure exists between the inlet and outlet, wherein such a pressure difference is indicative of whether cooling fluid is flowing through the cooling system.

Abstract: A gas turbine system comprises a compressor that takes in suction air on the inlet side and compresses it to compressor end air that is available on the outlet side, a combustor in which a fuel is burned by using the compressor end air while resulting in the formation of hot gas, as well as a turbine in which the hot gas is expanded while providing work output. In a method for cooling this gas turbine system, compressed air is removed from the compressor, is fed as cooling air for cooling inside an internal cooling channel through thermally loaded components of the combustor and/or the turbine, is then compressed and added to the compressor end air. An improved cooling without disadvantage for the efficiency of the system is achieved in that, in the manner of a targeted leakage, a small part of the cooling air is fed for film cooling into the turbine stream through drilled film cooling openings provided on the components.

Abstract: A process is provided for the control of the amount of cooling air of a gas turbine set. Suitable means are provided in the cooling system to enable the amount of cooling air to be varied. The control of this means takes place in dependence on an operating parameter (X). This operating parameter is made dependent on the fuel type used.

Abstract: A method and device for decoupling combustor attenuation and pressure fluctuation from turbine attenuation and pressure fluctuation in a gas turbine engine. The engine has: a compressor; a combustor; and a turbine, that generate a flow of hot gas from the combustor to the turbine. An aerodynamic trip is disposed in at least one of; a combustor wall; and an inner shroud of the nozzle guide vane ring, and is adapted to emit jets of compressed air from cross flow ports into the flow of hot gas from the combustor. The air jets from the cross flow ports increase turbulence and equalize temperature distribution in addition to decoupling the attenuation and pressure fluctuations between the combustor and the turbine.

Abstract: A cooling system for a gas turbine engine includes a fuel deoxygenator for increasing the cooling capacity of the fuel. The fuel deoxygenator removes dissolved gases from the fuel to prevent the formation of insoluble deposits. The prevention of insoluble deposits increases the usable cooling capacity of the fuel. The increased cooling capacity of the deoxygenated fuel provides a greater heat sink for cooling air used to protect engine components. The improved cooling capacity of the cooling air provides for increased engine operating temperatures that improves overall engine efficiency.

Abstract: A system located with the housing of a gas turbine engine for increasing the pressure of working fluid from the compressor that will be utilized to cool a component. In one form the system includes a pump rotatable with a turbine component to increase the pressure of the working fluid.

Abstract: A replaceable section (25) of force-cooling tubing assembly (21, 22) for attachment to a transition piece (5) of a gas turbine is comprised of two ends (54, 70) fashioned for attachment by removable unions (52) to adjoining parts of a force-cooling tubing assembly. When assembled thereto to complete the assembly, the replaceable section 25 provides for fluid communication between a manifold (3) and the transition piece (5) for either the supply or return of cooling fluid. A transition piece (5) in combination with two such assemblies, one a supply assembly (21), the other a return assembly (22), comprises a field-installable transition piece assembly (10) that provides for rapid and easy installation.

Abstract: A system for optimizing cooling air supply pressure includes a high pressure fluid source which receives bleed air from the exit of a high pressure compressor and a nozzle and ejector assembly for supplying fluid to a point of use at a pressure sufficient to maintain a required cooling airflow and backflow margin at the point of use, and for reducing leakage of the cooling fluid between the high pressure source and the point of use.

Abstract: A gas turbine set, with a cooling air system through which at least one cooling air mass flow flows from a compressor to thermally highly loaded components of the gas turbine set. Pressure increasing ejectors are arranged in a cooling air duct of the cooling air system for increasing the pressure of flowing cooling air. The ejectors are operable with a working fluid. The working fluid mass flow is less than twenty percent of a driven cooling air mass flow.

Abstract: A rotary internal combustion engine with intermittent burning cycle, converting heat energy into the rotary motion, with a multi stage compression rotor (3) and centrifugally-axial air compression impellers (3.2), an air valve (5.1), which alternatively filling up the combustors (4.3) with the compressed air and closing them. A fuel pump (6) compresses fuel and through fuel distributor (6.1) supplies it to the combustors (4.3) to achieve combustion when sprinkled on the hot-red spirals of the ignition plugs (4.5). The gas generated in the combustors (4.3) passes through the cascade turbines (5.10; 5.3; 5.11), driving the power rotor (5) and through the coupling-gear (5.4) transferring rotary motion to a drive shaft (7). A hollow impeller (3.5), mounted onto the compression rotor (3), pumps air through a diffusor-shade (3.8) for cooling the compressor from the outside and the inside.

Abstract: A gas turbine engine having a turbine section cooling system is disclosed which is able to cool the turbine section of the gas turbine engine to an optimal operating temperature range, while at the same time not substantially degrading engine performance. The disclosure provides a number of different structures for doing so including the provision of turning foils, turning holes, and turning grooves within a secondary air flow cavity which turns or directs cooling and parasitic leakage air into the turbine section in a direction in substantial alignment with a gas flow path through the turbine section.

Abstract: A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a first drive shaft; a booster compressor positioned downstream of and in at least partial flow communication with the fan section including a plurality of stages, each stage including a stationary compressor blade row and a rotating compressor blade row connected to a drive shaft and interdigitated with the stationary compressor blade row; a core system positioned downstream of the compressor, where the core system further includes an intermediate compressor positioned downstream of and in flow communication with the booster compressor, the intermediate compressor being connected to a second drive shaft, and a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet; and, a low pressure turbine

Abstract: In an exhaust gas housing (1) of a thermal engine, a radially outer housing casing (9) and a radially inner housing casing (10) arranged on the hub side are connected to one another via at least one thermally insulated carrying rib (3) acted upon by a cooling medium. A carrying rib (3) has at least two passage ducts (7) and (8) for the cooling medium, at least one passage duct (7) possessing a cooling medium supply (6) and at least one passage duct (8) possessing a cooling medium outlet (12), and these passage ducts (7) and (8) being in communicating connection in the radially inner hub-side end region via a deflection duct (11). The cooling medium is led from an external pressure source (5) through the carrying rib (3) to the region of the deflecting duct (11) arranged on the hub-side casing (10) and from there through the carrying rib (3) back again into a collecting duct (15) which issues preferably into an annular duct (26) for cooling the exhaust gas housing flange (24).

Abstract: A gas turbine (1), in particular in a power plant, has at least one combustion chamber (2) and an enclosed inner liner (3) which surrounds the combustion chamber (2) and an enclosed outer liner (4), having a stator (5) which has at least one vane row (6) with a plurality of vanes (7), a rotor (8) which has at least one blade row (9) with a plurality of blades (10), an air cooling arrangement (31) which is designed for cooling parts of the gas turbine (1) with air (L), and a steam cooling arrangement (32) which is designed for cooling parts of the gas turbine (1) with steam (D).

Abstract: A gas turbine electric powerplant, preferably driven by an aeroderivative turbine engine of split shaft design. The gas turbine engine is coupled to a speed reducer, which is in turn coupled to an electric generator. An engine mount is provided that ensures that the gas turbine engine will remain in proper alignment with the speed reducer and generator, even during the thermal expansion or contraction thereof. Preferably, the components comprising the powerplant are mounted to a common, transportable base, so that the powerplant can be delivered to various locations. An overspeed control system is provided for ensuring that a runaway condition of the gas turbine engine does not occur should the gas turbine engine become disconnected from the speed reducer or generator. Sensors are used to monitor multiple operating conditions of the powerplant.

Abstract: A turbomachine including a high pressure compressor with a downstream cone, a diffuser extending downstream by an inner casing disposed radially outside the downstream compressor cone, a combustion chamber disposed radially outside the inner diffuser casing, and a high pressure turbine whose rotor is connected to the downstream cone of the compressor by a connection drum, the inner diffuser casing and the downstream compressor cone defining between them an under-chamber cavity that is situated downstream from a discharge labyrinth in which cooling air flows downstream, wherein the engine also includes a stator shroud that is installed under the inner diffuser casing downstream from the discharge labyrinth.

Abstract: A co-generation turbocharged turbine system that utilizes a gas turbine (10) connecting the exhaust to a turbocharger (24) to introduce pressurized air into the inlet of the turbine for improving power and efficiency. A second embodiment places a work load (22) on an extension of the same power shaft (30), thereby connecting both the turbocharger driven rotor and drive rotor together in the form of an extended power shaft (30?). The turbocharger is in fluid communication from the gas turbine exhaust and is connected to the gas turbine air intake from the drive rotor of the turbocharger. A third embodiment consists of the same elements as the second embodiment except the work load (34) is driven by a turbocharger double-extended power shaft (30?) which extends from a vapor generating sub-system that has been added to the invention.

Abstract: The aim of the invention is a device for coolant cooling in a gas turbine which, with a relatively simple construction and low plant complexity permits a particularly high degree of efficiency in using the heat produced on cooling the coolant from a gas turbine. Said aim is achieved, whereby a number of interconnected evaporator tubes for a flow medium, are arranged in a coolant channel, connected to a gas turbine, to form forced throughflow steam generator. Said device is preferably used in a gas and steam unit with a waste heat steam generator on the exhaust gas side of a gas turbine, the heating surfaces of which are connected into the water-steam circuit of a steam turbine. The evaporator tubes of the device are thus connected on the inlet side by means of supply line to the feed water train of the water-steam circuit of the steam turbine.

Abstract: An object is to improve the operational reliability of a gas turbine by suppressing thermal stress and thermal deformation acting on the rotor of the gas turbine. The gas turbine has a rotor shaft constructed by arranging, in an axial direction in turn, a plurality of discs each having a plurality of combustion gas-driven moving blades annularly arranged on the peripheral portion and spacers arranged between the discs, and is characterized in that gap portions are formed between a region, on the rotor shaft center portion side, of the above-mentioned discs facing the spacers and spacers adjacent thereto, contact surfaces are formed both of which contact on both a region, on the rotor peripheral side, of the above-mentioned discs facing the spacers and adjacent spacers thereto, and a third flow path leading fluid to the above-mentioned gap portions is provided.

Abstract: An object is to improve the operational reliability of a gas turbine by suppressing thermal stress and thermal deformation acting on the rotor of the gas turbine. The gas turbine has a rotor shaft constructed by arranging, in an axial direction in turn, a plurality of discs each having a plurality of combustion gas-driven moving blades annularly arranged on the peripheral portion and spacers arranged between the discs, and is characterized in that gap portions are formed between a region, on the rotor shaft center portion side, of the above-mentioned discs facing the spacers and spacers adjacent thereto, contact surfaces are formed both of which contact on both a region, on the rotor peripheral side, of the above-mentioned discs facing the spacers and adjacent spacers thereto, and a third flow path leading fluid to the above-mentioned gap portions is provided.

Abstract: Bleeding is operated by annularly providing plural stationary blades at an interior side of a vehicle; annularly providing plural moving blades around rotor disk adjacent to stationary blades; providing plural stage units comprising the stationary and moving blades; introducing bleed air into each stage unit from a compressor; supplying bleed air extracted from a final stage of the compressor into a first stage unit; and supplying bleed air extracted from compressed air which has not yet arrived at a final stage unit of the compressor.

Abstract: A gas turbine engine includes a compressor powered by a turbine. The turbine includes a nozzle having vanes extending between outer and inner bands. Each vane includes an internal cooling plenum and a bypass tube extending through the bands. First and second manifolds surround the outer band and are disposed in flow communication with the plenums and bypass tubes, respectively. A bleed circuit joins the compressor to the manifolds for providing pressurized air thereto. A control valve modulates airflow to the first manifold and in turn through the cooling plenums of the vanes.

Abstract: The invention relates to a method of supplying cooling air to the hot portions of a turbojet that comprises, from upstream to downstream: a compressor; a diffuser; a combustion chamber; a distributor; and a turbine driving said compressor, in which method a flow of air is bled from the flow of air delivered by the compressor, the bled-off flow is cooled in a heat exchanger situated radially outside the combustion chamber, and is then directed radially inwards via the stationary blades of the distributor and used for cooling the moving wheel of the turbine, the method wherein the flow of cooling air is bled from the zone of the end of the combustion chamber that surrounds the diffuser, and wherein the stationary blades of the distributor are also cooled by a second flow of air bled from the diffuser.

Abstract: A gas turbine including a compressor, combustor and turbine further includes (1) a cooling air system for supplying part of air compressed by a compressor to a high temperature section of the turbine, (2) a heat exchanger for exchanging heat of part of air compressed by the compressor, this heat exchanger being mounted on the cooling air system, and (3) a system for adjusting the air temperature downstream from the heat exchanger in conformity to the operation period of the turbine.

Abstract: Fuel supplied to a rotary fluid trap is centrifugally accelerated within a first cavity adjacent a first side of a rotor, and is then directed though a plurality of first passages extending through the rotor between and proximate to the blades, and shaped so as to at least partially conform to the shape of the blades. Second passages extend within the blades from the first passages and terminate within associated cavities proximate to the tips of the blades. Relatively cooler fuel in the first passages is thermosiphon exchanged for relatively hotter fuel in the second passages so as to cool the blades. The heated fuel flows into a second cavity adjacent to a second side of the rotor and is discharged from the rotating frame of reference directly into the combustion chamber through a second rotary fluid trap. A separate fuel distribution circuit is used for starting and warm-up.

Abstract: A gas turbine has a cooling air system supplying air for cooling a high temperature part of the gas turbine and a spray air system supplying air for spraying fuel into a combustor and is formed so that a part of high-pressure air compressed by a gas turbine compressor is used as air of the cooling air system and spray air system, wherein a heat exchanger and a boost compressor are arranged downstream of the outlet of compressed air of the gas turbine compressor, and the boost compressor is composed of a parallel connection of a compressor driven by the turbine shaft and ae compressor driven by a driven source other than the turbine shaft, and pressurized air from the boost compressor is used as air for the cooling air system and the spray air system.

Abstract: A device and method for recuperating a gas turbine engine comprises a compressor being configured to receive a coolant fluid stream, to compress the coolant fluid stream and to discharge the compressed coolant fluid stream to a turbine in fluid communication with the compressor. The compressed coolant fluid stream undergoing thermal exchange within the turbine, exit the turbine thereafter. A source of a working fluid stream is in fluid communication with the turbine. The working fluid stream is fluidly isolated from a portion of the coolant fluid stream and undergoing thermodynamic expansion through the turbine to extract energy therefrom. Where desired, the entire coolant fluid stream is fluidly isolated from the working fluid stream. At least a portion of the coolant fluid stream is channeled downstream of the turbine to supply a preheated process fluid stream to an adjacent system.

Abstract: A steam temperature Ts and a casing air temperature Ta are measured by thermometers. The measurement results are taken into measuring devices and converted into electric signals. The electric signal is A/D converted by the measuring device and then, is sent to a control apparatus where a difference between both the temperature is calculated by a subtracter of a processor provided in the control apparatus. When an absolute value ?T=|Ta?Ts of this difference is contained within 10° C. continuously ten times, a control signal is sent from a computing unit to a controller which is the control section, a pressure adjusting valve and the like are controlled and a cooling medium is switched to steam.

Abstract: Aspects of the invention relate to a turbine engine configuration and method for overcoming a turbine blade tip clearance problem that can arise when the turbine inlet temperature is maintained at a high level during part load operation of the turbine. Aspects of the invention relate to reducing rotor cooling air to a temperature below the design temperature level by using, for example, additional heat extraction devices or by reconfiguring or resizing existing heat exchanger devices. Upon exposure to the cooled air, the discs and blades of the turbine will shrink so as to provide a clearance between the blade tips and surrounding stationary support structure. The design rotor cooling air temperature can be from about 350 degrees Fahrenheit to about 480 degrees Fahrenheit. Aspects of the present invention can be used to decrease the rotor cooling air to about 150 degrees Fahrenheit at about 70 percent load.

Abstract: A combined cycle gas turbine system is improved to enhance a gas turbine efficiency and a combined efficiency by effecting steam-cooling of a combustor transition piece and a turbine blade. The combined cycle system has a gas turbine (8) having a generator (1), a compressor (2), a combustor (3), a blade cooling air cooler (4) and a turbine (6), a steam turbine (29) having a high pressure turbine (21), an intermediate pressure turbine (22) and a low pressure turbine (23), and a waste heat recovery boiler (9). Saturated water of a high pressure pump (27) is partially led into a demineralizer (118) and a water sprayer (116) for cooling steam to be supplied into a moving blade (52). The steam, after being used for cooling, is recovered into a reheater (20). Outlet steam of the high pressure turbine (21) is led into a stationary blade (53) for cooling thereof and the steam is then recovered into an inlet of the intermediate pressure turbine (22).

Abstract: A casing air temperature Ta and a steam temperature Ts are measured, and if an absolute value ?T of a difference between these two temperatures is within a predetermined temperature, the gas turbine is connected to the generator. After the connection is done, the load is gradually increased, and the coolant changeover signal is sent from a processor to a controller. The coolant is then changed to the steam, thereby completing the connection of the gas turbine with the generator and the changeover of the coolant.

Abstract: A turbine shroud cooling system used in a gas turbine engine for controlling tip clearance between a turbine shroud assembly and turbine rotor blades comprises a cooling air passage for selectively directing a cooling air flow between components to be cooled and a turbine shroud support assembly for controlling the tip clearance and then later re-directing the cooling air flow to cool a downstream turbine component.

Abstract: A system to feed cooling air to a gas turbine, wherein the cooling air is taken from a high-pressure source, inside the gas turgine, and is conveyed to radial accelerators (129, which give rise to the tangential acceleration of the air in the direction of the peripheral motion of the rotor surface. After it has been accelerated to the peripheral sepped of the rotor, the cooling air enters radial holes (13), and, whilst passing radially through the radial holes (13), undergoes a reduction in the quantity of tangential motion, Subsequently, the cooling air is released into the hollow rotor, with a correspondingly reduced outlet radius (14). A series of labyrinth seals combined with brush seals separate the chamber for combined with brush seals separate the chamber for feeding of air to the radial holes (13), from the low-pressure environment around the pad #2 (15) of the said gas turbine.

Abstract: A ventilation device for a high pressure turbine rotor in a turbomachine, the turbine comprising upstream and downstream turbine disks fitted with blades, the device comprising a cooling circuit being supplied by a cooling airflow D taken from the back of the combustion chamber. The circuit is such that the airflow passes through orifices formed in an upstream flange of the upstream disk, such that this airflow circulates in the axial direction towards the downstream side between an inner reaming of the upstream disk and a downstream flange of the downstream disk, the device also comprising a labyrinth inserted between the two disks, such that the airflow is divided into a first flow F1 and a second flow circulating on each side of labyrinth towards the blades.

Abstract: A system and method of cooling a steam turbine having internal moving components to a predetermined temperature by controlling a flow of nitrogen through the turbine, thus decreasing the downtime associated with maintaining the turbine. This provides a more efficient and cost effective method of operating a power plant.

Abstract: In a cooling-air cooler for a gas-turbine plant of a power plant, in which cooling-air cooler first means for spraying water into the cooling-air flow and second means for generating steam are arranged in a pressure vessel, through which the cooling air flows, between a cooling-air inlet and a cooling-air outlet in the cooling-air flow, simplified operation is achieved in that a water separator is provided on the cooling-air side in the direction of flow downstream of the first means.

Abstract: Aspects of the present invention relate to hollow elongated turbine engine components, such as transition ducts, having a cooling system. The component can have first and second ends, and can be defined by an outer peripheral wall. The cooling system extends longitudinally within the wall between the first and second ends. While configured to cool the entire component, the cooling system is particularly configured to reduce thermal gradients that can occur in regions proximate to the first and second ends of the component caused by higher loads in those regions. The cooling system can include at least one pair of cooling channels. Each channel pair includes a first channel and a second channel. The first and second channels can be arranged such that a coolant supplied to each channel can exchange heat with itself in at least the region substantially proximate to one of the ends.

Abstract: A turbine rotor includes a plurality of moving blades having cooling paths through which a refrigerant flows inside, a plurality of wheels having the moving blades mounted in the outer periphery thereof and at least one spacer member installed between the wheels. The spacer member has a plurality of flow paths through which a refrigerant flows after cooling the moving blades. The plurality of flow paths have flow paths interconnecting to the cooling paths in the moving blades on a first wheel adjacent the spacer member and interconnecting to a space formed on a side wall surface with which a second wheel adjacent the spacer member and the spacer member are in contact.

Abstract: Part of compressed air discharged from a compressor is cooled by a cooler, and merged with a working fluid for a turbine to cool rotor blades of the turbine. Further, air from the compressor is bled through introduction passages, and the bled air is cooled by coolers and introduced to stationary blades of the turbine. The stationary blades of the turbine are cooled with air cooled by the coolers. A sufficient cooling effect is obtained by a small amount of air, and the amount of compressed air bled through the introduction passages is decreased. Compression power is converted into turbine output effectively.

Abstract: A gas turbine having a rotor driven by a shaft with a cooling medium passageway therein has a heat shield pipe portion mounted inside the passageway which forms an air layer with a wall of the passageway to protect a bearing member metal supporting the shaft from the heat of a cooling medium and a seal shaft portion for forming a labyrinth seal to prevent leakage of the cooling medium are formed as one piece and constitute a heat shied seal pipe.

Abstract: A gas turbine plant comprises an air compressor, gas turbine including at least one high temperature section, a driven equipment, which are operatively connected in series, a gas turbine combustor arranged between the air compressor and the gas turbine, a fuel system disposed for supplying a fuel to the gas turbine combustor, and a heat exchange section for heating the fuel from the fuel by means of a high pressure air as a heating medium fed from the air compressor.

Abstract: The invention relates to a gas turbine (3) with reheat in the turbine section (11), whereby a fuel injection device (37) is provided for the injection of fuel (33) in to a combustion air flow (61) in the direction of flow before a cooling air flow (65) branching. By means of the early mixing of fuel (63) in the combustion air flow (61) and concomitant additional use of the cooling air flow (65) as combustion air, the specific output and the efficiency of the gas turbine (3) are improved.

Abstract: An adiabatic combustor for a gas turbine including: (a) a combustion chamber designed and configured to produce high-pressure combustion gases for the turbine, the combustion chamber having a primary combustion zone containing a substantially vitiated-air zone; (b) at least one primary air inlet providing air to the primary combustion zone, and (c) a fuel injector for injecting fuel, disposed such that the fuel is directly introduced to the vitiated-air zone, wherein the primary air inlet is positioned and directed, and the combustion chamber is designed and configured, so as to produce an internal recirculation that generates a toroidal vortex within the primary combustion zone, thereby producing therein the vitiated-air zone and maintaining therein a state of flameless oxidation.

Abstract: A gas turbine has a cooling air system supplying air for cooling a high temperature part of the gas turbine and a spray air system supplying air for spraying fuel into a combustor and is formed so that a part of high-pressure air compressed by a gas turbine compressor is used as air of the cooling air system and spray air system, wherein a heat exchanger and a boost compressor are arranged downstream of the outlet of compressed air of the gas turbine compressor, and the boost compressor is composed of a parallel connection of a compressor driven by the turbine shaft and ae compressor driven by a driven source other than the turbine shaft, and pressurized air from the boost compressor is used as air for the cooling air system and the spray air system.

Abstract: In a reheat gas turbine engine for power generation, fuel is burnt with compressed air from a compressor in a first or primary combustor, the combustion products are passed through a high pressure turbine, the exhaust of the high pressure turbine is then burnt together with further fuel in a reheat combustor to consume the excess air, and the exhaust of the second combustor is passed through a lower pressure turbine. Excess air is supplied to the first combustor, thereby enabling so-called “lean burn” combustion for production of low levels of pollutants in the exhaust of the engine. Some turbine components of the turbines, e.g., blades or vanes, are cooled by cooling air supplies tapped off from the compressor.

Abstract: A gas turbine engine rotor has a body and blades, which are attached to the rotor by a root portion, for receiving a hot fluid flow and a compressed fluid from a compressor. Each blade has a base portion, an external airfoil surface, a blade flow portion of the airfoil, and a passage system communicating with the compressor and with the blade flow portion to create a thermal insulating boundary between the heated fluid and the external airfoil surface. The rotor has a substantially planar platform member, extending substantially transversally to the body of the blade and having an upstream portion protruding in the direction toward the combustion zone. The platform member has at least a pair of openings on either side of the airfoil, the openings being positioned in series along the blade flow portion. One opening of the pair defines an inclined passage in the platform member, and the other opening is separated by a partition from the first opening.

Abstract: This invention relates to an arrangement for the cooling of the casing 1 of an aircraft gas turbine with a bypass duct 2 and at least one cooling air tube 3, this cooling air tube 3 having its inlet 4 arranged in the bypass duct 2 and entering at least one cooling air chamber 5 for the supply of cooling air to the casing 1, characterized in that a movable air deflector 6 is arranged in the bypass duct 2 upstream of the inlet 4 of the cooling air tube 3.

Abstract: A process is provided for the control of the amount of cooling air of a gas turbine set. Suitable means are provided in the cooling system to enable the amount of cooling air to be varied. The control of this means takes place in dependence on an operating parameter (X). This operating parameter is made dependent on the fuel type used.