CA2888777A1 - Composite blade with uni-tape airfoil spars - Google Patents

Composite blade with uni-tape airfoil spars Download PDF

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Publication number
CA2888777A1
CA2888777A1 CA2888777A CA2888777A CA2888777A1 CA 2888777 A1 CA2888777 A1 CA 2888777A1 CA 2888777 A CA2888777 A CA 2888777A CA 2888777 A CA2888777 A CA 2888777A CA 2888777 A1 CA2888777 A1 CA 2888777A1
Authority
CA
Canada
Prior art keywords
blade
airfoil
spars
core section
chordwise
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Abandoned
Application number
CA2888777A
Other languages
French (fr)
Inventor
Nicholas Joseph Kray
Ian Francis Prentice
Tod Winton Davis
Dong-Jin Shim
Pranav Dhoj Shah
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.)
General Electric Co
Original Assignee
General Electric Co
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 Co filed Critical General Electric Co
Publication of CA2888777A1 publication Critical patent/CA2888777A1/en
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
    • F01D5/28—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

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Materials Engineering (AREA)
  • Composite Materials (AREA)
  • Architecture (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)

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 (1F 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 (1F 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
2 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
3 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 5 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 10 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
4 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
5 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 1F 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
6 airfoil modes 1F 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
7
8 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 (26)

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.
CA2888777A 2012-10-23 2013-09-24 Composite blade with uni-tape airfoil spars Abandoned CA2888777A1 (en)

Applications Claiming Priority (3)

Application Number Priority Date Filing Date Title
US13/658,209 2012-10-23
US13/658,209 US20140112796A1 (en) 2012-10-23 2012-10-23 Composite blade with uni-tape airfoil spars
PCT/US2013/061294 WO2014065968A1 (en) 2012-10-23 2013-09-24 Composite blade with uni-tape airfoil spars

Publications (1)

Publication Number Publication Date
CA2888777A1 true CA2888777A1 (en) 2014-05-01

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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)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US11982205B1 (en) * 2022-12-28 2024-05-14 General Electric Company Airfoil having a spar assembly for a turbine engine

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EP2917496A1 (en) 2015-09-16
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