Abstract: A vane carrier temperature control system for use in a gas turbine engine includes a first cooling air source, a second cooling air source, and an air temperature control system. The first cooling air source supplies a first portion of vane carrier cooling air extracted from a compressor section of the engine to a first section of a vane carrier that supports a plurality of rows of vanes within a turbine section of the engine. The second cooling air source supplies a second portion of vane carrier cooling air extracted from the compressor section to a second section of the vane carrier spaced from the first section in an axial direction defined by a direction of hot working gas flow through the turbine section. The air temperature control system controls a temperature of at least one of the first and second portions of vane carrier cooling air.

Abstract: A compressor of a gas turbine engine may have a bypass that routes a compressed air flow from within the compressor and directs the compressed air flow to a combustor. The bypass may have an inlet positioned just ahead of a downstream stage of the compressor and an outlet positioned to route the compressed air flow from the bypass to a diffuser or directly to a combustor. A valve may be used within the bypass and may be located near the inlet, near the outlet, or both. The valve may have the form of an annular sleeve in some embodiments and may be actuated with an actuator. The various arrangements allow for a compressor having a variable compression ratio.

Abstract: An air conditioning system in an aircraft is provided. The system comprises at least one air cycle air-conditioning pack with a compressor comprising a bleed air outlet and with a compressed-air line connected to the bleed air connection, which compressed-air line is connectable to a starter turbine for a main engine. With the use of air from a compressor of an air conditioning pack it is not necessary to switch off the air conditioning pack; bleed air removal from the auxiliary power units becomes obsolete; and, since there is no unnecessary discharge of compressed air from a bleed air system, noise pollution on the ground is reduced.

Abstract: A system and method for controlling aircraft anti-ice bleed air flow in an anti-ice system. The system includes a modulating shut-off valve that is moveable between a closed position and a plurality of open positions. Bleed air pressure upstream of the modulating shut-off valve is measured when the modulating shut-off valve is in the closed position. When an anti-ice initiation signal is received in a processor, the processor determines, based on at least the measured bleed air pressure when the anti-ice initiation signal was received, an initial open position of the modulating shut-off valve. The processor also commands the modulating shut-off valve to move to the initial open position, determines when a predetermined event occurs, and commands the modulating shut-off valve to remain in the initial open position until the predetermined event occurs period.

Abstract: A cooling system is provided for an aero gas turbine engine. The system has a duct which diverts a portion of a bypass air stream of the engine. A heat exchanger located in the duct receives cooling air for cooling components of the engine. The cooling air is cooled in the heat exchanger by the diverted bypass air stream. After cooling the cooling air, the spent diverted air stream is routed to a tail cone located at the exit of the engine and ejected through a nozzle at the tail cone.

Abstract: A method for monitoring a servo-control loop (3) of an actuator system (2) for actuating variable-geometry components of a turbojet, said method comprising: an estimation step of estimating a plurality of monitoring parameters from operating data of the servo-control loop (2); an evaluation step of evaluating a plurality of indicators from the monitoring parameters; an evaluation step for evaluating at least one signature matrix, each signature matrix being representative of the values of at least some of the indicators; and a detection and location step of detecting and locating a degradation affecting the servo-control loop as a function of said at least one signature matrix.

Abstract: A gas turbine engine variable bleed apparatus includes a variable bleed valve door disposed in a bleed inlet in a transition duct, rotatable about two or more separate pivot points, operable to open and close an aft bleed slot extending outwardly from transition duct, and operable to open and close a forward bleed slot extending inwardly into transition duct. Door is operable to transition between a first position with aft bleed slot open and forward bleed slot closed to a second position with aft bleed slot closed and forward bleed slot open without fully closing door. Door is rotatable about an axis translatable between the two or more separate pivot points. Transition duct having a transition duct conical angle at least about 10 degrees greater than a booster conical angle of a booster outer shroud upstream of transition duct.

Abstract: A cooling fluid air injection system for use in a gas turbine engine includes at an external cooling fluid source, at least one rotor cooling pipe, which is used to inject cooling fluid from the source into a rotor chamber, a piping system that provides fluid communication between the source and the rotor cooling pipe(s), a blower system for conveying the cooling fluid through the piping system and the rotor cooling pipe(s) into the rotor chamber, and a valve system. The valve system is closed during full load engine operation to prevent cooling fluid from the source from passing through the piping system, and open during less than full load engine operation to allow cooling fluid from the source to pass through the piping system.

Abstract: A gas turbine engine includes a buffer system that can communicate buffer supply air to a portion of the gas turbine engine. The buffer system can include a first circuit and a second circuit. The first circuit selects between a first bleed air supply having a first pressure and a second bleed air supply having a second pressure that is greater than the first pressure to render a first buffer supply air having an intermediate pressure. The second circuit selects between a third bleed air supply and a fourth bleed air supply to communicate a second buffer supply air.

Abstract: An integrated inducer heat exchanger is provided. The integrated inducer heat exchanger includes multiple airfoil devices disposed in an annular array within an inner circular casing and an outer circular casing forming multiple passages for allowing a flow of fluid from a forward side to an aft side of the integrated inducer heat exchanger. The integrated inducer heat exchanger also includes multiple annular manifolds arranged about the outer circular casing configured for supplying a flow of coolant at low temperature from one or more coolant sources and returning the flow of coolant at high temperature to the one or more coolant sources via an external heat exchanger for dissipating heat and multiple transfer tubes connecting the multiple annular manifolds with the multiple airfoil devices for transferring the flow of coolant within the airfoil devices for exchanging heat between the coolant and the fluid passing through the multiple passages.

Abstract: A cooling system for a gas turbine engine (10) having a compressor (14), a combustor (16) and a turbine section arranged to receive combustion products from the combustor. The cooling system includes ducting (32,44) defining a flow path from the compressor to a component, such as a turbine blade (26) to be cooled within the turbine section. The ducting (32) bypasses the combustor (16). A heat exchanger (34) may be arranged in the flow path to extract heat from the flow between the compressor and the component. One or more valves (36;64;66) in the flow path are actuated under the control of a controller (48;60) at a predetermined frequency of actuation so as to pulse the flow between the heat exchanger and the component, typically to improve the aerodynamic efficiency of turbine blades in use.

Abstract: A gas turbine engine includes a buffer system that includes a first circuit and a second circuit. The first circuit can communicate a first buffer supply air to a first portion of the gas turbine engine and the second circuit can communicate a second buffer supply air to a second portion of the gas turbine engine.

Abstract: One embodiment of the present invention is a unique gas turbine engine. Another embodiment is a unique cooling system for a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for cooling one or more objects of cooling. Further embodiments, forms, features, aspects, benefits, and advantages of the present application shall become apparent from the description and figures provided herewith.

Abstract: In one aspect, the present disclosure is directed to a method for controlling a position of a bleed valve of a gas turbine engine. The onset of an off-load transient may be determined. Values representative of the turbine rotor inlet temperature and the exhaust outlet temperature may be determined. Also, the amount of time elapsed since the onset of the off-load transient may be determined. Three provisional bleed valve command positions may be determined based on value representative of the turbine rotor inlet temperature, the value representative of the exhaust outlet temperature, and the amount of time elapsed, respectively. The provisional bleed valve command position associated with the lowest relative value may be selected. Then, the bleed valve position may be adjusted to match the selected bleed valve command position.

Abstract: A disclosed bleed air system utilizes high pressure air from a high pressure compressor to drive the turbo compressor to increase a pressure of bleed air drawn from the low pressure compressor. Air drawn from the low pressure compressor is at a lower temperature and pressure than that encountered from the high pressure compressor. The turbo compressor increases the pressure of airflow and provides that airflow into the main bleed air passage to be communicated to systems utilizing the bleed air.

Abstract: An integrated turbomachine oxygen plant includes a turbomachine and an air separation unit. One or more compressor pathways flow compressed air from a compressor through one or more of a combustor and a turbine expander to cool the combustor and/or the turbine expander. An air separation unit is operably connected to the one or more compressor pathways and is configured to separate the compressed air into oxygen and oxygen-depleted air. A method of air separation in an integrated turbomachine oxygen plant includes compressing a flow of air in a compressor of a turbomachine. The compressed flow of air is flowed through one or more of a combustor and a turbine expander of the turbomachine to cool the combustor and/or the turbine expander. The compressed flow of air is directed to an air separation unit and is separated into oxygen and oxygen-depleted air.

Abstract: The disclosure includes a load ramp system for a heat recovery steam generator (HRSG) of a combined cycle power plant and a method of increasing the rate of the load ramp process of a gas turbine system of the combined cycle power plant. In one embodiment, the load ramp system for the HRSG of the combined cycle power plant includes a conduit in fluid communication with a supercharger positioned upstream of a compressor of a gas turbine system in the combined cycle power plant. The conduit of the load ramp system is configured to direct a portion of supercharged air for use in cooling the HRSG during a load ramp process of the gas turbine system in the combined cycle power plant.

Abstract: The invention concerns a turbojet for aircraft including engine located in nacelle, and thermal exchanger intended to cool a fluid participating in the engine propulsive system, characterized in that said thermal exchanger is located on engine external wall, an interstitial space within which air can circulate being arranged between the engine external wall and a lower wall of said thermal exchanger. The invention also concerns an aircraft provided with at least one such turbojet.

Abstract: A turbine has a first and second inner wall, an inner casing and a shield element. The first and second inner walls are mounted to the inner casing such that an inner volume for working fluid is separated from an outer volume for cooling fluid. The first and second inner walls and inner casing are arranged to form a cavity in outer volume. The shield element is arranged inside cavity to separate cavity in an inner and outer region in a radially outer position to the inner region. A gap is formed between the first and second inner walls for working fluid between the inner volume and inner region. The shield element is arranged in cavity to form a fluid inlet for injecting cooling fluid from outer region to inner region for generating a predefined recirculation of working fluid and cooling fluid inside inner region.

Abstract: The gas turbine engine has a bleed air aperture formed in the radially outer wall upstream from the combustor and a bearing cavity formed within the radially inner wall, at least two bearing seals enclosing at least one bearing in the bearing cavity and separating the bearing cavity from associated buffer air entry points, an oil supply system including oil paths leading to each of the bearings; a buffer air supply system including buffer air paths leading to each of the entry points.

Abstract: A method and system for an integrated ejector valve assembly is provided. The integrated ejector valve assembly includes a first valve assembly configured to control a flow of relatively lower pressure fluid from a first inlet port, a second valve assembly configured to control a flow of relatively higher pressure fluid from a second inlet port, a first actuation chamber configured to close the first valve assembly, a second actuation chamber configured to close the second valve assembly, and a third actuation chamber configured to open the second valve assembly.

Abstract: A combined cycle power plant includes a compressor section including a compressor inlet and a compressor outlet, and a turbine section operatively connected to the compressor section. The turbine section includes a turbine inlet and a turbine outlet. A heat recovery steam generator (HRSG) is fluidly connected to the turbine outlet. A combustor includes a head end and a combustor discharge. The head end is fluidly connected to the compressor outlet and the combustor discharge is fluidly connected to the turbine inlet. A carbon dioxide collection system is fluidly connected to one of the compressor outlet and the head end of the combustor. The carbon dioxide collection system is configured and disposed to extract a first fluid comprising carbon dioxide and a second fluid from a substantially oxygen free fluid flow passed from the one of the compressor outlet and the head end of the combustor.

Abstract: An operation method for a gas turbine is provided which includes compressing a working fluid by a compressor, feeding this compressed working fluid into at least one combustion chamber where it is overheated, expanding this overheated working fluid in at least one expansion turbine to produce energy, carrying out a first tapping of the compressed work fluid from the compressor to feed it into a first cavity of the turbine for cooling, carrying out a second tapping of the working fluid downstream from the first tapping to feed it into a second cavity of the turbine upstream from the first cavity, for cooling, and fluidly connecting the first tapping to the second tapping to selectively feed the first tapping by a part of the second tapping during partial load operation conditions to keep the temperature of the first cavity within the acceptable limits for the resistance of the materials.

Abstract: A method and apparatus is provided for cooling the external surface of an aircraft APU eductor assembly. A processor is configured to open a surge valve by a predetermined amount when the surge valve is closed and the temperature of the exhaust gas exceeds a predetermined temperature in order to cool the surge plenum surfaces.

Abstract: The present application provides an inlet bleed heat system for supplying a flow of bleed air to a flow of incoming air into a compressor of a gas turbine engine. The inlet bleed heat system may include an air knife and a silencer panel. The air knife may include a compressor bleed air port in communication with the flow of bleed air and a discharge gap to discharge the flow of bleed air into the flow of incoming air. The air knife may and the silencer panel may form an integrated air knife/silencer panel.

Abstract: An air injection system for use in a gas turbine engine includes at least one outlet port through which air is extracted from the engine only during less than full load operation, at least one rotor cooling pipe, which is used to inject the air extracted from the outlet port(s) into a rotor chamber, a piping system that provides fluid communication between the one outlet port(s) and the rotor cooling pipe(s), a blower system for extracting air from the engine through the outlet port(s) and for conveying the extracted air through the piping system and the rotor cooling pipe(s) into the rotor chamber, and a valve system. The valve system is closed during full load engine operation to prevent air from passing through the piping system, and open during less than full load engine operation to allow air to pass through the piping system.

Abstract: An end-wall component of the mainstream gas annulus of a gas turbine engine has a cooling arrangement including one or more circumferentially extending rows of ballistic cooling holes through which, in use, dilution cooling air is jetted into the mainstream gas to reduce the mainstream gas temperature adjacent the end-wall. A portion of the cooling holes are first cooling holes angled such that the direction of the dilution cooling air jetted therethrough has, on entry into the mainstream gas annulus, a component in one tangential direction. A portion of the cooling holes are second cooling holes angled such that the direction of the dilution cooling air jetted therethrough has, on entry into the mainstream gas annulus, a component in the opposite tangential direction. The first and second cooling holes are arranged such that the cooling air from jets having entry components in opposing tangential directions collide and coalesce.

Abstract: Embodiments of methods and apparatus for providing compressor extraction cooling are provided. According to one example embodiment, a method is disclosed for controlling compressor extraction cooling. The method can include providing a cooling medium. The method can include extracting air from a compressor associated with a gas turbine. The method can also include introducing the cooling medium to the compressor extraction air, wherein the compressor extraction air is cooled by the cooling medium prior to or during introduction to the turbine section. Furthermore, method can include selectively controlling at least one of the compressor extraction air or the cooling medium based at least in part on a characteristic associated with the gas turbine.

Abstract: Thermal control is provided for a gas turbine casing by supplying thermal control gas from a compressor to a space between an outer casing and an inner casing, and transferring the thermal control gas from the space through the opening in the inner casing via a plurality of holes defined through a plate attached to an outer surface of the inner casing. The holes are arranged with a predetermined non-uniform distribution corresponding to a desired preferential impingement pattern for providing non-uniform heat transfer. A gas turbine thermal control assembly includes structure providing preferential heat transfer from the inner casing during operation of the gas turbine via a thermal control gas flow path from radially outside of the inner casing into the interior of the gas turbine.

Abstract: A combustor liner which reduces cooling flow to a combustion chamber and augments pressure drop split between impingement holes and effusion holes is disclosed. The combustor liner may further include accelerating channels, trip strips, pedestals, and cone-shaped effusion holes to provide further cooling of the liner. The combustor liner may reduce NOx production and the temperature of the combustion chamber of a gas turbine engine or the like.

Abstract: An air-oil cooler (AOC) for a gas turbine engine is disclosed. The AOC may comprise an oil inlet, an oil outlet, and heat exchange elements between the oil inlet and the oil outlet. The AOC may be longitudinally positioned between a fan and a V-groove of the engine and radially spaced between a low pressure compressor and a low pressure compressor panel. A gas turbine engine comprising an AOC is disclosed. The AOC of the engine may comprise an oil inlet, an oil outlet, and heat exchange elements between the oil inlet and the oil outlet. The AOC of the engine may be longitudinally positioned between a fan and a V-groove of the engine and radially spaced between a low pressure compressor and a low pressure compressor panel. A method of operating an AOC for use on a gas turbine engine is also disclosed.

Abstract: A system includes a turbine combustor containing a first wall portion disposed about a combustion zone, a second wall portion disposed about the first wall portion, and a plurality of vanes disposed within an annulus between the first and second wall portions. The plurality of vanes is configured to swirl air flowing within the annulus.

Abstract: In a jet engine with compressor air circulation provided for stabilizing the flow conditions, the compressed hot compressor air tapped from the flow path of the compressor (1) is first cooled in the bypass duct (3) and then resupplied to the upstream compressor via line (12). This reduces the thermal load of the blades and improves efficiency.

Abstract: A combustor liner grommet is disclosed. The grommet may include a peripheral wall defining a hole in a combustor liner and further including at least one cooling air flow channel. The cooling air flow channel in the grommet wall may be a slot or a hole. The channel may increase cooling flow to the grommet and the combustor liner around the grommet to prevent cracking from heat stress.

Abstract: A turbine engine has a compressor for delivery of compressed air to a combustor. The combustor delivers hot combustion gas through an outlet to a turbine. The turbine includes a nozzle assembly, downstream turbine blades, and shroud assemblies adjacent radially distal ends of turbine rotor blades. The nozzle and shroud assemblies include internal cooling passages for receiving compressed air from the compressor and, cooling air apertures opening through walls of the vanes and shrouds into the hot gas path to release film cooling air. The number of apertures, the aperture area, and the aperture pattern are varied in relation to the circumferential temperature profile of the combustion gas with a higher aperture area and/or higher number of apertures in high temperature regions and a lower aperture area and/or lower number of apertures in low temperature regions.

Abstract: Systems and methods for controlling compressor extraction air flows from a compressor of a turbine system during engine turn down are provided. In one embodiment, a method for controlling compressor extraction air flows from a compressor of a turbine system during turn down includes a control unit monitoring one or more operating parameters of a turbine system associated with an exit temperature of a combustor of the turbine system. The method further includes the control unit detecting one or more operating parameters meeting or exceeding a threshold associated with a decrease in the exit temperature of the combustor. In response to detecting one or more operating parameter meeting or exceeding the threshold, a control signal is transmitted to at least one variable orifice located in the turbine system causing at least one variable orifice to alter at least one extraction air flow from the compressor.

Abstract: Discrete bleed behind stator vane platform is provided. The discrete bleed behind stator vane platform relates to a system for bleeding off a working fluid from an inner volume (Vi) of a turbo-machine. The system includes a vane carrier with an annular rail and a vane device comprising at least one vane element, a vane platform and a vane root. The vane element is mounted to the vane platform and the vane root is mounted to the annular rail. A first annular cavity is formed between the vane platform and the annular rail, wherein an annular gap is formed between an edge of the vane platform and the vane carrier such that a part of the working fluid of the turbo-machine is bleedable through the annular gap into the first annular cavity. A second annular cavity is formed between the vane root, the annular rail and the vane carrier, wherein at least one inlet hole is formed into the annular rail for coupling the first annular cavity and the second annular cavity.

Abstract: A combined cycle power plant is provided and includes a gas turbine engine to generate power, a heat recovery steam generator (HRSG) to produce steam from high energy fluids produced from the generation of power in the gas turbine engine, a steam turbine engine to generate additional power from the steam produced in the HRSG and a thermal load reduction system to reduce thermal loading of components of the HRSG and/or the steam turbine engine during at least startup and/or part load operations, which includes an eductor by which a mixture of compressor discharge air and entrained ambient air is injectable into the HRSG and/or an attemperator to cool superheated steam to be transmitted to the steam turbine engine.

Abstract: In one aspect, an embodiment of the present disclosure provides an Integrated Gasification Combined Cycle (IGCC) apparatus. The apparatus includes a saturator configured to saturate NPG with water vapor, and a heat recovery steam generator (HRSG), a low pressure steam loop through the saturator, wherein the HRSG is configured to heat the low pressure steam loop. The apparatus further includes a compressor and a heat exchanger configured to heat the NPG using waste process heat and extraction air from the compressor, wherein the heated NPG thereby becomes diluent nitrogen.

Abstract: The disclosed overboard extractor provides overboard extraction of compressor fluid at multiple extraction points. The overboard extractor includes multiple extraction manifolds and a delivery manifold for extracting the compressor fluid from the multiple extraction points. The overboard extractor also includes extraction valves and a delivery valve to control the extraction flow through each extraction manifold. The overboard extractor can operate in a delivery mode, a closed loop mode, or a mixed mode.

Abstract: A bleed valve module for mounting to a support structure within an aircraft includes a pipe, a plurality of bleed valves connected to the pipe, a plurality of pneumatic actuators connected to the plurality of bleed valves, and a vibration isolating element. The pipe contains a flow of bleed air and includes a plurality of inlets and an outlet. The plurality of bleed valves controls the flow of bleed air in the pipe. The plurality of pneumatic actuators actuate the plurality of bleed valves. The vibration isolating element connects the bleed valve module to the support structure within the aircraft and isolates the bleed valve module from vibration.

Abstract: A cooling circuit of a gas turbine passes an airflow through a combustor section that includes a plurality of mixing tubes for transporting a fuel/air mixture and a perforated plate including a plurality of impingement holes and a plurality of tube holes for accommodating the mixing tubes. The tube holes and the mixing tubes form a plurality of annulus areas between the perforated plate and the mixing tubes. The impingement holes and the annulus areas are configured to pass the airflow through the perforated plate. A flow management device modifies an effective size of the annulus areas to control a distribution of the airflow through the impingement holes and the annulus areas of the perforated plate to enhance cooling efficiency.

Abstract: A method for cooling components of a gas turbine engine, having first and second sets of components that require cooling, includes removing thermal energy from cooling air, directing the cooling air from which thermal energy has been removed to the first set of components of the gas turbine engine, and transferring thermal energy away from the second set of components of the gas turbine engine using a vaporization cycle of a vapor cooling assembly. Use of the vaporization cycle offsets thermal energy removed from the cooling air.

Abstract: A system includes a gas turbine, an inlet bleed circuit, and a controller. The gas turbine includes a compressor and a turbine. The compressor is configured to produce pressurized air and bleed air. The turbine is configured to produce a first output. The inlet bleed circuit includes a turbo-expander configured to produce a second output from a non-zero first portion of the bleed air. The inlet bleed circuit is also configured to direct a part of the bleed air to an inlet of the compressor. The controller is configured to adjust the gas turbine and the inlet bleed circuit to control the second output of the turbo-expander.

Abstract: A turbine power generation system with enhanced cooling provided by a stream of carbon dioxide from a carbon dioxide source and a method of using a stream of carbon dioxide to cool hot gas path components. The turbine power generating system includes a compressor, a combustor, a turbine, a generator, and at least one shaft linking the compressor and turbine and generator together such that mechanical energy produced from the turbine is used to drive the compressor and the generator. Carbon dioxide that is sequestered from the exhaust of the turbine may be stored and injected back into the turbine to cool hot gas path components of the turbine.

Abstract: An aircraft engine includes a precooler disposed within a nacelle. The precooler defines an ambient air passage and a bleed air passage. Air passes through the precooler from an ambient air inlet to an ambient air outlet. Bleed air passes through the precooler from a bleed air inlet to a bleed air outlet. Heat is transferred between the air and the bleed air via heat exchange within the precooler. At least one pressure relief door disposed proximate to the air outlet of the engine. A controller is operatively connected to the pressure relief door. The controller at least opens the pressure relief door based on a demand for increased flow of air through the ambient air passage.

Abstract: An example method of thermal energy exchange within a turbomachine includes heating a liquid using thermal energy from compressed air, injecting the liquid into a first portion of a turbomachine, and using the compressed air to cool a second portion of the turbomachine.

Abstract: Provided in one embodiment is a two-shaft gas turbine that exhibits improved reliability, output power, and efficiency. The turbine operates stably by establishing a balance between the driving force of a compressor and the output power of a high-pressure turbine in the case where the two-shaft gas turbine is applied to a system, in which the flow rate of a fluid flowing into a combustor is higher than a simple cycle gas turbine. A portion of the fluid driving the high-pressure turbine is allowed to flow not into the high-pressure turbine but into a low-pressure turbine.

Abstract: A method and apparatus for controlling fluid flow is provided. The valve assembly includes a valve body including a longitudinal axis and a radially inner internal flow passage extending axially between an inlet opening and an outlet opening, a disk selectively positionable to modulate a flow of working fluid through the flow passage, and a radially outer external surface. The valve assembly further includes a first actuator positioned radially outboard of the external surface, a second actuator positioned radially outboard of the external surface, and a first linkage including a first actuator end configured to couple to the first actuator, a second actuator end configured to couple to the second actuator, and a rod end configured to couple to the disk, wherein an actuating force applied from each actuator to each respective actuator end is combined through the first linkage to the rod end.

Abstract: A cooling air inlet which is particularly suitable for use in an engine bleed air system of an aircraft air conditioning system. The cooling air inlet comprises a cooling air inlet opening which opens into an inlet portion of a cooling air line. A cooling air inlet valve which is positionable in a first open position or a second open position is arranged in the inlet portion of the cooling air line. The cooling air inlet valve is adapted to effect the build-up of a first cooling air admission pressure in the inlet portion of the cooling air line in its first open position and the build-up of a second cooling air admission pressure in the inlet portion of the cooling air line in its second open position. The second cooling air admission pressure is greater than the first cooling air inlet pressure.