WO2014065968A1 - Composite blade with uni-tape airfoil spars - Google Patents

Composite blade with uni-tape airfoil spars Download PDF

Info

Publication number
WO2014065968A1
WO2014065968A1 PCT/US2013/061294 US2013061294W WO2014065968A1 WO 2014065968 A1 WO2014065968 A1 WO 2014065968A1 US 2013061294 W US2013061294 W US 2013061294W WO 2014065968 A1 WO2014065968 A1 WO 2014065968A1
Authority
WO
WIPO (PCT)
Prior art keywords
blade
airfoil
spars
core section
chordwise
Prior art date
Application number
PCT/US2013/061294
Other languages
French (fr)
Inventor
Nicholas Joseph Kray
Ian Francis Prentice
Tod Winton DAVIS
Dong-Jin Shim
Pranav Dhoj SHAH
Original Assignee
General Electric Company
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 General Electric Company filed Critical General Electric Company
Priority to EP13774557.6A priority Critical patent/EP2917496A1/en
Priority to BR112015009060A priority patent/BR112015009060A2/en
Priority to CN201380054437.7A priority patent/CN104981586A/en
Priority to CA2888777A priority patent/CA2888777A1/en
Priority to JP2015537713A priority patent/JP6179961B2/en
Publication of WO2014065968A1 publication Critical patent/WO2014065968A1/en

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
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/282—Selecting composite materials, e.g. blades with reinforcing filaments
    • 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
    • F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
    • F04D—NON-POSITIVE-DISPLACEMENT PUMPS
    • F04D29/00—Details, component parts, or accessories
    • F04D29/02—Selection of particular materials
    • F04D29/023—Selection of particular materials especially adapted for elastic fluid pumps
    • 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
    • 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
    • F01D5/16—Form or construction for counteracting blade vibration
    • 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
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
    • 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
    • F05D2240/305—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 pressure side of a rotor blade
    • 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
    • F05D2240/306—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 suction side of a rotor blade
    • 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
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • F05D2260/941—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF] particularly aimed at mechanical or thermal stress reduction
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/603—Composites; e.g. fibre-reinforced
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/60—Properties or characteristics given to material by treatment or manufacturing
    • F05D2300/603—Composites; e.g. fibre-reinforced
    • F05D2300/6034—Orientation of fibres, weaving, ply angle
    • 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

  • the invention relates to gas turbine engine blades and, particularly, to composite blades.
  • composited in a light-weight matrix have been developed for aircraft gas turbine engines.
  • the blades are light-weight having high strength.
  • composite has come to be defined as a material containing a reinforcement such as fibers or particles supported in a binder or matrix
  • the composites used for the blades disclosed herein are made of a unidirectional tape material and an epoxy resin matrix. A discussion of this and other suitable materials may be found in the "Engineering Materials
  • the composite blades disclosed herein are made from the non-metallic type made of a material containing a fiber such as a carbonaceous, silica, metal, metal oxide, or ceramic fiber embedded in a resin material such as Epoxy, PMR15, BMI, PEEU, etc.
  • the fibers are unidirectionally aligned in a tape that is impregnated with a resin, formed into a part shape, and cured via an autoclaving process or press molding to form a light weight, stiff, relatively homogeneous article having laminates or plies within.
  • Composite fan blades have been developed for aircraft gas turbine engines to reduce weight and cost, particularly for fan blades in larger engines.
  • a large engine composite wide chord fan blades offer a significant weight savings over a large engine having standard chorded fan blades.
  • all gas turbine engine blades face resonance or flexural modes.
  • a gas turbine engine composite fan blade (10) includes an airfoil (12) having pressure and suction sides (41, 43) extending outwardly in a spanwise direction (S) from a blade root (20) of the blade (10) along a span (S) to a blade tip (47) .
  • a core section (50) of the blade (10) includes composite quasi-isotropic plies (52) extending spanwise outwardly through the blade (10) including the root (20) and the airfoil (12) towards the tip (47) .
  • One or more spars (54, 56) including a stack (62) of uni-tape plies (63) having a preferential 0 degree fiber orientation with respect to the span (S) spanwise outwardly through the root
  • the chordwise extending portion (58) of the core section (50) may be centered about a maximum thickness location (61) of the airfoil (12) .
  • the spars (54, 56) may have a spanwise height (H) , chordwise width (W) , and spar thickness (TS) that avoids flexural airfoil modes such as first and second flexural airfoil modes.
  • the one or more spars may include pressure and suction side spars (54, 56) sandwiching a chordwise extending portion (58) of the core section (50) in the airfoil (12) which may be located near or along the pressure and suction sides (41, 43)
  • the one or more spars include chordwise spaced apart upstream and downstream pressure side spars (74, 76) and chordwise spaced apart upstream and downstream suction side spars (78, 80)
  • FIG. 1 is a perspective view illustration of an aircraft gas turbine engine composite fan blade having a composite uni- tape spar.
  • FIG. 2 is a cross-sectional illustration of the
  • FIG. 3 is a perspective diagrammatical view
  • FIG. 4 is a perspective diagrammatical view
  • FIG. 5 is perspective diagrammatical view illustration of -P degree, 0 degree, and +P degree plies of the composite fan blade illustrated in FIG. 2.
  • FIG. 6 is a perspective view illustration of an
  • FIG. 7 is a cross-sectional illustration of the
  • FIGS. 1 and 2 Illustrated in FIGS. 1 and 2 is a composite fan blade 10 for a high bypass ratio fanjet gas turbine engine (not shown) having a composite airfoil 12.
  • Composite fan blade 10 is made up of filament reinforced laminations 30 formed from a composite material lay-up 36 of filament reinforced composite plies 40 (illustrated in FIG. 5) .
  • laminate and “ply” are synonymous.
  • the airfoil 12 includes pressure and suction sides 41, 43 extending outwardly in a spanwise direction from a fan blade root 20 along a span S to a blade tip 47.
  • the root 20 includes an integral dovetail 28 that enables the fan blade 10 to be mounted to a rotor disk.
  • the airfoil 12 extends along a chord C between chordwise spaced apart leading and trailing edges LE, TE . Thickness T of the airfoil 12 varies in both chordwise and spanwise directions C, S and extends between pressure and suction sides 41, 43 of the blade 10 also referred to as convex and concave sides of the blade or airfoil.
  • the airfoil 12 may be mounted on and be integral with a hub to form an integrally bladed rotor (IBR) or integrally with a disk in a BLISK
  • the plies 40 are generally all made from a
  • the plies 40 are essentially those plies that form the airfoil 12 and root 20 of the blade 10 as illustrated in FIGS. 1 and 3.
  • the composite fan blade 10 is made up of filament reinforced laminations 30 formed from a composite material lay-up 36 of different filament reinforced airfoil plies 40.
  • the blade 10 uses filament reinforced laminations or plies with a filament orientation of 0 degrees, +P degrees, and -P degrees as illustrated in FIG. 5.
  • the angle P is a
  • predetermined angle as measured from 0 degrees which corresponds to a generally radially extending axis of the airfoil which may be its centerline or stacking line and is typically about 45 degrees.
  • the composite fan blade 10 includes a core section 50 of composite quasi-isotropic plies 52.
  • Pressure and suction side spars 54, 56 sandwich a chordwise extending portion 58 of the core section 50 made of composite quasi-isotropic plies 52 generally near or along the pressure and suction sides 41, 43 respectively in the airfoil 12.
  • the chordwise extending portion 58 of the core section 50 extends chordwise partially through the airfoil 12.
  • the chordwise extending portion 58 of the core section 50 is generally centered chordwise in the airfoil
  • the exemplary embodiment of the chordwise extending portion 58 illustrated herein extends chordwise about 1/3 through the airfoil 12 and is generally centered chordwise about in the middle of the airfoil 12.
  • the composite quasi-isotropic ply chordwise extending portion 58 of the core section 50 is preferably limited to a thicker cross sectional area of the airfoil 12 around or centered about a maximum thickness Tmax location 61 of the airfoil 12, as illustrated in FIG. 2, so as to be most effective.
  • the Tmax location 61 is about a middle third of the airfoil between the leading and trailing edges LE, TE in the chordwise direction C for the exemplary airfoil illustrated herein.
  • the pressure and suction side spars 54, 56 are made from stacks 62 of preferential 0 degree uni-tape plies 63 (see FIG. 5) with a 0 degree fiber orientation with respect to the span S .
  • the pressure and suction side spars 54, 56 extends spanwise S through the fan blade root 20 and through a portion 53 of the airfoil 12 to a spar tip 57.
  • the pressure and suction side spars 54, 56 have a spanwise height H as measured from the fan blade root 20 to the spar tip 57 which is less than the span S of the airfoil.
  • the pressure and suction side spars 54, 56 extend all the way through the root 20 including the dovetail 28.
  • the quasi-isotropic ply core section 50 generally include alternating plies of tape with different +P, 0, and -P fiber orientations.
  • the pressure and suction side spars 54, 56 include uni-tape plies with a predominately 0 degree fiber orientation.
  • An exemplary blade ply lay-up is
  • the ply lay-up disclosed in 5,375,978 is referred to as a standard quasi-isotropic lay-up sequence of 0. degree, +45 degree, 0 degree, -45 degree fiber orientations with the plies having the numerous ply shapes.
  • the stacks 62 of the spars include uni-tape plies with a predominately 0 degree fiber orientation. A few of the plies may have another fiber orientation. An example is a stack having a total of 8 plies with 4 plies of 0 degree fiber orientation on both sides of two plies having +30 and a - 30 degree plies. This ply layup may be represented or denoted by 0 , 0 , 0 , 0 , +30 , -30 , 0 , 0 , 0 , 0.
  • the spars have a spanwise height H, chordwise width W, and spar thickness TS designed to increase radial or spanwise stiffness of the airfoil 12 without increasing the weight of the blade.
  • the spars are also designed or tailored or tuned to avoid flexural airfoil modes such as first and second flexural airfoil modes IF and 2F.
  • the spanwise height H and the spar thickness TS are designed or tailored to tuned or avoid flexural airfoil modes such as first and second flexural airfoil modes IF and 2F.
  • the exemplary embodiment of the composite blade illustrated herein is a fan blade but the composite blade with a quasi-isotropic ply core section and spars made from stacks 62 of 0 degree uni-tape plies 63 may also be used for other gas turbine engine blades such as compressor blades.
  • the exemplary embodiment of the composite blade 10 illustrated herein includes one or more outer cover plies 66 around the core section 50, made of composite
  • a leading edge metallic shield 68 is bonded around the leading edge LE .
  • the shield is often referred to as metallic cladding.
  • an alternative spar design for the composite fan blade 10 includes a core section 50 of composite quasi-isotropic plies and two sets of pressure and suction side spars.
  • the two sets include chordwise spaced apart upstream and downstream pressure side spars 74, 76 and chordwise spaced apart upstream and downstream suction side spars 78, 80 sandwiching the chordwise extending portion 58 of the core section 50 made of composite quasi-isotropic plies generally near or along the pressure and suction sides

Abstract

A gas turbine engine composite blade (10) includes an airfoil (12) having pressure and suction sides (41, 43) extending outwardly in a spanwise direction (S) from a blade root (20) along a span (S) to a blade tip (47). A core section (50) of the blade (10) including composite quasi-isotropic plies (52) extends spanwise outwardly through the blade (10). One or more spars (54, 56) including a stack (62) of uni-tape plies (63) having predominately a 0 degree fiber orientation with respect to the span (S) and extending spanwise outwardly through the root (20) and a portion (53) of the airfoil (12) towards the tip (47). Spars may include pressure and suction side spars (54, 56) sandwiching a chordwise extending portion (58) of the core section (50) in the airfoil (12) and which be located near or along the pressure and suction sides (41, 43) respectively. Chordwise extending portion (58) may be centered about a maximum thickness location (61) of the airfoil (12). Spars (54, 56) may have height (H), width (W), and thickness (T) that avoids flexural airfoil modes.

Description

COMPOSITE BLADE WITH UNI-TAPE AIRFOIL SPARS
BACKGROUND OF THE INVENTION FIELD OF THE INVENTION
The invention relates to gas turbine engine blades and, particularly, to composite blades.
DESCRIPTION OF RELATED ART
Composite blades made from elongated filaments
composited in a light-weight matrix have been developed for aircraft gas turbine engines. The blades are light-weight having high strength. The term composite has come to be defined as a material containing a reinforcement such as fibers or particles supported in a binder or matrix
material. Many composites are used in the aerospace
industry including both metallic and non-metallic
composites. The composites used for the blades disclosed herein are made of a unidirectional tape material and an epoxy resin matrix. A discussion of this and other suitable materials may be found in the "Engineering Materials
Handbook" by ASM INTERNATIONAL, 1987-1989 or later editions.
The composite blades disclosed herein are made from the non-metallic type made of a material containing a fiber such as a carbonaceous, silica, metal, metal oxide, or ceramic fiber embedded in a resin material such as Epoxy, PMR15, BMI, PEEU, etc. The fibers are unidirectionally aligned in a tape that is impregnated with a resin, formed into a part shape, and cured via an autoclaving process or press molding to form a light weight, stiff, relatively homogeneous article having laminates or plies within.
Composite fan blades have been developed for aircraft gas turbine engines to reduce weight and cost, particularly for fan blades in larger engines. A large engine composite wide chord fan blades offer a significant weight savings over a large engine having standard chorded fan blades. Among the problems, all gas turbine engine blades face resonance or flexural modes. Large composite fan blades for high bypass ratio aircraft gas turbine engines with
relatively wide diameter fans are faced with this problem. This is particularly true for the frequencies that cause the blade to experience first and second flexural airfoil modes (IF and 2F) .
It is highly desirable to provide light-weight and strong aircraft gas turbine engine fan blades that avoid passing through or experiencing assonance and flexural modes and, in particular, first and second flexural airfoil modes (IF and 2F) .
SUMMARY OF THE INVENTION
A gas turbine engine composite fan blade (10) includes an airfoil (12) having pressure and suction sides (41, 43) extending outwardly in a spanwise direction (S) from a blade root (20) of the blade (10) along a span (S) to a blade tip (47) . A core section (50) of the blade (10) includes composite quasi-isotropic plies (52) extending spanwise outwardly through the blade (10) including the root (20) and the airfoil (12) towards the tip (47) . One or more spars (54, 56) including a stack (62) of uni-tape plies (63) having a preferential 0 degree fiber orientation with respect to the span (S) spanwise outwardly through the root
(20) and through a portion (53) of the airfoil (12) towards the tip (47) .
The chordwise extending portion (58) of the core section (50) may be centered about a maximum thickness location (61) of the airfoil (12) . The spars (54, 56) may have a spanwise height (H) , chordwise width (W) , and spar thickness (TS) that avoids flexural airfoil modes such as first and second flexural airfoil modes. The one or more spars may include pressure and suction side spars (54, 56) sandwiching a chordwise extending portion (58) of the core section (50) in the airfoil (12) which may be located near or along the pressure and suction sides (41, 43)
respectively .
In one embodiment of the blade (10), the one or more spars include chordwise spaced apart upstream and downstream pressure side spars (74, 76) and chordwise spaced apart upstream and downstream suction side spars (78, 80)
sandwiching a chordwise extending portion (58) of the core section (50) in the airfoil (12) .
BRIEF DESCRIPTION OF THE DRAWINGS
The foregoing aspects and other features of the
invention are explained in the following description, taken in connection with the accompanying drawings where: FIG. 1 is a perspective view illustration of an aircraft gas turbine engine composite fan blade having a composite uni- tape spar.
FIG. 2 is a cross-sectional illustration of the
composite fan blade through 2-2 in FIG. 1.
FIG. 3 is a perspective diagrammatical view
illustration of an alternative aircraft gas turbine engine composite fan blade having a composite uni-tape spar.
FIG. 4 is a perspective diagrammatical view
illustration of the composite uni-tape spar illustrated in FIG. 3.
FIG. 5 is perspective diagrammatical view illustration of -P degree, 0 degree, and +P degree plies of the composite fan blade illustrated in FIG. 2.
FIG. 6 is a perspective view illustration of an
alternative aircraft gas turbine engine composite fan blade having a composite uni-tape spar.
FIG. 7 is a cross-sectional illustration of the
composite fan blade through 7-7 in FIG. 6.
DETAILED DESCRIPTION OF THE INVENTION
Illustrated in FIGS. 1 and 2 is a composite fan blade 10 for a high bypass ratio fanjet gas turbine engine (not shown) having a composite airfoil 12. Composite fan blade 10 is made up of filament reinforced laminations 30 formed from a composite material lay-up 36 of filament reinforced composite plies 40 (illustrated in FIG. 5) . As used herein, the terms "lamination" and "ply" are synonymous. The airfoil 12 includes pressure and suction sides 41, 43 extending outwardly in a spanwise direction from a fan blade root 20 along a span S to a blade tip 47. In the exemplary embodiment, the root 20 includes an integral dovetail 28 that enables the fan blade 10 to be mounted to a rotor disk.
The exemplary pressure and suction sides 41, 43
illustrated herein are concave and convex respectively. The airfoil 12 extends along a chord C between chordwise spaced apart leading and trailing edges LE, TE . Thickness T of the airfoil 12 varies in both chordwise and spanwise directions C, S and extends between pressure and suction sides 41, 43 of the blade 10 also referred to as convex and concave sides of the blade or airfoil. The airfoil 12 may be mounted on and be integral with a hub to form an integrally bladed rotor (IBR) or integrally with a disk in a BLISK
configuration .
The plies 40 are generally all made from a
unidirectional fiber filament ply material, preferably a tape, as it is often referred to, arranged generally in order of span and used to form a composite airfoil 12 as shown in FIG. 1. The plies 40 are essentially those plies that form the airfoil 12 and root 20 of the blade 10 as illustrated in FIGS. 1 and 3.
The composite fan blade 10 is made up of filament reinforced laminations 30 formed from a composite material lay-up 36 of different filament reinforced airfoil plies 40.
The blade 10 uses filament reinforced laminations or plies with a filament orientation of 0 degrees, +P degrees, and -P degrees as illustrated in FIG. 5. The angle P is a
predetermined angle as measured from 0 degrees which corresponds to a generally radially extending axis of the airfoil which may be its centerline or stacking line and is typically about 45 degrees. An exemplary arrangement is more particularly pointed out and explained in U.S. Patent No. 4,022,547 by Stanley.
Referring to FIGS. 1-4, the composite fan blade 10 includes a core section 50 of composite quasi-isotropic plies 52. Pressure and suction side spars 54, 56 sandwich a chordwise extending portion 58 of the core section 50 made of composite quasi-isotropic plies 52 generally near or along the pressure and suction sides 41, 43 respectively in the airfoil 12. The chordwise extending portion 58 of the core section 50 extends chordwise partially through the airfoil 12. The chordwise extending portion 58 of the core section 50 is generally centered chordwise in the airfoil
12. The exemplary embodiment of the chordwise extending portion 58 illustrated herein extends chordwise about 1/3 through the airfoil 12 and is generally centered chordwise about in the middle of the airfoil 12. The composite quasi-isotropic ply chordwise extending portion 58 of the core section 50 is preferably limited to a thicker cross sectional area of the airfoil 12 around or centered about a maximum thickness Tmax location 61 of the airfoil 12, as illustrated in FIG. 2, so as to be most effective. The Tmax location 61 is about a middle third of the airfoil between the leading and trailing edges LE, TE in the chordwise direction C for the exemplary airfoil illustrated herein. The pressure and suction side spars 54, 56 are made from stacks 62 of preferential 0 degree uni-tape plies 63 (see FIG. 5) with a 0 degree fiber orientation with respect to the span S .
Referring to FIGS. 3 and 4, the pressure and suction side spars 54, 56 (and the uni-tape plies they are made from extend) spanwise S through the fan blade root 20 and through a portion 53 of the airfoil 12 to a spar tip 57. The pressure and suction side spars 54, 56 have a spanwise height H as measured from the fan blade root 20 to the spar tip 57 which is less than the span S of the airfoil. In the embodiment of the composite fan blade 10 illustrated herein, the pressure and suction side spars 54, 56 (and the uni-tape plies are made from) extend all the way through the root 20 including the dovetail 28.
The quasi-isotropic ply core section 50 generally include alternating plies of tape with different +P, 0, and -P fiber orientations. The pressure and suction side spars 54, 56 include uni-tape plies with a predominately 0 degree fiber orientation. An exemplary blade ply lay-up is
disclosed in United States Patent No. 5,375,978, entitled "Foreign Object Damage Resistant Composite Blade and
Manufacture" to Evans, which issued December 27, 1994, is assigned to the same assignee of this patent, and is
incorporated herein by reference. The ply lay-up disclosed in 5,375,978 is referred to as a standard quasi-isotropic lay-up sequence of 0. degree, +45 degree, 0 degree, -45 degree fiber orientations with the plies having the numerous ply shapes.
The stacks 62 of the spars include uni-tape plies with a predominately 0 degree fiber orientation. A few of the plies may have another fiber orientation. An example is a stack having a total of 8 plies with 4 plies of 0 degree fiber orientation on both sides of two plies having +30 and a - 30 degree plies. This ply layup may be represented or denoted by 0 , 0 , 0 , 0 , +30 , -30 , 0 , 0 , 0 , 0.
Referring to FIGS. 1, 2, and 3, the spars have a spanwise height H, chordwise width W, and spar thickness TS designed to increase radial or spanwise stiffness of the airfoil 12 without increasing the weight of the blade. The spars are also designed or tailored or tuned to avoid flexural airfoil modes such as first and second flexural airfoil modes IF and 2F. The spanwise height H and the spar thickness TS are designed or tailored to tuned or avoid flexural airfoil modes such as first and second flexural airfoil modes IF and 2F. The uni-tape ply spar with
predominately a 0 degree fiber orientation allows for a stiffer blade without adding thickness and without adding weight and performance penalties. The exemplary embodiment of the composite blade illustrated herein is a fan blade but the composite blade with a quasi-isotropic ply core section and spars made from stacks 62 of 0 degree uni-tape plies 63 may also be used for other gas turbine engine blades such as compressor blades.
The exemplary embodiment of the composite blade 10 illustrated herein includes one or more outer cover plies 66 around the core section 50, made of composite
quasi-isotropic plies, and the pressure and suction side spars 54, 56. A leading edge metallic shield 68 is bonded around the leading edge LE . The shield is often referred to as metallic cladding.
Referring to FIGS. 6 and 7, an alternative spar design for the composite fan blade 10 includes a core section 50 of composite quasi-isotropic plies and two sets of pressure and suction side spars. The two sets include chordwise spaced apart upstream and downstream pressure side spars 74, 76 and chordwise spaced apart upstream and downstream suction side spars 78, 80 sandwiching the chordwise extending portion 58 of the core section 50 made of composite quasi-isotropic plies generally near or along the pressure and suction sides
41, 43 respectively.
The present invention has been described in an
illustrative manner. It is to be understood that the terminology which has been used is intended to be in the nature of words of description rather than of limitation.
While there have been described herein, what are considered to be preferred and exemplary embodiments of the present invention, other modifications of the invention shall be apparent to those skilled in the art from the teachings herein and, it is, therefore, desired to be secured in the appended claims all such modifications as fall within the true spirit and scope of the invention.
Accordingly, what is desired to be secured by Letters Patent of the United States is the invention as defined and differentiated in the following claims:

Claims

CLAIMS :
What is claimed is: 1. A gas turbine engine composite blade (10) comprising: an airfoil (12) having pressure and suction sides (41, 43) extending outwardly in a spanwise direction (S) from a blade root (20) of the blade (10) along a span (S) to a blade tip (47) ,
a core section (50) of the blade (10) including
composite quasi-isotropic plies (52) extending spanwise outwardly through the blade (10) including the root (20) and the airfoil (12) towards the tip (47),
one or more spars (54, 56) including a stack (62) of uni-tape plies (63) having predominately a 0 degree fiber orientation with respect to the span (S) , and
the one or more spars (54, 56) extending spanwise outwardly through the root (20) and through a portion (53) of the airfoil (12) towards the tip (47) .
2. The blade (10) as claimed in claim 1 further comprising the one or more spars including pressure and suction side spars (54, 56) sandwiching a chordwise extending portion (58) of the core section (50) in the airfoil (12) .
3. The blade (10) as claimed in claim 2 further comprising the chordwise extending portion (58) of the core section
(50) located near or along the pressure and suction sides
(41, 43) respectively.
4. The blade (10) as claimed in claim 1 further comprising the chordwise extending portion (58) of the core section (50) centered about a maximum thickness location (61) of the airfoil ( 12 ) .
5. The blade (10) as claimed in claim 4 further comprising the one or more spars including pressure and suction side spars (54, 56) sandwiching a chordwise extending portion (58) of the core section (50) in the airfoil (12) .
6. The blade (10) as claimed in claim 5 further comprising the chordwise extending portion (58) of the core section (50) located near or along the pressure and suction sides (41, 43) respectively.
7. The blade (10) as claimed in claim 1 further comprising the spars (54, 56) having a spanwise height (H) , chordwise width (W) , and spar thickness (TS) that avoids flexural airfoil modes.
8. The blade (10) as claimed in claim 7 further comprising the flexural airfoil modes including first and second flexural airfoil modes.
9. The blade (10) as claimed in claim 8 further comprising the one or more spars including pressure and suction side spars (54, 56) sandwiching a chordwise extending portion (58) of the core section (50) in the airfoil (12) .
10. The blade (10) as claimed in claim 9 further comprising the chordwise extending portion (58) of the core section
(50) located near or along the pressure and suction sides (41, 43) respectively.
11. The blade (10) as claimed in claim 8 further comprising the chordwise extending portion (58) of the core section
(50) centered about a maximum thickness location (61) of the airfoil ( 12 ) .
12. The blade (10) as claimed in claim 11 further
comprising the one or more spars including pressure and suction side spars (54, 56) sandwiching a chordwise extending portion (58) of the core section (50) in the airfoil ( 12 ) .
13. The blade (10) as claimed in claim 12 further
comprising the chordwise extending portion (58) of the core section (50) located near or along the pressure and suction sides (41, 43) respectively.
14. The blade (10) as claimed in claim 1 further comprising the one or more spars including chordwise spaced apart upstream and downstream pressure side spars (74, 76) and chordwise spaced apart upstream and downstream suction side spars (78, 80) sandwiching a chordwise extending portion (58) of the core section (50) in the airfoil (12) .
15. The blade (10) as claimed in claim 14 further
comprising the chordwise extending portion (58) of the core section (50) located near or along the pressure and suction sides (41, 43) respectively.
16. The blade (10) as claimed in claim 15 further
comprising the chordwise extending portion (58) of the core section (50) centered about a maximum thickness location (61) of the airfoil (12) .
17. The blade (10) as claimed in claim 16 further
comprising the upstream and downstream pressure side spars (74, 76) and the chordwise spaced apart upstream and
downstream suction side spars (78, 80) having a spanwise height (H) , chordwise width (W) , and spar thickness (TS) that avoids flexural airfoil modes.
18. The blade (10) as claimed in claim 17 further
comprising the flexural airfoil modes including first and second flexural airfoil modes.
19. The blade (10) as claimed in claim 17 further
comprising the chordwise extending portion (58) of the core section (50) located near or along the pressure and suction sides (41, 43) respectively.
20. The blade (10) as claimed in claim 19 further
comprising the flexural airfoil modes including first and second flexural airfoil modes.
21. The blade (10) as claimed in claim 1 further
comprising :
the root (20) includes an integral dovetail (28), one or more outer cover plies (66) around the core section (50) , and
a leading edge metallic shield (68) bonded around the leading edge (LE) .
22. The blade (10) as claimed in claim 21 further
comprising the spars (54, 56) having a spanwise height (H) , chordwise width (W) , and spar thickness (TS) that avoids flexural airfoil modes.
23. The blade (10) as claimed in claim 22 further
comprising the flexural airfoil modes including first and second flexural airfoil modes.
24. The blade (10) as claimed in claim 23 further
comprising the one or more spars including pressure and suction side spars (54, 56) sandwiching a chordwise
extending portion (58) of the core section (50) in the airfoil ( 12 ) .
25. The blade (10) as claimed in claim 24 further
comprising the chordwise extending portion (58) of the core section (50) centered about a maximum thickness location (61) of the airfoil (12) .
26. The blade (10) as claimed in claim 25 further
comprising the chordwise extending portion (58) of the core section (50) located near or along the pressure and suction sides (41, 43) respectively.
PCT/US2013/061294 2012-10-23 2013-09-24 Composite blade with uni-tape airfoil spars WO2014065968A1 (en)

Priority Applications (5)

Application Number Priority Date Filing Date Title
EP13774557.6A EP2917496A1 (en) 2012-10-23 2013-09-24 Composite blade with uni-tape airfoil spars
BR112015009060A BR112015009060A2 (en) 2012-10-23 2013-09-24 composite gas turbine engine blade.
CN201380054437.7A CN104981586A (en) 2012-10-23 2013-09-24 Composite blade with uni-tape airfoil spars
CA2888777A CA2888777A1 (en) 2012-10-23 2013-09-24 Composite blade with uni-tape airfoil spars
JP2015537713A JP6179961B2 (en) 2012-10-23 2013-09-24 Composite blade with unidirectional tape airfoil girder

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US13/658,209 US20140112796A1 (en) 2012-10-23 2012-10-23 Composite blade with uni-tape airfoil spars
US13/658,209 2012-10-23

Publications (1)

Publication Number Publication Date
WO2014065968A1 true WO2014065968A1 (en) 2014-05-01

Family

ID=49326854

Family Applications (1)

Application Number Title Priority Date Filing Date
PCT/US2013/061294 WO2014065968A1 (en) 2012-10-23 2013-09-24 Composite blade with uni-tape airfoil spars

Country Status (7)

Country Link
US (1) US20140112796A1 (en)
EP (1) EP2917496A1 (en)
JP (1) JP6179961B2 (en)
CN (1) CN104981586A (en)
BR (1) BR112015009060A2 (en)
CA (1) CA2888777A1 (en)
WO (1) WO2014065968A1 (en)

Families Citing this family (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10465703B2 (en) * 2016-04-11 2019-11-05 United Technologies Corporation Airfoil
US11168566B2 (en) * 2016-12-05 2021-11-09 MTU Aero Engines AG Turbine blade comprising a cavity with wall surface discontinuities and process for the production thereof
FR3081496B1 (en) 2018-05-24 2020-05-01 Safran Aircraft Engines FABRIC COMPRISING ARAMID FIBERS FOR PROTECTING A DAWN FROM IMPACTS
FR3081914B1 (en) 2018-06-05 2020-08-28 Safran Aircraft Engines BLOWER VANE IN COMPOSITE MATERIAL WITH LARGE INTEGRATED GAME
FR3098544B1 (en) 2019-07-11 2021-06-25 Safran Aircraft Engines Blower blade
FR3107300B1 (en) * 2020-02-18 2022-07-08 Safran Aircraft Engines COMPOSITE BLADE FOR TURBOMACHINE ROTOR
US11898464B2 (en) 2021-04-16 2024-02-13 General Electric Company Airfoil for a gas turbine engine

Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4022547A (en) 1975-10-02 1977-05-10 General Electric Company Composite blade employing biased layup
US5375978A (en) 1992-05-01 1994-12-27 General Electric Company Foreign object damage resistant composite blade and manufacture
EP1555391A2 (en) * 2004-01-15 2005-07-20 General Electric Company Hybrid ceramic matrix composite turbine blade
EP1980714A2 (en) * 2007-04-11 2008-10-15 General Electric Company Turbomachine blade

Family Cites Families (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5279892A (en) * 1992-06-26 1994-01-18 General Electric Company Composite airfoil with woven insert
US5655883A (en) * 1995-09-25 1997-08-12 General Electric Company Hybrid blade for a gas turbine
US6607358B2 (en) * 2002-01-08 2003-08-19 General Electric Company Multi-component hybrid turbine blade
US7575417B2 (en) * 2003-09-05 2009-08-18 General Electric Company Reinforced fan blade
FR2884550B1 (en) * 2005-04-15 2010-09-17 Snecma Moteurs PIECE FOR PROTECTING THE EDGE OF A BLADE
US7334997B2 (en) * 2005-09-16 2008-02-26 General Electric Company Hybrid blisk
FR2892339B1 (en) * 2005-10-21 2009-08-21 Snecma Sa PROCESS FOR MANUFACTURING A COMPOSITE TURBOMACHINE BLADE, AND BLADE OBTAINED BY THIS PROCESS
US7547194B2 (en) * 2006-07-31 2009-06-16 General Electric Company Rotor blade and method of fabricating the same
JP5457110B2 (en) * 2009-09-03 2014-04-02 住友精密工業株式会社 Ship propeller
FR2964411B1 (en) * 2010-09-06 2012-08-24 Messier Dowty Sa DUST OF TURBOREACTOR, IN PARTICULAR A DRAWER OF RECTIFIER, AND TURBOJET RECEIVING SUCH AS AUBES

Patent Citations (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4022547A (en) 1975-10-02 1977-05-10 General Electric Company Composite blade employing biased layup
US5375978A (en) 1992-05-01 1994-12-27 General Electric Company Foreign object damage resistant composite blade and manufacture
EP1555391A2 (en) * 2004-01-15 2005-07-20 General Electric Company Hybrid ceramic matrix composite turbine blade
EP1980714A2 (en) * 2007-04-11 2008-10-15 General Electric Company Turbomachine blade

Non-Patent Citations (1)

* Cited by examiner, † Cited by third party
Title
"Engineering Materials Handbook", 1987, ASM INTERNATIONAL

Also Published As

Publication number Publication date
EP2917496A1 (en) 2015-09-16
JP6179961B2 (en) 2017-08-16
JP2015537143A (en) 2015-12-24
BR112015009060A2 (en) 2017-07-04
CN104981586A (en) 2015-10-14
US20140112796A1 (en) 2014-04-24
CA2888777A1 (en) 2014-05-01

Similar Documents

Publication Publication Date Title
WO2014065968A1 (en) Composite blade with uni-tape airfoil spars
US8061997B2 (en) Damping device for composite blade
JP3924333B2 (en) Composite blade
US10066491B2 (en) Fibre composite component for the rotor blade of a wind turbine
US5375978A (en) Foreign object damage resistant composite blade and manufacture
EP1462606B1 (en) Multi-component hybrid turbine blade
US8419374B2 (en) Gas turbine engine composite blade
US9309772B2 (en) Hybrid turbine blade including multiple insert sections
EP2149711B1 (en) Blade, associated manufacturing system and manufacturing method
US5885059A (en) Composite tip cap assembly for a helicopter main rotor blade
US20050079334A1 (en) Method for making a pin reinforced, crack resistant fiber reinforced composite article
US8556579B2 (en) Composite aerofoil blade with wear-resistant tip
CN108026778B (en) Blade comprising a leading edge guard and method of producing the same
WO2013184491A1 (en) Composite structure with low density core and composite stitching reinforcement
EP2932042B1 (en) Attachment of composite article
EP3022396B1 (en) Vibration-damped composite airfoils and manufacture methods
US20140322482A1 (en) Composite article including composite to metal interlock and method of fabrication
US20130287588A1 (en) Fan blade
US9777579B2 (en) Attachment of composite article
US20230003129A1 (en) Composite airfoils with frangible tips

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: 13774557

Country of ref document: EP

Kind code of ref document: A1

ENP Entry into the national phase

Ref document number: 2888777

Country of ref document: CA

ENP Entry into the national phase

Ref document number: 2015537713

Country of ref document: JP

Kind code of ref document: A

NENP Non-entry into the national phase

Ref country code: DE

WWE Wipo information: entry into national phase

Ref document number: 2013774557

Country of ref document: EP

REG Reference to national code

Ref country code: BR

Ref legal event code: B01A

Ref document number: 112015009060

Country of ref document: BR

ENP Entry into the national phase

Ref document number: 112015009060

Country of ref document: BR

Kind code of ref document: A2

Effective date: 20150422