US20140030109A1 - low-Modulus Gas-Turbine Compressor Blade - Google Patents

low-Modulus Gas-Turbine Compressor Blade Download PDF

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Publication number
US20140030109A1
US20140030109A1 US13/953,934 US201313953934A US2014030109A1 US 20140030109 A1 US20140030109 A1 US 20140030109A1 US 201313953934 A US201313953934 A US 201313953934A US 2014030109 A1 US2014030109 A1 US 2014030109A1
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Prior art keywords
compressor blade
compressor
blade
accordance
modulus
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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.)
Abandoned
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US13/953,934
Inventor
Karl Schreiber
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.)
Rolls Royce Deutschland Ltd and Co KG
Rolls Royce PLC
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Rolls Royce Deutschland Ltd and Co KG
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Assigned to ROLLS-ROYCE DEUTSCHLAND LTD & CO KG reassignment ROLLS-ROYCE DEUTSCHLAND LTD & CO KG ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: SCHREIBER, KARL
Publication of US20140030109A1 publication Critical patent/US20140030109A1/en
Assigned to ROLLS-ROYCE DEUTSCHLAND LTD & CO KG reassignment ROLLS-ROYCE DEUTSCHLAND LTD & CO KG ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: WEBSTER, JOHN RICHARD
Assigned to ROLLS-ROYCE PLC, ROLLS-ROYCE DEUTSCHLAND LTD & CO KG reassignment ROLLS-ROYCE PLC ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: ROLLS-ROYCE DEUTSCHLAND LTD & CO KG
Abandoned legal-status Critical Current

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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
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow 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
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • 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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/26—Rotors specially for elastic fluids
    • F04D29/32—Rotors specially for elastic fluids for axial flow pumps
    • F04D29/321—Rotors specially for elastic fluids for axial flow pumps for axial flow compressors
    • F04D29/324—Blades
    • 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
    • F05D2240/00—Components
    • F05D2240/20—Rotors
    • F05D2240/30—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor
    • F05D2240/303—Characteristics of rotor blades, i.e. of any element transforming dynamic fluid energy to or from rotational energy and being attached to a rotor related to the leading edge of a rotor blade

Definitions

  • This invention relates to a gas-turbine compressor blade, in particular to an aircraft gas turbine or a stationary turbine, having an airfoil fastened to a blade root.
  • the compressor blade has a leading edge or inflow edge, which is also called the blade front edge.
  • the erosion-resistant materials known from the state of the art are unsuitable as materials for compressor blades. Erosion-resistant coatings too have proven in practice to be unsuitable, as they are only effective for very small particle sizes.
  • the object underlying the present invention is to provide an aircraft gas-turbine compressor blade which, while being simply designed and easily and cost-effectively producible, has a high erosion resistance.
  • At least the area of the leading edge of the compressor blade is made from a material which has a very low modulus of elasticity.
  • a suitable and low modulus of elasticity is in the range of 50 GPa.
  • the compressor blade in accordance with the invention thus has, due to the low modulus of elasticity, a relatively high elasticity in the area of the leading edge, so that an impact of a foreign object, for example a sand particle or stone, does not lead to damage thanks to the elasticity of the material in the area of the leading edge.
  • a foreign object for example a sand particle or stone
  • zero or only very low erosion occurs when compared with compressor blades known from the state of the art.
  • At least the area of the leading edge is made from a titanium alloy of types Ti-28Nb-1Fe-0.5Si, Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn.
  • a titanium alloy of this type it is for example possible to reduce erosive wear by about 50% compared with compressor blades made from a conventional alloy, for example Ti64.
  • erosive wear can be reduced by 30%.
  • leading edge is designed in the form of a leading-edge element which is connected to the airfoil of the compressor blade by means of a welding process.
  • a welding process is particularly advantageous.
  • the airfoil itself can, in accordance with the invention, be made from one of the conventional titanium alloys, for example Ti64, Ti6246 or Ti6242.
  • leading-edge element extends preferably up to just in front of the blade root, and in the direct transition of the blade root it is not necessary to apply or use the alloy in accordance with the invention.
  • the compressor blade in accordance with the invention which can be designed as a single compressor blade, as a blisk blade or as a fan blade, is thus to a high degree resistant against erosion and damage from foreign objects.
  • the invention thus also relates to a method for the manufacture of a compressor blade in which a leading-edge element is manufactured separately and is connected to the airfoil by means of a welding process, in particular by a laser welding process, where the leading-edge element is made from a low-modulus (in respect of the modulus of elasticity) titanium alloy, for example from Ti-28Nb-1Fe-0.5Si, from Ti-36Nb-2Ta-3Zr-0.3O or from Ti-24Nb-4Zr-8Sn and/or an alloy with a modulus of elasticity of substantially 50 to 70 GPa and where the airfoil is made from an alloy known from the state of the art.
  • a low-modulus (in respect of the modulus of elasticity) titanium alloy for example from Ti-28Nb-1Fe-0.5Si, from Ti-36Nb-2Ta-3Zr-0.3O or from Ti-24Nb-4Zr-8Sn and/or an alloy with a modulus of elasticity of substantially 50 to 70 GPa and
  • the invention also relates to the use of a low-modulus (of elasticity) titanium alloy, for example Ti-28Nb-1Fe-0.5Si or Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn, in particular for the leading-edge area of a compressor blade.
  • a low-modulus (of elasticity) titanium alloy for example Ti-28Nb-1Fe-0.5Si or Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn, in particular for the leading-edge area of a compressor blade.
  • FIG. 1 shows a schematized representation of a gas-turbine engine in accordance with the present invention
  • FIG. 2 shows a schematized side view of a compressor blade in accordance with the present invention
  • FIG. 3 shows a representation of the curve of the modulus of elasticity as a function of the blade length.
  • the gas-turbine engine 10 in accordance with FIG. 1 is an example of a turbomachine where the invention can be used. The following however makes clear that the invention can also be used in other turbomachines.
  • the engine 10 is of conventional design and includes in the flow direction, one behind the other, an air inlet 11 , a fan 12 rotating inside a casing, an intermediate-pressure compressor 13 , a high-pressure compressor 14 , combustion chambers 15 , a high-pressure turbine 16 , an intermediate-pressure turbine 17 and a low-pressure turbine 18 as well as an exhaust nozzle 19 , all of which being arranged about a central engine axis 1 .
  • the intermediate-pressure compressor 13 and the high-pressure compressor 14 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 20 , generally referred to as stator vanes and projecting radially inwards from the engine casing 21 in an annular flow duct through the compressors 13 , 14 .
  • the compressors furthermore have an arrangement of compressor rotor blades 22 which project radially outwards from a rotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine 16 or the intermediate-pressure turbine 17 , respectively.
  • the turbine sections 16 , 17 , 18 have similar stages, including an arrangement of fixed stator vanes 23 projecting radially inwards from the casing 21 into the annular flow duct through the turbines 16 , 17 , 18 , and a subsequent arrangement of turbine blades 24 projecting outwards from a rotatable hub 27 .
  • the compressor drum or compressor disk 26 and the blades 22 arranged thereon, as well as the turbine rotor hub 27 and the turbine rotor blades 24 arranged thereon rotate about the engine axis 1 during operation.
  • FIG. 2 shows a schematic side view of a compressor blade in accordance with the invention, which can be a conventional compressor blade, a fan blade or a blisk blade.
  • the blade has for example a blade root 31 to which is connected an airfoil 29 designed in the usual way.
  • a leading-edge element 30 is connected by means of a laser weld 32 in the area of a leading edge or inflow edge.
  • the leading-edge element 30 extends up to just in front of the blade root 31 , and does not need to be connected directly to the blade root 31 .
  • FIG. 3 shows a representation of the curve of the modulus of elasticity in schematic form. It can be seen that the modulus of elasticity of the leading-edge element 30 is lower than the modulus of elasticity of the airfoil 29 , where a soft transition is obtained in the area of the weld 32 between the lower and the higher moduli of elasticity.

Abstract

The present invention relates to an aircraft gas-turbine compressor blade having an airfoil with a leading edge, with at least the area of the leading edge being made from a low-modulus titanium alloy.

Description

  • This invention relates to a gas-turbine compressor blade, in particular to an aircraft gas turbine or a stationary turbine, having an airfoil fastened to a blade root. The compressor blade has a leading edge or inflow edge, which is also called the blade front edge.
  • Aircraft gas turbines always face the problem that the compressor blades are subjected to heavy erosion due to sucked-in particles. These are for example grains of sand or the like which impact the compressor blades at high velocity.
  • As a result of the erosion occurring, it is necessary to replace or repair the compressor blades. Repair is, in the case of blisks in particular, very complex from the engineering viewpoint and also very cost-intensive.
  • The erosion-resistant materials known from the state of the art are unsuitable as materials for compressor blades. Erosion-resistant coatings too have proven in practice to be unsuitable, as they are only effective for very small particle sizes.
  • The object underlying the present invention is to provide an aircraft gas-turbine compressor blade which, while being simply designed and easily and cost-effectively producible, has a high erosion resistance.
  • It is a particular object of the present invention to provide solution to the above problematics by a combination of the features of Claim 1. Further advantageous embodiments of the invention become apparent from the sub-claims.
  • It is provided in accordance with the invention that at least the area of the leading edge of the compressor blade is made from a material which has a very low modulus of elasticity. A suitable and low modulus of elasticity is in the range of 50 GPa.
  • The compressor blade in accordance with the invention thus has, due to the low modulus of elasticity, a relatively high elasticity in the area of the leading edge, so that an impact of a foreign object, for example a sand particle or stone, does not lead to damage thanks to the elasticity of the material in the area of the leading edge. In particular, zero or only very low erosion occurs when compared with compressor blades known from the state of the art.
  • In accordance with the invention, it is particularly advantageous when at least the area of the leading edge is made from a titanium alloy of types Ti-28Nb-1Fe-0.5Si, Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn. By means of an alloy of this type, it is for example possible to reduce erosive wear by about 50% compared with compressor blades made from a conventional alloy, for example Ti64. Compared with compressor blades known from the state of the art made from nickel-based alloys, for example IN718, erosive wear can be reduced by 30%.
  • It is possible in accordance with the invention to either manufacture the entire compressor blade from the above titanium alloy with the low modulus of elasticity or only to manufacture parts of the compressor blade from this material.
  • In a particularly favourable embodiment of the invention, it is provided that the leading edge is designed in the form of a leading-edge element which is connected to the airfoil of the compressor blade by means of a welding process. Particularly advantageous is the use of a laser welding process.
  • The airfoil itself can, in accordance with the invention, be made from one of the conventional titanium alloys, for example Ti64, Ti6246 or Ti6242.
  • If a separate leading-edge element is provided, the latter extends preferably up to just in front of the blade root, and in the direct transition of the blade root it is not necessary to apply or use the alloy in accordance with the invention.
  • In the case of separate manufacture of the leading-edge element in accordance with the invention and connection by means of a laser welding process, there is the advantage that due to the thermal influence during the laser welding operation a soft transition is obtained between the low modulus of elasticity of the leading-edge element and the higher modulus of elasticity of the material in the rest of the airfoil. The result is thus a gradual transition of the elastic properties, which acts advantageously on the behaviour of the entire compressor blade component. This effect is based on the fact that during heating of the low-modulus titanium alloy in temperature ranges above around 500° C., the modulus of elasticity rises again to values of conventional titanium alloys, and hence the direct area of the weld is homogeneous.
  • The compressor blade in accordance with the invention, which can be designed as a single compressor blade, as a blisk blade or as a fan blade, is thus to a high degree resistant against erosion and damage from foreign objects.
  • The invention thus also relates to a method for the manufacture of a compressor blade in which a leading-edge element is manufactured separately and is connected to the airfoil by means of a welding process, in particular by a laser welding process, where the leading-edge element is made from a low-modulus (in respect of the modulus of elasticity) titanium alloy, for example from Ti-28Nb-1Fe-0.5Si, from Ti-36Nb-2Ta-3Zr-0.3O or from Ti-24Nb-4Zr-8Sn and/or an alloy with a modulus of elasticity of substantially 50 to 70 GPa and where the airfoil is made from an alloy known from the state of the art.
  • The invention also relates to the use of a low-modulus (of elasticity) titanium alloy, for example Ti-28Nb-1Fe-0.5Si or Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn, in particular for the leading-edge area of a compressor blade.
  • The present invention is described in the following in light of the accompanying drawing, showing an exemplary embodiment. In the drawing,
  • FIG. 1 shows a schematized representation of a gas-turbine engine in accordance with the present invention,
  • FIG. 2 shows a schematized side view of a compressor blade in accordance with the present invention, and
  • FIG. 3 shows a representation of the curve of the modulus of elasticity as a function of the blade length.
    •
  • The gas-turbine engine 10 in accordance with FIG. 1 is an example of a turbomachine where the invention can be used. The following however makes clear that the invention can also be used in other turbomachines. The engine 10 is of conventional design and includes in the flow direction, one behind the other, an air inlet 11, a fan 12 rotating inside a casing, an intermediate-pressure compressor 13, a high-pressure compressor 14, combustion chambers 15, a high-pressure turbine 16, an intermediate-pressure turbine 17 and a low-pressure turbine 18 as well as an exhaust nozzle 19, all of which being arranged about a central engine axis 1.
  • The intermediate-pressure compressor 13 and the high-pressure compressor 14 each include several stages, of which each has an arrangement extending in the circumferential direction of fixed and stationary guide vanes 20, generally referred to as stator vanes and projecting radially inwards from the engine casing 21 in an annular flow duct through the compressors 13, 14. The compressors furthermore have an arrangement of compressor rotor blades 22 which project radially outwards from a rotatable drum or disk 26 linked to hubs 27 of the high-pressure turbine 16 or the intermediate-pressure turbine 17, respectively.
  • The turbine sections 16, 17, 18 have similar stages, including an arrangement of fixed stator vanes 23 projecting radially inwards from the casing 21 into the annular flow duct through the turbines 16, 17, 18, and a subsequent arrangement of turbine blades 24 projecting outwards from a rotatable hub 27. The compressor drum or compressor disk 26 and the blades 22 arranged thereon, as well as the turbine rotor hub 27 and the turbine rotor blades 24 arranged thereon rotate about the engine axis 1 during operation.
  • FIG. 2 shows a schematic side view of a compressor blade in accordance with the invention, which can be a conventional compressor blade, a fan blade or a blisk blade. The blade has for example a blade root 31 to which is connected an airfoil 29 designed in the usual way. A leading-edge element 30 is connected by means of a laser weld 32 in the area of a leading edge or inflow edge. The leading-edge element 30 extends up to just in front of the blade root 31, and does not need to be connected directly to the blade root 31.
  • FIG. 3 shows a representation of the curve of the modulus of elasticity in schematic form. It can be seen that the modulus of elasticity of the leading-edge element 30 is lower than the modulus of elasticity of the airfoil 29, where a soft transition is obtained in the area of the weld 32 between the lower and the higher moduli of elasticity.
    • LIST OF REFERENCE NUMERALS
    • 1 Engine axis
    • 10 Gas-turbine engine/core engine
    • 11 Air inlet
    • 12 Fan
    • 13 Intermediate-pressure compressor (compressor)
    • 14 High-pressure compressor
    • 15 Combustion chambers
    • 16 High-pressure turbine
    • 17 Intermediate-pressure turbine
    • 18 Low-pressure turbine
    • 19 Exhaust nozzle
    • 20 Guide vanes
    • 21 Engine casing
    • 22 Compressor rotor blades
    • 23 Stator vanes
    • 24 Turbine blades
    • 26 Compressor drum or disk
    • 27 Turbine rotor hub
    • 28 Exhaust cone
    • 29 Airfoil
    • 30 Leading-edge element
    • 31 Blade root
    • 32 Weld
    • 33 Leading edge

Claims (9)

1. Gas-turbine compressor blade having an airfoil with a leading edge, with at least the area of the leading edge being made from a low-modulus titanium alloy.
2. Compressor blade in accordance with claim 1, characterized in that at least the area of the leading edge is made from Ti-28Nb-1Fe-0.5Si or Ti-36Nb-2Ta-3Zr-0.3O or Ti-24Nb-4Zr-8Sn.
3. Compressor blade in accordance with claim 1, characterized in that the leading edge is designed in the form of a leading-edge element which is connected to the airfoil by means of a welding process.
4. Compressor blade in accordance with claim 3, characterized in that the leading-edge element is connected to the airfoil by means of a laser welding process.
5. Compressor blade in accordance with claim 1, characterized in that the area of the leading edge has a modulus of elasticity of substantially 50 to 70 GPa.
6. Compressor blade in accordance with claim 1, characterized in that the airfoil is made from Ti64 or Ti6246 or Ti6242 or IN718.
7. Compressor blade in accordance with claim 1, characterized in that the compressor blade is designed as a conventional compressor blade.
8. Compressor blade in accordance with claim 1, characterized in that the compressor blade is designed as a fan blade.
9. Compressor blade in accordance with claim 1, characterized in that the compressor blade is designed as a blisk blade.
US13/953,934 2012-07-30 2013-07-30 low-Modulus Gas-Turbine Compressor Blade Abandoned US20140030109A1 (en)

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DE102012015137.3 2012-07-30
DE102012015137.3A DE102012015137A1 (en) 2012-07-30 2012-07-30 Low-modulus gas turbine compressor blade

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EP (1) EP2692989A3 (en)
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US20160243979A1 (en) * 2015-02-24 2016-08-25 GM Global Technology Operations LLC Headlight for a motor vehicle
US20180001814A1 (en) * 2016-07-01 2018-01-04 Ford Global Technologies, Llc Vehicle headlamp alignment system and method
US10047924B2 (en) 2014-11-03 2018-08-14 Zkw Group Gmbh Light system for a motor vehicle

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US5509933A (en) * 1989-12-21 1996-04-23 Smith & Nephew Richards, Inc. Medical implants of hot worked, high strength, biocompatible, low modulus titanium alloys
US5320487A (en) * 1993-01-19 1994-06-14 General Electric Company Spring clip made of a directionally solidified material for use in a gas turbine engine
US5725354A (en) * 1996-11-22 1998-03-10 General Electric Company Forward swept fan blade
US6073439A (en) * 1997-03-05 2000-06-13 Rolls-Royce Plc Ducted fan gas turbine engine
US6200685B1 (en) * 1997-03-27 2001-03-13 James A. Davidson Titanium molybdenum hafnium alloy
US7156622B2 (en) * 2003-02-22 2007-01-02 Rolls-Royce Deutschland Ltd & Co Kg Compressor blade for an aircraft engine
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US20080152506A1 (en) * 2006-12-21 2008-06-26 Karl Schreiber Fan blade for a gas-turbine engine
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Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10047924B2 (en) 2014-11-03 2018-08-14 Zkw Group Gmbh Light system for a motor vehicle
US20160243979A1 (en) * 2015-02-24 2016-08-25 GM Global Technology Operations LLC Headlight for a motor vehicle
US20180001814A1 (en) * 2016-07-01 2018-01-04 Ford Global Technologies, Llc Vehicle headlamp alignment system and method

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DE102012015137A1 (en) 2014-02-13
EP2692989A3 (en) 2017-06-14
EP2692989A2 (en) 2014-02-05

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