EP2513431A1 - Agencement et procédé pour le refroidissement à circuit fermé d'un composant de moteur de turbine à gaz - Google Patents

Agencement et procédé pour le refroidissement à circuit fermé d'un composant de moteur de turbine à gaz

Info

Publication number
EP2513431A1
EP2513431A1 EP09852351A EP09852351A EP2513431A1 EP 2513431 A1 EP2513431 A1 EP 2513431A1 EP 09852351 A EP09852351 A EP 09852351A EP 09852351 A EP09852351 A EP 09852351A EP 2513431 A1 EP2513431 A1 EP 2513431A1
Authority
EP
European Patent Office
Prior art keywords
cooling
medium
gas turbine
cooling medium
turbine engine
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Withdrawn
Application number
EP09852351A
Other languages
German (de)
English (en)
Inventor
Anders Hellgren
Hans MÅRTENSSON
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
GKN Aerospace Sweden AB
Original Assignee
Volvo Aero AB
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Volvo Aero AB filed Critical Volvo Aero AB
Publication of EP2513431A1 publication Critical patent/EP2513431A1/fr
Withdrawn legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01K—STEAM ENGINE PLANTS; STEAM ACCUMULATORS; ENGINE PLANTS NOT OTHERWISE PROVIDED FOR; ENGINES USING SPECIAL WORKING FLUIDS OR CYCLES
    • F01K7/00—Steam engine plants characterised by the use of specific types of engine; Plants or engines characterised by their use of special steam systems, cycles or processes; Control means specially adapted for such systems, cycles or processes; Use of withdrawn or exhaust steam for feed-water heating
    • F01K7/16—Steam engine plants characterised by the use of specific types of engine; Plants or engines characterised by their use of special steam systems, cycles or processes; Control means specially adapted for such systems, cycles or processes; Use of withdrawn or exhaust steam for feed-water heating the engines being only of turbine type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/16—Arrangement of bearings; Supporting or mounting bearings in casings
    • F01D25/162—Bearing supports
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
    • F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D9/00—Stators
    • F01D9/06—Fluid supply conduits to nozzles or the like
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01K—STEAM ENGINE PLANTS; STEAM ACCUMULATORS; ENGINE PLANTS NOT OTHERWISE PROVIDED FOR; ENGINES USING SPECIAL WORKING FLUIDS OR CYCLES
    • F01K27/00—Plants for converting heat or fluid energy into mechanical energy, not otherwise provided for
    • F01K27/02—Plants modified to use their waste heat, other than that of exhaust, e.g. engine-friction heat
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F02C7/16—Cooling of plants characterised by cooling medium
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/205—Cooling fluid recirculation, i.e. after cooling one or more components is the cooling fluid recovered and used elsewhere for other purposes
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • This invention relates to an arrangement for cooling a gas turbine engine component according to the preamble of claim 1.
  • the invention relates to a load-carrying ring-structured component arranged to guide a main engine gas flow in a hot turbine part of a gas turbine engine.
  • the invention also relates to a gas turbine engine provided with such a component and a method for cooling such a component.
  • Turbine components of a gas turbine engine such as turbine casings, vanes and blades, are exposed to very high temperatures. Cooling of turbine components are often required to avoid material fatigue and failure.
  • turbine vanes comprises a load-carrying strut or shaft and an airfoil where the strut connects an inner structure of the gas turbine engine with an outer casing and where the airfoil is arranged around the strut for guiding a general gas flow through the engine.
  • compressed air tapped off from the compressor of the gas turbine engine is guided to the vane and led in a radial direction through a space between the strut and the airfoil.
  • the airfoil and the cooling air protect the load carrying strut from being overheated.
  • US 6261054 refers to cooling requirements for vanes/blades of a gas turbine engine and shows an airfoil/vane provided with radially directed cooling channels through which a cooling medium flows.
  • the cooling medium is in the form of air from an off-board source, steam or compressed air from earlier stage of gas turbine engine.
  • the airfoil/vane is arranged between flanges that guide and distribute the cooling medium to and from the airfoil/vane in such a way that the cooling medium flows through the vane in one direction and then back through the vane in an opposite direction before being discharged.
  • the term "closed-circuit" is used to describe that the flow is re-directed and passes through the vane in both directions.
  • An object of this invention is to provide means for cooling a gas turbine engine component in a more efficient way compared to conventional technique. This object is achieved by the arrangement defined by the technical features contained in independent claim 1. It is also achieved by the engine and method defined in claims 17 and 18.
  • the dependent claims contain advantageous embodiments, further developments and variants of the invention.
  • the invention concerns an arrangement for cooling a gas turbine engine component, said arrangement comprising: a gas turbine engine component provided with at least one cooling channel through which a cooling medium is intended to flow during operation of the arrangement, a feeding system configured to supply cooling medium to the cooling channel, a cooling channel inlet, and a cooling channel outlet.
  • the invention is characterized in that the feeding system is arranged in flow communication with both the inlet and the outlet of the cooling channel such as to form a closed flow system.
  • the inventive arrangement makes water and other liquid media suitable as cooling medium. Water is a much more effective cooling medium than e.g. air that commonly is used.
  • the invention further provides possibilities to extract useful energy from the arrangement, e.g. by producing useful heat or electricity from the heat extracted when cooling the heated cooling medium.
  • the invention enables the use of a more energy efficient cooling system. This is in turn useful for increasing the overall energy efficiency, i.e.
  • cooling medium e.g. is generated by tapping off air from a compressor and where the cooling medium simply is discharged after use (i.e. after having taken up heat).
  • a more efficient cooling e.g. achieved by using water as cooling medium, is advantageous in that it reduces the need for additional structures (additional heat shields etc.) and reduces the temperature of the cooled part which in turn increases the durability and reduces the need for using high-temperature resistant materials.
  • the feeding system comprises a heat extraction device configured to extract thermal energy from the cooling medium.
  • a suitable device for this purpose is a steam engine, such as a turbine driving a generator that produces electricity.
  • a heat exchanger that transfers heat from the cooling medium to a second medium for the purpose of e.g. pre-heating fuel or producing hot water or steam, which in turn can be used for e.g. heating or for generating electricity.
  • the heat extraction device is a heat exchanger arranged to transfer heat from the cooling medium to a second medium.
  • the second medium is arranged to form a second closed flow system, wherein the second medium is allowed to evaporate in the heat exchanger when the heat is transferred from the cooling medium, wherein the second closed flow system comprises a steam engine arranged to be driven by the evaporated second medium.
  • the second closed flow system preferably comprises, in flow order from an inlet to an outlet of the heat exchanger: a turbine in which second medium that has evaporated during or after its transport through the heat exchanger is allowed to expand and thereby drive the turbine, a condenser in which the second medium is condensed to a liquid form, and a pump for feeding second medium in liquid form to the heat exchanger inlet, wherein the turbine is operatively connected to a generator capable of generating electricity.
  • the heat extraction device is a steam engine configured to be driven by evaporated cooling medium.
  • the feeding system comprises, in flow order from cooling channel outlet to inlet: a cooling medium turbine in which cooling medium that has evaporated during or after its transport through the cooling channel is allowed to expand and thereby drive the cooling medium turbine; a condenser in which the cooling medium is condensed to a liquid form, and a pump for feeding cooling medium in liquid form to the cooling channel inlet, wherein the cooling medium turbine is operatively connected to a generator capable of generating electricity.
  • the gas turbine engine component comprises an inner ring element and an outer ring element connected by a plurality of circumferentially spaced elements extending in a radial direction of the ring elements, wherein the component is arranged to transmit load or thrust from a main shaft to a casing of a gas turbine engine.
  • the outer ring element comprises an outer annular part and an inner annular part which outer and inner parts are connected by, in relation to the gas turbine engine component, radially and axially extending load-carrying wall elements, wherein axially extending cooling channels are formed between the wall elements.
  • the structure of the outer ring element has a multifunction in that it provides both a load carrying function and channels suitable for being used as cooling channels.
  • the cooling arrangement is arranged in association with a gas turbine engine that, during operation of the engine, generates i) a first main gas flow that passes at an inside of an annular inner casing through e.g. a turbine of the engine and ii) a second main gas flow that passes at an outside of said annular casing, such as a turbojet engine, wherein the condenser is arranged in relation to the second main gas flow in such a way that the second main gas flow is allowed to cool and condense the second medium or the cooling medium.
  • Air with high velocity, as in the second main gas stream, provides for efficient heat transfer which makes it possible to make use of a small heat exchanger as condenser. Further, the heat transferred from the heated medium to the second airflow increases the thrust of a turbojet engine since increased temperature gives a greater volume flow.
  • the invention also concerns a gas turbine engine provided with a component of the above type as well as a method for cooling such a component.
  • Figure 1 shows, in a schematic view, a turbo-jet engine provided with a component according to the invention
  • Figure 2 shows, in a perspective and partly schematic view, an embodiment of the inventive cooling arrangement
  • Figure 3 shows, in a partly sectional and partly schematic view, another embodiment of the inventive cooling arrangement
  • Figures 4-9 show an alternative design of the cooling channels for the inventive cooling arrangement.
  • Fig. 1 shows a turbojet engine.
  • the turbojet engine comprises a central body 1 , an annular outer casing 2 (fan casing), an annular inner casing 3 (engine casing), a fan or blower 4, a low pressure compressor 5, a high pressure compressor 6, a combustion chamber 7, a high pressure turbine 8 and a low pressure turbine 9.
  • It further comprises a set of arms 10 extending in a radial direction from the inner casing 3 to an outer ring element 14 forming part of the outer casing 2.
  • the arms 10 comprise aerodynamic vanes 11 primarily provided to act as guide vanes for air passing through the annular channel between the inner casing 3 and the outer casing 2 in an axial direction, i.e. a longitudinal direction, of the engine.
  • the arms 10 further comprise structural arms or load carrying vanes 12 primarily provided to guarantee a certain mechanical strength of the construction.
  • a gas turbine engine component 20 associated with a cooling arrangement according to the invention is in this example positioned between the high pressure turbine 8 and the low pressure turbine 9.
  • said component 20 comprises an inner ring element 22 and an outer ring element 24 connected by a plurality of circumferentially spaced elements 26 (vanes) extending in a radial direction of the ring elements 22, 24.
  • the inner ring element 22 forms part of the central body 1
  • the outer ring element 24 forms part of the annular inner casing 3.
  • the component 20 is a load carrying structure that transmits load from an engine shaft/bearing housing via the inner ring element 22 and the vanes 26 to the outer ring element 24.
  • the flow of gas through the turbojet engine is divided into two major streams, a first one of which passes through an annular channel between the central body 1 and the inner casing 3, and passes the compressors 5, 6, the combustion chamber 7 and the turbines 8, 9.
  • the first stream passes through the component 20 in an axial direction thereof; through the spaces between the inner and outer ring elements 22, 24 and between the vanes 26.
  • the component 20 thus defines a gas channel for a main gas flow through a gas turbine engine.
  • a second stream passes through the annular channel between the inner casing 3 and the outer casing 2.
  • a temperature of the second stream is in operation lower than a temperature of the first stream, but the second stream substantially increases the thrust of the turbojet engine.
  • An engine mount (not shown in figure 1) is arranged onto the outer ring element 14 by means of which the turbojet engine is attached to and held in position in relation to an aircraft.
  • FIG. 2 shows in a perspective and partly schematic view, an embodiment of the inventive cooling arrangement.
  • the cooling arrangement comprises the above mentioned gas turbine component 20 that in turn comprises the inner and outer ring elements 22, 24 connected by the circumferentially spaced elements 26 (vanes) extending in the radial direction of the ring elements 22, 24.
  • the component 20 is further provided with a system of cooling channels that can be arranged in different ways inside the component 20 (see figures 3-10).
  • a cooling medium in the form of water flows through the cooling channels during operation of the arrangement.
  • the arrangement further comprises a main cooling channel inlet 27 configured to provide an inlet for the cooling medium to the cooling channels arranged in the component 20, and a main cooling channel outlet 28 configured to provide an outlet for the (heated) cooling medium from the cooling channels arranged in the component 20.
  • the main inlet 27 is arranged at an annular inlet manifold 30 that extends in a circumferential direction around the component 20 along the outer ring element 24 and that distributes cooling medium from the main inlet 27 to a plurality of cooling channels arranged inside the component 20 (see figure 3).
  • the main outlet 28 is arranged at an annular outlet manifold 31 that extends in a circumferential direction around the component 20 along the outer ring element 24 and that guides cooling medium from the plurality of cooling channels arranged inside the component 20 to the main outlet 28 (see figure 3).
  • the annular inlet and outlet manifolds 30, 31 are arranged at a distance from each other as seen in an axial direction of the component 20. As shown in figure 3, this allows cooling channels to be arranged in an axial direction in fluid communication with both manifolds 30, 31. Also the manifolds 30, 31 may be regarded as cooling channels.
  • An internal wall section 25 forms a closing end section in each of the manifolds and prevents the cooling medium from taking a short-cut from the inlet 27 to the outlet 28.
  • the wall section 25 forces the cooling medium to pass almost 360° around the component 20.
  • the arrangement further comprises a feeding system 40 configured to supply cooling medium to the system of cooling channels.
  • the feeding system 40 is arranged in flow communication with both the main inlet 27 and the main outlet 28 of the cooling channel system such as to form a closed flow system. This means that the cooling medium, in this case the water, that flows out through the main outlet 28 is re-circulated such as to also flow in through the main inlet 27.
  • the arrangement is designed in such a way that the cooling water evaporates while flowing through the cooling channels.
  • the water is generally in liquid form when passing the main inlet 27 and generally in gas form when passing the main outlet 28.
  • Various designs may be suitable depending on e.g. the type of component, where the component is positioned in the gas turbine engine and type of engine. For a given component in a given engine, the degree of evaporation can be adjusted by adjusting the mass flow rate of the cooling medium.
  • the feeding system 40 further comprises, in flow order from cooling channel outlet to inlet 28, 27,: a cooling medium turbine 41 in which cooling medium that has evaporated during its transport through the cooling channel and manifolds 30, 31 is allowed to expand and thereby drive the cooling medium turbine 41 ; a condenser 42 in which the cooling medium is condensed to a liquid form, and a pump 43 for feeding cooling medium in liquid form to the cooling channel inlet 27.
  • the cooling medium turbine 41 is operatively connected to a generator 44 for generating electricity.
  • the feeding system 40 comprises a heat extraction device in the form of a turbine 41 configured to extract thermal energy from the cooling medium.
  • the turbine 41 converts this thermal energy to kinetic (rotational) energy which in turn is converted to electricity by the generator 44.
  • This electricity can, for instance, be used in electric systems of an aircraft provided with the inventive arrangement and the associated gas turbine engine as to replace electricity generated by an on-board APU or by the gas turbine engine (via the power take-off shaft).
  • the turbine 41 and the generator 44 together forms an energy conversion device configured to convert cooling medium thermal energy to another (useful) form of energy, in this case electricity.
  • Cooling water can be kept in liquid form during its transport through the cooling channels by keeping the pressure sufficiently high. Evaporation can in such a case be brought about by decreasing the pressure before the water reaches the turbine 41 , e.g. by letting the cooling water pass a release valve arranged between the outlet 28 and the turbine 41.
  • the condenser 42 is a heat exchanger, and thus also a form of heat extraction device, where a secondary cooling medium is used to cool and condense the primary cooling medium used to cool the component 20.
  • the condenser 42 is arranged in the annular channel between the inner casing 3 and the outer casing 2 such that the second (colder) gas stream works as the secondary cooling medium. Air with high velocity, as in the second gas stream, provides for efficient heat transfer which makes it possible to make use of a small heat exchanger as condenser. Further, the heat transferred from the (primary) cooling medium to the second airflow increases the thrust of the turbojet engine since increased temperature gives a greater volume flow.
  • the condenser/heat exchanger 42 can form part of or be integrated in the fan outlet guide vanes or other already existing static fan structure.
  • FIG. 3 shows, in a partly sectional and partly schematic view, another embodiment of the inventive cooling arrangement.
  • This embodiment is in principle similar to the embodiment shown in figure 2 and therefore the same reference numbers has been used for the majority of parts involved.
  • the gas turbine component has been given reference number 200 as compared to 20 in figure 2 as to indicate that the cooling arrangement exemplified in figure 3 concerns only a circumferential section of the outer ring element 24, i.e. the section delimited by wall sections 25a and 25b.
  • This section can, in principal, have any length in the circumferential direction of the outer ring element 24, which means that the section e.g. can extend 360° around the outer ring element 24 such that the wall sections 25a and 25b coincide and such that the embodiment in figure 3 becomes similar to the embodiment shown in figure 2.
  • the outer ring element 24 of the gas turbine component 200 is provided with a plurality of parallel and, in relation to the component 200, axially extending cooling channels 50 separated by wall elements 55.
  • the outer ring element 24 is a load carrying structure comprising an outer annular part and an inner annular part which outer and inner parts are connected by the radially (and axially) extending wall elements 55.
  • the cooling channels 50 are thus formed between the load carrying wall elements 55.
  • the structure of the outer ring element 24 thus provides both a load carrying function and channels suitable for being used as cooling channels.
  • an inlet manifold 30 provided with a main inlet 27 is arranged at an inlet side of the cooling channels 50 and an outlet manifold 31 provided with a main outlet 28 is arranged at an outlet side of the cooling channels 50.
  • Each cooling channel 50 is arranged in fluid communication with the inlet and outlet manifolds 30, 31 by means of inlet openings 51 and outlet opening 52, respectively.
  • the closed flow system for the cooling medium is indicated with an arrowed line and can be described in the following way:
  • the cooling medium in this case water
  • the gas turbine component 200 becomes heated by the hot gases flowing through the component 200. Heat absorbed by (the section of) the outer ring element 24 is transferred to the cooling water flowing through the cooling channels 50 which in this case results in that the water evaporates.
  • the cooling medium now in the form of steam, enters the outlet manifold 31 via outlet opening 52.
  • the steam flows inside the inlet manifold 30 along its direction of extension (circumferentially along the outer ring element 24) until it reaches the main outlet 28 at the end of the outlet manifold 31.
  • the steam flows to and drives the turbine 41 that in turn drives the generator 44 that generates electricity.
  • the steam is condensed to water which is forced to the main inlet 27 by means of the pump 43.
  • FIGS 4-9 show, for a component similar to what is described above, an alternative design of the cooling channels for the inventive cooling arrangement where cooling channels are provided not only in the outer ring element 24 but also in the vanes 26 and in the inner ring element 22.
  • Figure 4 shows a load-carrying vane 260, the shape of which is only schematic, provided with first and second inlet cooling channels 261 , 262 arranged at a leading edge and a trailing edge, respectively, thereof.
  • First and second outlet cooling channels 263, 264 are arranged along both sides (i.e. pressure and suction sides) of the vane 260. All of these cooling channels extend through the vane 260 from one end to the other in a radial direction with reference to the annular component in which the vane 260 is intended to be arranged.
  • the vane 260 is further provided with centrally located channels 265, in this case three in number, for leading e.g. oil to a bearing of the gas turbine engine.
  • Figure 5 shows the vane 260 arranged and extending radially between an inner ring element 220 and an outer ring element 240.
  • the vane 260 and the parts of the inner and outer ring elements 220, 240 shown in figure 5 form a part of a gas turbine engine component similar to the component shown in figures 1-3.
  • the outer ring element 240 has a flange 241 for attachment to adjacent components in the gas turbine engine and an engine mount 242 for mounting of the engine to e.g. an aircraft.
  • Figure 5 further shows two main inlets 270 each of which forms a passage for a cooling medium to an inlet manifold 300.
  • the component also comprises two main outlets 280 in fluid communication with outlet manifolds 310.
  • there are in this example two partly separated cooling channel systems arranged inside the component, and therefore there are two main inlets 270 and two main outlets 280.
  • the partly separated systems may be combined in different ways if only one inlet and one outlet is desired.
  • a feeding system (not shown), similar to the feeding system 40 described above, is arranged in flow communication with the main inlets 270 and the main outlets 280 of the cooling channel system such as to form a closed flow system.
  • the flows of cooling medium through the two separate cooling channel systems are combined outside of the component so that one combined flow of cooling medium is supplied to and from the feeding system.
  • the cooling medium enters the inlet manifold 300 via the main inlet 270 and flows further through the trailing (or leading) edge of the vane towards the inner ring element 220.
  • Figure 6 shows, with arrows, how the cooling medium flows via the inlet manifolds 300 to the first or second inlet cooling channels 261 , 262 arranged at the leading and trailing edge, respectively, of the vane 260.
  • relatively cool cooling medium is guided to the leading edge of the vane 260 which is an advantage since the leading edge is likely to be the part of the vane 260 that requires the most efficient cooling during operation of the gas turbine engine.
  • Figure 6 also shows how the cooling medium leaves the first and second outlet cooling channels 263, 264 of the vane 260 and enters the outlet manifolds 310.
  • Figure 7 shows that a first cavity 222 is arranged in the inner ring element 220 below the vane 260 at a position where the second inlet cooling channel 262 ends such that cooling medium flowing through the second inlet channel 262 can enter the first cavity 222.
  • An opening 224 provides a further fluid communication to inner ring cooling channels 225 arranged in the inner ring element 220.
  • a further opening 228 provides for fluid communication between the inner ring cooling channels 225 and a second cavity 226 arranged below and in fluid communication with the first outlet cooling channel 263 of the vane 260.
  • a still further opening 223 can be arranged between the first and second cavities 222, 226 as to (partly) bypass the inner ring cooling channels 225.
  • the cooling medium enters the inlet manifold 300 via the inlet 270 and flows in a radial direction through the second inlet cooling channel 262 of the vane 260 and further via the first cavity 222 and opening 224 into the inner ring cooling channels 225. Via the further opening 228 the cooling medium enters the second cavity 226 and flows further in a radial direction through the second outlet cooling channel 263 of the vane 260.
  • This channel 263 ends at the outlet manifold 310 which guides the cooling medium to the outlet 280.
  • the cooling medium is, in this example, in evaporated form.
  • the medium then enters the feeding system where its heat is utilized, in this case for generating electricity as described above. Finally, the cooling medium, in liquid form, is fed back to the inlet 270.
  • the other partly separated flow of cooling medium through the component is in this example in principal similar to the flow described above but makes instead use of e.g. the first inlet cooling channel 261 , a third cavity 222' corresponding to the first cavity 222, second inner ring cooling channels 225' and the second outlet cooling channel 264 of the vane 260. This other partly separated flow also flows through a further cavity corresponding to the second cavity 226. The two flows are mixed in the two second cavities 226.
  • Figure 8 shows part of what is described above and also that the flow that leaves the third cavity 222' is split so that half the flow goes through the second inner ring cooling channels 225' and the other half through similar channels arranged on the other (left and nearer) side of the position of the vane 260.
  • the outlet manifold 310 may be arranged as to combine the flows exiting the first and second outlet cooling channels 263, 264 of the vane 260 and may be provided with only one outlet 280. Further, the outlet manifold 310 may extend circumferentially along the outer ring 240 such as to receive flows from several vanes 260. Also the inlet manifolds 300 may extend along the outer ring 240 and be arranged such as to feed several vanes 26 with a flow of cooling medium.
  • Figure 9 shows some of the parts described above but in another view.
  • Figure 9 also shows a load-carrying structure 316 that connects a bearing house 317 to the inner ring element 220.
  • the bearing house 317 is arranged in association with a main shaft of the gas turbine engine.
  • the gas turbine engine component is arranged to transmit load or thrust from the engine shaft to e.g. an aircraft via the bearing house 317, the load-carrying structure 316, the inner ring element 220, the vane 260 (not shown in figure 9), the outer ring element 240 and the engine mount 242.
  • An advantage of the inventive cooling arrangement according to figures 4-9 compared to the traditional solution with a two-part turbine vane structure that includes a first load-carrying part and a second heat-protecting/air-guiding part, is that only one part (with at least one cooling channel) is needed which provides for a more cost-efficient production.
  • the gas turbine component of the present invention is a gas turbine engine housing component located in the turbine region of the engine. Such a component must be configured to withstand a high thermal load from a passing gas flow during operation of the gas turbine engine.
  • the typical component can also be described as a gas turbine component that defines a gas channel for a main gas flow through a gas turbine engine or as a gas turbine component that is configured so that an external surface of the gas turbine component is in direct contact with a passing main gas flow through a gas turbine engine. Since such a main gas flow has its highest temperature in the region of the turbine (downstream of the combustion chamber) said component is typically arranged in the turbine region of a gas turbine engine.
  • the cooling medium can be pressurized so that no evaporation of cooling medium occurs.
  • Heat can in such a case be extracted from the cooling medium by using a heat exchanger that is arranged to transfer heat from the cooling medium to a second medium.
  • a heat exchanger is arranged in association with a flow connection between the between the inlet 27, 270 and outlet 28, 280.
  • This second medium can be air (e.g. by using the second gas stream as the second medium as described above for the condenser 42 or by using another flow of air for the purpose of heating the cabin space) or fuel (for preheating purposes) or another medium.
  • the second medium is arranged to form a second closed flow system wherein the second medium is evaporated in the heat exchanger when the heat is transferred from the cooling medium.
  • the second closed flow system comprises, in flow order from heat exchanger outlet to inlet, the turbine 41 (and the generator 44), the condenser 42 and the pump 43, which work in the same way as described above but where the second medium, instead of the cooling medium, drives the turbine etc.
  • the closed cooling flow system preferably comprises a cooling medium pump for circulating the cooling medium.
  • the second medium is suitably water.
  • the cooling medium that is to flow through the cooling channels of the component, which cooling medium in this variant is not intended to evaporate, can be water but other substances may now be used, such as oil.
  • turbine 41 and generator 44 can be replaced by another form of steam engine.

Abstract

La présente invention concerne un agencement pour refroidir un composant de moteur de turbine à gaz, ledit agencement comprenant : un composant de moteur de turbine à gaz (20, 200) équipé d'au moins un canal de refroidissement (50, 261-264, 225) à travers lequel un support de refroidissement doit s'écouler lors du fonctionnement de l'agencement, un système d'alimentation (40) configuré pour fournir un support de refroidissement au canal de refroidissement (50, 261-264, 225), une entrée de canal de refroidissement (27, 51, 270), et une sortie de canal de refroidissement (28, 52, 280). L'invention est caractérisée en ce que le système d'alimentation (40) est agencé en communication fluidique à la fois avec l'entrée (27, 51, 270) et avec la sortie (28, 52, 280) du canal de refroidissement (50, 261-264, 225) de façon à former un système à circuit fermé. L'invention concerne également un moteur de turbine à gaz équipé d'un tel composant et un procédé pour refroidir ce composant.
EP09852351A 2009-12-17 2009-12-17 Agencement et procédé pour le refroidissement à circuit fermé d'un composant de moteur de turbine à gaz Withdrawn EP2513431A1 (fr)

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PCT/SE2009/000526 WO2011075013A1 (fr) 2009-12-17 2009-12-17 Agencement et procédé pour le refroidissement à circuit fermé d'un composant de moteur de turbine à gaz

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EP2513431A1 true EP2513431A1 (fr) 2012-10-24

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Families Citing this family (36)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN102943711B (zh) * 2012-11-12 2015-02-11 湖南航翔燃气轮机有限公司 冷却涡轮导向器组的装置
US9739171B2 (en) 2012-11-16 2017-08-22 United Technologies Corporation Turbine engine cooling system with an open loop circuit
US20140165570A1 (en) * 2012-12-18 2014-06-19 United Technologies Corporation Oscillating heat pipe for thermal management of gas turbine engines
US10036258B2 (en) 2012-12-28 2018-07-31 United Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US10018052B2 (en) 2012-12-28 2018-07-10 United Technologies Corporation Gas turbine engine component having engineered vascular structure
US9422063B2 (en) * 2013-05-31 2016-08-23 General Electric Company Cooled cooling air system for a gas turbine
US9617917B2 (en) 2013-07-31 2017-04-11 General Electric Company Flow control assembly and methods of assembling the same
US20150107249A1 (en) * 2013-10-22 2015-04-23 Access Energy Llc Extracting Heat From A Compressor System
ES2674241T3 (es) * 2014-03-13 2018-06-28 Bae Systems Plc Intercambiador de calor
EP2987966A1 (fr) 2014-08-21 2016-02-24 Siemens Aktiengesellschaft Turbine à gaz dotée de canal de refroidissement divisé en sections annulaires
US20160102566A1 (en) * 2014-10-13 2016-04-14 Pw Power Systems, Inc. Power turbine air strut
FR3028576B1 (fr) * 2014-11-19 2016-12-23 Snecma Secteur d'aubage de stator d'une turbomachine comprenant des canaux de circulation de fluide chaud
US10094287B2 (en) 2015-02-10 2018-10-09 United Technologies Corporation Gas turbine engine component with vascular cooling scheme
FR3034474B1 (fr) * 2015-04-01 2019-08-09 Safran Aircraft Engines Turbomachine equipee d'un secteur d'aubage et d'un circuit de refroidissement
US10077664B2 (en) 2015-12-07 2018-09-18 United Technologies Corporation Gas turbine engine component having engineered vascular structure
US11125160B2 (en) * 2015-12-28 2021-09-21 General Electric Company Method and system for combination heat exchanger
US10221694B2 (en) 2016-02-17 2019-03-05 United Technologies Corporation Gas turbine engine component having vascular engineered lattice structure
US10260523B2 (en) 2016-04-06 2019-04-16 Rolls-Royce North American Technologies Inc. Fluid cooling system integrated with outlet guide vane
US10247106B2 (en) * 2016-06-15 2019-04-02 General Electric Company Method and system for rotating air seal with integral flexible heat shield
US10443436B2 (en) 2016-07-01 2019-10-15 General Electric Company Modular annular heat exchanger
FR3054263B1 (fr) * 2016-07-20 2018-08-10 Safran Aircraft Engines Carter intermediaire de turbomachine d'aeronef realise d'une seule piece de fonderie avec une canalisation de lubrifiant
US10830056B2 (en) * 2017-02-03 2020-11-10 General Electric Company Fluid cooling systems for a gas turbine engine
US20180266325A1 (en) * 2017-03-20 2018-09-20 General Electric Company Extraction cooling system using evaporative media for stack cooling
US11242767B2 (en) * 2017-05-01 2022-02-08 General Electric Company Additively manufactured component including an impingement structure
US10815830B2 (en) 2017-12-21 2020-10-27 Raytheon Technologies Corporation Lightweight tierod
CN110388271A (zh) * 2018-04-23 2019-10-29 云南靖创液态金属热控技术研发有限公司 一种航空发动机的液态金属散热系统及航空发动机
US10774653B2 (en) 2018-12-11 2020-09-15 Raytheon Technologies Corporation Composite gas turbine engine component with lattice structure
CN111350549B (zh) * 2019-12-30 2022-07-26 中国科学院工程热物理研究所 一种适用于富燃工质涡轮高温静叶的冷却结构
US11378341B2 (en) * 2020-01-03 2022-07-05 Raytheon Technologies Corporation Gas turbine engine heat exchanger for annular flowpaths
CN111594285A (zh) * 2020-05-27 2020-08-28 中国航空发动机研究院 用于航空发动机高压涡轮导叶的并联式冷却循环系统
CN111594286A (zh) * 2020-06-01 2020-08-28 中国航空发动机研究院 用于航空发动机高压涡轮导叶的串联式冷却循环系统
GB2599691A (en) * 2020-10-09 2022-04-13 Rolls Royce Plc A heat exchanger
GB2599693B (en) * 2020-10-09 2022-12-14 Rolls Royce Plc A heat exchanger
US11867121B2 (en) * 2021-03-24 2024-01-09 General Electric Company Gas turbine engines with heat recovery systems
CN114776400B (zh) 2022-04-11 2024-02-20 北京航空航天大学 一种航空发动机涡轮机匣及导叶一体化冷却系统
US20240084733A1 (en) * 2022-09-08 2024-03-14 General Electric Company Closed-loop cooling system for a gas turbine engine

Family Cites Families (11)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3756020A (en) * 1972-06-26 1973-09-04 Curtiss Wright Corp Gas turbine engine and cooling system therefor
JPS57140502A (en) * 1981-02-23 1982-08-31 Hitachi Ltd Cooling device of rotary blade of gas turbine
US4431371A (en) * 1982-06-14 1984-02-14 Rockwell International Corporation Gas turbine with blade temperature control
US5685693A (en) * 1995-03-31 1997-11-11 General Electric Co. Removable inner turbine shell with bucket tip clearance control
KR100389990B1 (ko) * 1995-04-06 2003-11-17 가부시끼가이샤 히다치 세이사꾸쇼 가스터빈
DE19643716A1 (de) * 1996-10-23 1998-04-30 Asea Brown Boveri Schaufelträger für einen Verdichter
DE60028446T2 (de) * 1999-04-23 2006-12-21 General Electric Co. Heiz- und Kühlkreislauf für das Innengehäuse einer Turbine
US6565312B1 (en) * 2001-12-19 2003-05-20 The Boeing Company Fluid-cooled turbine blades
JP4489808B2 (ja) * 2004-05-27 2010-06-23 ボルボ エアロ コーポレイション タービン又はコンプレッサ装置内の支持構造部及びその構造部の組み立て方法
US20080245054A1 (en) * 2005-09-20 2008-10-09 Volvo Aero Corporation Cooling System for an Aircraft, Aircraft Comprising the Cooling System and Cooling Method
DE102005060704A1 (de) * 2005-12-19 2007-06-28 Rolls-Royce Deutschland Ltd & Co Kg Gasturbinenbrennkammer

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
See references of WO2011075013A1 *

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