WO2014102381A1 - Aubes de turbine et procédé de construction - Google Patents

Aubes de turbine et procédé de construction Download PDF

Info

Publication number
WO2014102381A1
WO2014102381A1 PCT/EP2013/078146 EP2013078146W WO2014102381A1 WO 2014102381 A1 WO2014102381 A1 WO 2014102381A1 EP 2013078146 W EP2013078146 W EP 2013078146W WO 2014102381 A1 WO2014102381 A1 WO 2014102381A1
Authority
WO
WIPO (PCT)
Prior art keywords
aerofoil
rotor
sintering
inlet
outlet
Prior art date
Application number
PCT/EP2013/078146
Other languages
English (en)
Inventor
Heinz Peter Berg
Steffen Sebastian KIEßLING
Original Assignee
Lux Powertrain Sa
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 Lux Powertrain Sa filed Critical Lux Powertrain Sa
Publication of WO2014102381A1 publication Critical patent/WO2014102381A1/fr

Links

Classifications

    • 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/34—Rotor-blade aggregates of unitary construction, e.g. formed of sheet laminae
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/22—Manufacture essentially without removing material by sintering
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/183—Two-dimensional patterned zigzag
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/184—Two-dimensional patterned sinusoidal
    • 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
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F05D2250/185—Two-dimensional patterned serpentine-like
    • 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
    • F05D2250/00—Geometry
    • F05D2250/80—Size or power range of the machines
    • F05D2250/82—Micromachines

Definitions

  • This invention relates to a blade and a rotor for a turbine and a method of constructing a blade.
  • Gas turbines operate by means of a combustion in a combustion chamber which produces gas.
  • the gas is directed towards a blade, causing the blade to rotate.
  • the gas is directed (by means of vanes) in a direction so that it is introduced in a direction radial to the rotation of blades.
  • the kinetic energy of the gas is then converted into work in the form of mechanical energy.
  • Axial and diagonal turbines act in an analogous manner although the gas is directed towards the turbine axially, or diagonally, as the case may be.
  • the thermal efficiency of the work cycle of the turbine is dependent upon a high turbine entry temperature, which is limited by the ability of the material from which the various constituents of the turbine blades and the nozzle guide vanes are produced.lt is therefore generally desirable to cool the blade since this allows a greater operating temperature, and therefore efficiency, for the turbine.
  • a hollow blade or vane is provided and a cooling fluid is allowed to circulate in the interior of the blade.
  • effects contributing to internal cooling e.g. "impingement cooling” where a jet of the cooling fluid is directed to impinge on an inner wall of the blade or vane; “pin fin cooling” where shapes such as cylinders are introduced into the path of the cooling fluid, thereby increasing the turbulence of the fluid and improving the heat conducted to the fluid; and “dimple cooling” where dimples are formed on an inner surface of the blade or vane, these dimples also acting to increase turbulence (this is often preferred to pin fin cooling due to the relatively small pressure drop caused).
  • blades and vanes which are formed for inner cooling suffer from the disadvantage that they display poor castability. Such blades and vanes therefore often have a significantly shorter lifespan than film cooled vanes and blades.
  • the invention provides an aerofoil for a gas turbine comprising a body having at least one inner chamber, an inlet channel formed in a portion of the body and an outlet channel, wherein the inlet and the outlet channels place the inner chamber in fluid communication with one or more environments exterior to the aerofoil to allow a cooling fluid to flow from the exterior environment into the cooling chamber, and wherein the portion of the body having the inlet channel formed therein is formed by a process of sintering.
  • Embodiments of the invention extend to both blades (those structures of the rotor which, in reaction to incoming gas or other fluid cause rotation of the rotor) as well as vanes (those structures of the turbine which direct the incoming gas onto the blades). Therefore, the term 'aerofoil' encompasses both blades and vanes.
  • the aerofoil may further comprise a plurality of inlet channels formed in corresponding portions of the body of the aerofoil wherein each of the corresponding portions are formed by a process of sintering. It has been found that when the portions of the aerofoil which define the inlets are made by a process of sintering that the temperature of the fluid which enters through the inlet can be significantly higher when compared to aerofoils constructed with other manufacturing techniques.
  • the inlet is able to withstand temperatures of 1 800 °C.
  • the maximum temperature is 1 500 °C.
  • the maximum temperature is 1 250 °C.
  • the aerofoil may further comprise a plurality of inner chambers in fluid communication with one another.
  • Adjacent inner chambers may be placed in fluid communication with one another by means of a flow channel.
  • the flow channel may have a first linear dimension D and a distance between the flow channel and an opposing wall of an inner chamber may be S, in which case the ratio S:D may be between 2:1 and 4:1 .
  • the ratio S:D may be between 3 and 3.5. In an embodiment, the ratio S:D is about 3.25.
  • Adjacent inner channels may be separated by a structural element wherein the structural element reinforces the aerofoil.
  • the structural element may comprise one or more of a matrix, a lattice, an open- cell foam structure and a closed cell-foam structure.
  • At least one inner chamber may comprise a baffle which increases a fluid path length between the inlet channel and the outlet channel.
  • the aerofoil may further comprise cooling structures formed on an inner surface of at least one of the inner chambers.
  • the outlet channel may be positioned and/or shaped so that fluid exiting the outlet channel flows over a surface of the aerofoil, thereby cooling the surface.
  • the invention further extends to a rotor for a turbine comprising one or more blades, each of the blades being aerofoils as herein described.
  • the rotor may be a rotor for a radial turbine and, in this case, may comprise an integral whole comprising an axial portion and a plurality of blades integrated into a whole.
  • the rotor is formed as an integral whole by a process of sintering.
  • the rotor may form a single part and is not assembled from a plurality of partial sections. Only sintering may be used to manufacture the rotor.
  • the rotor may have a maximum diameter of 200 mm. Alternatively, a diameter measured between extremities of the rotor may be 200 mm or less.
  • the rotor may comprise one or more aerofoils, at least one of the aerofoils having more than three portions, each portion formed by a process of sintering wherein a first portion is sintered from a first material, a second portion is sintered from a second material and a third portion is sintered from a third material.
  • the first portion is sintered from a first material
  • a second portion is sintered from a second material
  • a third portion is sintered from a third material.
  • the rotor may further comprise a gas inlet and a gas outlet wherein the portion comprised of the first material is located closer to the inlet and the portion comprised of the third material is located closer to the outlet, the portion comprised of the second material being located between the portion comprised of the first material and the portion comprised of the second material.
  • the invention further extends to a turbine comprising a rotor as hereinbefore described.
  • the invention further extends to a method of manufacturing a turbine aerofoil, the turbine aerofoil comprising a body having at least one inner chamber, an inlet channel formed in a portion of the body and an outlet channel formed in the body, wherein the inlet and the outlet channels place the inner chamber in fluid communication with one or more environments exterior to the aerofoil, the method comprising:
  • the entire aerofoil may be made by a process of sintering.
  • the method may further comprise forming a plurality of inlet channels in corresponding portions of the body of the aerofoil wherein each of the corresponding portions are formed by a process of sintering.
  • the method may further comprise forming a plurality of inner chambers in fluid communication with one another.
  • the method may further comprise forming a flow channel such that adjacent chambers are placed in fluid communication with one another by the flow channel.
  • the flow channel may have a first linear dimension D and a distance between the flow channel and an opposing wall of an inner chamber may be S, in which case the ratio S:D may be between 2:1 and 4:1 .
  • Adjacent inner channels may be separated by a structural element wherein the structural element reinforces the aerofoil.
  • the aerofoil may be formed so that the structural element comprises one or more of a matrix, a lattice, an open-cell foam structure and a closed cell-foam structure.
  • the aerofoil may be formed so that at least one inner chamber comprises a baffle where the baffle increases a fluid path length between the inlet channel and the outlet channel.
  • the aerofoil may be formed with cooling structures on an inner surface of at least one of the inner chambers.
  • the outlet may be positioned and/or shaped so that fluid exiting the outlet flows over a surface of the aerofoil, thereby cooling the surface.
  • the aerofoil may be sintered in at least three different portions, each portion comprising one or more layers, wherein a first portion is sintered from a first material, a second portion is sintered from a second material and a third portion is sintered from a third material.
  • the aerofoil may be shaped so that portions comprised of the first material are located closer to a gas inlet and portions comprised of the third material are located closer to a gas outlet, portions comprised of the second material being located between portions comprised of the first material and portions comprised of the second material. Further embodiments of the invention extend to a rotor having a plurality of aerofoils manufactured by a single sintering process from a single material.
  • the material may be chosen so that the rotor as a maximum operating temperature of 1 800 °C. In a further embodiment, the maximum operating temperature is 1 500 °C. In yet a further embodiment, the maximum operating temperature is 1 250 °C. DESCRIPTION OF ACCOMPANYING FIGURES
  • Figure 1 is schematic diagram of a turbine rotor according to an embodiment of the invention
  • Figure 2 is a schematic diagram of a portion of a cross-section of a turbine rotor according to a further embodiment of the invention.
  • Figure 3 is a schematic diagram of a portion of a turbine wall of a turbine rotor according to an embodiment of the invention
  • Figure 4 is a schematic diagram of a portion of a cross-section of a turbine rotor according to a further embodiment of the invention.
  • Figure 5 is a process of constructing a turbine rotor according to an embodiment of the invention.
  • Figure 1 illustrates a turbine rotor 10 which includes a plurality of turbine blades 12 arranged to rotate about an axis 14.
  • the rotor 10 includes a number of inlet channels 16 to allow air to enter the rotor.
  • a seal isolates the inlet channel from the gas causing rotation of the rotor 10.
  • the blades are hollow and, since the inlet channels 16 are formed in an outer surface of the blades 12, air (or other cooling fluid in further embodiments) enters through the inlet channels 16 from an external environment and circulates in the blades 12 thereby cooling them.
  • the rotor 10 further comprises an outlet (not shown in this Figure) which allows the air to exit the rotor.
  • the rotor 10 is an integrated whole, the axial portion and the blades are formulated as an integrated part. This is particularly useful for radial turbines.
  • the rotor 10 is constructed by means of a laser sintering process described in greater detail below with reference to Figure 4. The rotor is therefore manufactured by depositing successive layers of material.
  • the rotor 10 is manufactured the rotor is manufactured by being built up, layer on layer, from the left to the right in the orientation shown in Figure 1 .
  • the rotor 10 is manufactured by layer deposition in alternate directions. It is to be realised that there are a number of advantages to constructing a blade for a turbine using sintering.
  • thermal properties of materials used for sintering are particularly well-suited for turbine blade construction and, by extension, rotor construction too.
  • the thermal properties of such materials have been found to allow for a significant increase in the inlet temperature of the turbine.
  • an inlet temperature of 1 800 °C is achievable whereas in others, the maximum temperature is 1 500 °C or 1250 °C.
  • the rotor 10 illustrated in Figure 1 is entirely constructed by using a sintering process.
  • sintering is not limited to constructing the entire rotor, or even an entire blade or vane, by sintering. Instead, only a portion of the rotor or blade may be constructed with this process. It has been found that the area surrounding the inlet is the area which is most likely to fail with cast rotors or blades. As such, in a further embodiment of the invention only the portion of a blade, such as the portion 18 shown in dotted outline in Figure 1 , is the portion of the blade constructed by a process of sintering.
  • Figure 2 illustrates a cross-section through a blade 20 from a rotor similar to the rotor 10 illustrated in Figure 1 .
  • the blade 20 includes a body 21 having a leading edge 22 and a trailing edge 24.
  • the leading edge 22 is the edge of the body 21 which first impacts the medium in which the rotor rotates.
  • the body 21 has nine inner chambers formed therein: chambers 31 , 32, 33, 34, 35, 36, 37, 38 and 39.
  • the blade 20 further comprises an inlet channel 26 formed closer to the leading edge 22 than to the trailing edge 24.
  • the inlet chamber has two exit ports 27 and 29. Exit port 27 introduces cooling fluid into chamber 32 and exit port 29 introduces cooling fluid into chamber 34.
  • the blade 20 further comprises an outlet channel 28. As the blade 20 rotates air (in the embodiment shown, although other embodiments may use other cooling fluids such as oil, steam or water) is introduced into the inlet channels 26, enters the inner chambers 32 and 34.
  • the cooling fluid is able to flow from the inlet channels, through the inner chambers 31 , 32, 33, 34, 35, 36, 37, 38 and 39, and out of the outlet channels 28, thereby cooling the blade 20.
  • the outlet channels 28 are situated close to the rear of the blade; air exiting through the outlet channel will then exit the chamber.
  • the outlet channel or channels of the blade may be located further towards an inlet of the turbine and directed so that air exiting from the outlet channel or channels is then directed over a surface of the blade. The air exiting the outlet may then contribute to the cooling of the blade through film cooling.
  • Figure 3 illustrates a portion of a wall 40 of a blade of a rotor according to an embodiment of the invention. It is to be realised that the blade of the rotor has a variable thickness. As illustrated in Figure 3a, the wall 40 comprises a first exterior wall 42 and a second exterior wall 44. In this embodiment, both exterior walls 42 and 44 form outer surfaces for the blade.
  • a partitioning wall 48 is formed between the first exterior wall 42 and the second exterior wall 44 between the first exterior wall 42 and the second exterior wall 44.
  • the partitioning wall 48 is formed as an undulating form which contacts both the interior and the exterior walls. Therefore, the partitioning wall 48 divides the interior of the blade into a number of chambers 46.
  • the undulating form of the partitioning wall 48 further provides the blade with structural support as the partitioning wall acts against deformation of the blade during use.
  • the undulation of partitioning wall 48 forms a repeating pattern. This repeating pattern has a length ⁇ .
  • the thickness of the blade illustrated here is denoted as T.
  • the ratio of t:l is between 4 and 4.5. It has been found that this ration provides effective cooling.
  • embodiments of the invention find particular application in microturbines.
  • Such turbines may be characterised by the thickness of the turbine blade. Therefore, in an embodiment, the thickness of the blade illustrated in Figure 3 (i.e. dimension T) is 2 mm. In a further embodiment, 1 mm ⁇ I ⁇ 2 mm. In a further embodiment, I ⁇ 1 mm.
  • manufacturing techniques utilising additive layers such as sintering, and laser sintering in particular are particularly well suited to forming the intricate structures such as the inner chambers, the outlet and inlet channels and any surface cooling structures which may additionally be incorporated (such as ribs, dimples, pins etc.).
  • Figure 3b illustrates a portion of the wall 40 illustrated in Figure 3a.
  • the undulating partitioning wall 48 includes a flow channel 50 formed therein.
  • Flow channel 50 allows air (in this embodiment) to flow between adjacent chambers.
  • each of the chambers is paced in fluid communication with one or more adjacent chambers by a corresponding flow channel.
  • the chambers are also thereby placed in fluid communication with an inlet channel and an outlet channel (as described above, both formed in an outer surface, i.e. exterior wall 42 of the blade) thereby allowing movement of air from outside the turbine blade, through each of the channels and out of the outlet channel, thereby cooling the blade during use.
  • the flow channel has a linear dimension D.
  • the flow channel is cylindrical and therefore the linear dimension D is the diameter of the flow channel.
  • the flow channel illustrated communicates with the chamber 46 illustrated in Figure 3b which is triangular in shape ( Figure 3a).
  • the distance between the flow channel 50 and the opposite wall is S.
  • the ratio S:D is about 3:1. In alternate embodiments, this ratio is between 2:1 and 4:1 .
  • the ratio of S:D is chosen depending on the functional requirements of the blade and the effect that the impingement cooling of the cooling fluid passing through the flow channel has. It has been found that a ratio of about 3:1 provides effective cooling.
  • the undulating form of the partitioning wall 40 increases the path length for the air flowing through the blade.
  • the flow channels 50 are positioned to encourage impingement cooling.
  • a hollow blade i.e. having one of more internal chambers
  • a process of sintering is particularly well suited to a blade which has multiple inlets. This too allows for the construction of an inner cooled blade having more than one chamber, and multiple inlet channels connected to multiple channels.
  • FIG 4 illustrates a blade 60 for a gas turbine according to a further embodiment of the invention.
  • the blade 60 is hollow and is formed with a plurality of chambers 62.
  • Each of the chambers 62 includes a baffle 64 and an inlet channel 66.
  • each of the chambers is also formed with an outlet channel.
  • the inlet channels 66 allow the ingress of air which enters the corresponding chamber.
  • the baffles 64 are situated in the chambers so as to increase the path length over which the air must travel between the inlet channel and the outlet channel. This increases the cooling properties of the arrangement.
  • the blade 60 illustrated in Figure 4 is manufactured through a process of sintering.
  • the entire blade, or a relatively small portion thereof, have been manufactured by sintering.
  • different portions of a rotor are produced using different sintering techniques.
  • One of the advantages of sintering is that the process may be applied to different materials. It is therefore relatively easy to manufacture different parts of the same blade from different materials depending on the desired characteristics.
  • the blade 20 as illustrated as divided up into three portions 90, 92 and 94 by dashed lines 82 and 84.
  • the different portions of a blade are sintered using different materials.
  • the materials of the different portions are chosen according to their characteristics. It is to be realised that portion 22 faces the gas inlet of the turbine and therefore experiences the highest temperatures. Therefore, a material able to withstand high temperatures is chosen for the sintering of portion 90 whereas those used for portions 92 and 94 may have slightly less favourable thermal response characteristics, but may be cheaper, or have other desirable characteristics encouraging their mix.
  • the inlet channels may be formed in only selected portions of the blade. Where different portions are sintered from different materials, the inlet channels are preferably formed in the portion or portions made from the material best suited to withstand the corrosive action of the air or other cooling fluid entering through the inlet channel. In an embodiment, this is the portion 90 sintered from Titanium.
  • FIG. 5 illustrates a process 100 for manufacturing a turbine rotor according to an embodiment of the invention.
  • the method 100 starts at step 102 where a computer software program is prepared.
  • the laser sintering is carried out by a computer aided manufacturing (CAM) machine which is controlled by a computer program. Therefore, the computer program will specify all of the design parameters of the rotor to be manufactured using this method.
  • CAM computer aided manufacturing
  • step 104 the raw materials are loaded into the laser sintering machine.
  • the machine is capable of switching between materials and therefore three different materials are loaded at once.
  • step 106 the laser is started and the method proceeds to step 108.
  • the material for the first region (corresponding to the first portion 90 of Figure 2) is deposited according to the pre-defined computer program.
  • the material for the mid-portion (portion 92) and the third portion (portion 94) are deposited in steps 1 10 and 1 12. Once the third portion has been completed, the process terminates at step 1 14.
  • the third portion is deposited first, then the mid- portion, and finally the first portion, with reference to Figure 1.
  • the rotors may be manufactured starting with the portion of the rotor having the greater diameter and progressing to portions of the rotor having smaller diameters.
  • the material used for the sintering is changed so that different regions are comprised of different materials.
  • a process of laser sintering has been described, it is to be realised that any process involving the accumulation of layers by amalgamating powdered building material under the influence of an energy source may be used.
  • electric-arc sintering may be used instead.
  • any metal or other material which may be provided in powder form and re-combined may be used as the material for the manufacture of blades and rotors according to embodiments of the invention.
  • Other materials which may be used include brass and ceramics.
  • other forms of additive manufacturing whereby the material of the construction is laid down in layers may be used.
  • the sintering process may be directed so as to form structures which encourage the cooling characteristics of the turbine blades.
  • the number of chambers and inlets may be varied.
  • the number of inlets may depend on the number of chambers to ensure that all of the chambers cool in a manner dependant on their operating characteristics (e.g. the temperature and mechanical strain/stress they are required to withstand in relation to the material from which they are composed).

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne une aube de turbine, l'aube comprenant une ou plusieurs chambres à des fins de refroidissement par l'introduction d'un fluide de refroidissement dans la chambre, l'aube étant construite par un procédé selon lequel des couches successives sont ajoutées l'une sur l'autre, comme le frittage. L'invention concerne en outre un rotor incorporant une telle aube et un procédé correspondant de fabrication.
PCT/EP2013/078146 2012-12-28 2013-12-30 Aubes de turbine et procédé de construction WO2014102381A1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
LU92125 2012-12-28
LU92125A LU92125B1 (en) 2012-12-28 2012-12-28 Turbine blades and method of construction

Publications (1)

Publication Number Publication Date
WO2014102381A1 true WO2014102381A1 (fr) 2014-07-03

Family

ID=47470066

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/EP2013/078146 WO2014102381A1 (fr) 2012-12-28 2013-12-30 Aubes de turbine et procédé de construction

Country Status (2)

Country Link
LU (1) LU92125B1 (fr)
WO (1) WO2014102381A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102018217056A1 (de) * 2018-10-05 2020-04-09 Continental Automotive Gmbh Turbinenläufer und Verfahren zur Herstellung desselben
US20230012375A1 (en) * 2021-07-09 2023-01-12 Raytheon Technologies Corporation Radial flow turbine rotor with internal fluid cooling

Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB602530A (en) * 1945-10-16 1948-05-28 Bristol Aeroplane Co Ltd Improvements in or relating to gas turbines
DE102009004881A1 (de) * 2009-01-16 2010-07-29 Bosch Mahle Turbo Systems Gmbh & Co. Kg Ladeeinrichtung
EP2230384A2 (fr) * 2009-03-18 2010-09-22 General Electric Company Dispositif d'augmentation du refroidissement à film et aube de turbine incorporant ce dispositif
WO2010137609A1 (fr) * 2009-05-26 2010-12-02 株式会社Ihi Hélice appliquée à un compresseur et son procédé de fabrication
GB2471119A (en) * 2009-06-17 2010-12-22 Nebb Technology As Sintered gas turbine blade
US20110038709A1 (en) * 2009-08-13 2011-02-17 George Liang Turbine Vane for a Gas Turbine Engine Having Serpentine Cooling Channels
EP2394974A1 (fr) * 2009-02-09 2011-12-14 IHI Corporation Procédé de fabrication de céramique frittée de sialon
EP2423434A2 (fr) * 2010-08-31 2012-02-29 General Electric Company Préforme de forgeage de rotor compact d'alimentation et rotor de turbine compact à poudre forgé et leurs procédés de fabrication
US20120067054A1 (en) * 2010-09-21 2012-03-22 Palmer Labs, Llc High efficiency power production methods, assemblies, and systems
US20120082563A1 (en) * 2010-09-30 2012-04-05 Florida Turbine Technologies, Inc. Cooed IBR for a micro-turbine

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB0709838D0 (en) * 2007-05-23 2007-07-04 Rolls Royce Plc A hollow blade and a method of manufacturing a hollow blade

Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB602530A (en) * 1945-10-16 1948-05-28 Bristol Aeroplane Co Ltd Improvements in or relating to gas turbines
DE102009004881A1 (de) * 2009-01-16 2010-07-29 Bosch Mahle Turbo Systems Gmbh & Co. Kg Ladeeinrichtung
EP2394974A1 (fr) * 2009-02-09 2011-12-14 IHI Corporation Procédé de fabrication de céramique frittée de sialon
EP2230384A2 (fr) * 2009-03-18 2010-09-22 General Electric Company Dispositif d'augmentation du refroidissement à film et aube de turbine incorporant ce dispositif
WO2010137609A1 (fr) * 2009-05-26 2010-12-02 株式会社Ihi Hélice appliquée à un compresseur et son procédé de fabrication
GB2471119A (en) * 2009-06-17 2010-12-22 Nebb Technology As Sintered gas turbine blade
US20110038709A1 (en) * 2009-08-13 2011-02-17 George Liang Turbine Vane for a Gas Turbine Engine Having Serpentine Cooling Channels
EP2423434A2 (fr) * 2010-08-31 2012-02-29 General Electric Company Préforme de forgeage de rotor compact d'alimentation et rotor de turbine compact à poudre forgé et leurs procédés de fabrication
US20120067054A1 (en) * 2010-09-21 2012-03-22 Palmer Labs, Llc High efficiency power production methods, assemblies, and systems
US20120082563A1 (en) * 2010-09-30 2012-04-05 Florida Turbine Technologies, Inc. Cooed IBR for a micro-turbine

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102018217056A1 (de) * 2018-10-05 2020-04-09 Continental Automotive Gmbh Turbinenläufer und Verfahren zur Herstellung desselben
DE102018217056B4 (de) 2018-10-05 2022-09-29 Vitesco Technologies GmbH Turbinenläufer und Verfahren zur Herstellung desselben
US20230012375A1 (en) * 2021-07-09 2023-01-12 Raytheon Technologies Corporation Radial flow turbine rotor with internal fluid cooling

Also Published As

Publication number Publication date
LU92125B1 (en) 2014-06-30

Similar Documents

Publication Publication Date Title
US11951532B2 (en) Additively manufactured core
JP6928995B2 (ja) 翼のためのテーパした冷却チャネル
EP1998004B1 (fr) Composant de turbine avec canaux de refroidissement pourvus des "microcircuits", décalés en direction axiale, ayant un écoulement radial
US8066483B1 (en) Turbine airfoil with non-parallel pin fins
US7625180B1 (en) Turbine blade with near-wall multi-metering and diffusion cooling circuit
EP2911815B1 (fr) Noyau de coulée pour un arrangement de refroidissement pour un élément de turbine à gaz
US8951004B2 (en) Cooling arrangement for a gas turbine component
US8303253B1 (en) Turbine airfoil with near-wall mini serpentine cooling channels
EP1533475A2 (fr) Elément chaud avec des dispositifs de refroidissement
CN105715306B (zh) 用于制造燃气涡轮发动机零件的增材制造的铸造制品
US9995150B2 (en) Cooling configuration for a gas turbine engine airfoil
JP2016505754A (ja) ターボ機械ブレード、対応するターボ機械、およびタービンブレードを製造する方法
EP3090145B1 (fr) Turbulateur de passage de refroidissement de composant de turbine à gaz
EP3063389B1 (fr) Socles de distribution de films refroidis par un trou
US20190316472A1 (en) Double wall airfoil cooling configuration for gas turbine engine
JP7458742B2 (ja) 複数の壁および内部遮熱コーティングを備えるタービン翼形部
EP3372787B1 (fr) Aube de turbine, turbine à gaz et procédé de fabrication d'aube de turbine
JP5868609B2 (ja) ガスタービン動翼及びその製造方法
JP2017115861A (ja) 物品及び物品を形成する方法
JP2014114814A (ja) タービンから熱を除去するためのシステムおよび方法
WO2014102381A1 (fr) Aubes de turbine et procédé de construction
EP3287598B1 (fr) Éléments de refroidissement à géométrie de chevrons en trois dimensions
KR20220008914A (ko) 에어포일용 니어 월 리딩 에지 냉각 채널
US8382431B1 (en) Turbine rotor blade
EP1533481A2 (fr) Composant pour guider des gaz chauds avec une structure de refroidissement réticulée comprenant des bosses

Legal Events

Date Code Title Description
121 Ep: the epo has been informed by wipo that ep was designated in this application

Ref document number: 13818235

Country of ref document: EP

Kind code of ref document: A1

NENP Non-entry into the national phase

Ref country code: DE

32PN Ep: public notification in the ep bulletin as address of the adressee cannot be established

Free format text: NOTING OF LOSS OF RIGHTS PURSUANT TO RULE 112(1) EPC (EPO FORM 1205A DATED 23.10.2015)

122 Ep: pct application non-entry in european phase

Ref document number: 13818235

Country of ref document: EP

Kind code of ref document: A1