JP6179961B2 - Composite blade with unidirectional tape airfoil girder - Google Patents

Description

  The present invention relates to gas turbine engine blades, and more particularly, to composite blades.

  Composite blades consisting of elongated filaments synthesized into a lightweight matrix have been developed for aircraft gas turbine engines. This blade is high strength and lightweight. The term composite is being defined as a material that includes a reinforcing material, such as fibers or particles held in a binder or matrix. Many composites are used in the aerospace industry, including both metallic and non-metallic composites. The composite material used for the blade disclosed in the present specification is formed of a unidirectional tape material and an epoxy resin base material. A description of this material, and other suitable materials, can be found in “Engineered Materials Handbook” (1987-1989 or later edition) by ASM INTERNATIONAL.

  The composite blade disclosed in the present specification is made of a material including a fiber such as carbon, silica, metal, metal oxide, or ceramic fiber embedded in a resin material such as epoxy, PMR15, BMI, or PEEU. Made of metal. The fibers are impregnated with resin, formed into part shapes, and autoclaved to form lightweight, rigid, relatively homogeneous articles that are unidirectionally arranged on the tape and have a laminate or layer inside. And cured through press molding.

  Composite fan blades have been developed for aircraft gas turbine engines, particularly for larger engine fan blades, to reduce weight and cost. Large engine composite wide chord fan blades provide significant weight savings compared to large engines with standard chord fan blades. The problem is that all gas turbine engine blades face resonance and bending modes. Large composite fan blades for high bypass ratio aircraft gas turbine engines with relatively wide fan diameters face this problem. This is especially true for frequencies where the blade undergoes first and second bent airfoil modes (1F, 2F).

  To provide a lightweight and powerful aircraft gas turbine engine fan blade that does not experience or experience similar or bending modes, particularly the first and second bent airfoil modes (1F, 2F). Is highly desired.

European Patent Application No. 1980714

  The composite fan blade (10) of the gas turbine engine is blade lengthwise (S) outward from the blade root (20) of the blade (10) along the length (S) to the blade tip (47). And an airfoil (12) having a pressure side (41) and a suction side (43). The core portion (50) of the blade (10) passes through the blade (10) having the root (20) and the airfoil portion (12) toward the tip (47), and is a composite material extending outward in the blade length direction. Includes a pseudo-isotropic layer (52). The one or more girders (54, 56) comprise a laminate (62) of unidirectional tape layers (63) having a fiber orientation of predominantly 0 degrees relative to the length (S) In the direction, it passes through the root (20) and passes through a part (53) of the airfoil (12) toward the tip (47).

  The chord extension (58) of the core (50) can be centered on the maximum thickness position (61) of the airfoil (12). The spar (54, 56) avoids bending airfoil modes such as the first and second bending airfoil modes, the wing length height (H), the chord width (W), and It can have a girder thickness (TS). The one or more spar may include a pressure side spar (54) and a suction side spar (56) sandwiching the chord extension (58) of the core (50) in the airfoil (12). Can be disposed near each of the pressure side (41) and suction side (43) or along each of the pressure side (41) and suction side (43).

  In one embodiment of the blade (10), the one or more girders are spaced upstream in the chord direction, sandwiching the chord extension (58) of the core (50) at the airfoil (12). And downstream pressure side girders (74, 76) and upstream and downstream suction side girders (78, 80) spaced in the chord direction.

  The above aspects and other features of the invention are described in the following description with reference to the accompanying drawings.

1 is a perspective view of an aircraft gas turbine engine composite fan blade having composite unidirectional tape girders. FIG. FIG. 2 is a cross-sectional view of the composite fan blade through 2-2 of FIG. FIG. 6 is a perspective view of another aircraft gas turbine engine composite fan blade having composite unidirectional tape girders. FIG. 4 is a perspective view of the girder of the composite unidirectional tape shown in FIG. 3. FIG. 3 is a perspective view of a −P degree, 0 degree, and + P degree layer of the composite fan blade shown in FIG. 2. FIG. 5 is a perspective view of another aircraft gas turbine engine composite fan blade having composite unidirectional tape girders. FIG. 7 is a cross-sectional view of the composite fan blade taken through 7-7 of FIG.

  1 and 2 illustrate a composite fan blade 10 for a high bypass ratio fan jet gas turbine engine (not shown) having a composite airfoil 12. Composite fan blade 10 is comprised of a filament reinforced laminate 30 formed from a composite layup 36 of a layer 40 of filament reinforced composite (see FIG. 5). As used herein, the terms “stack” and “layer” are synonymous. The airfoil 12 has a pressure side 41 and a suction side 43 that extend outwardly in the blade length direction from the root 20 of the fan blade to the blade tip 47 along the length S. In the exemplary embodiment, root 20 includes an integral dovetail 28 that allows attachment of fan blade 10 to a rotor disk.

  The exemplary pressure side 41 and suction side 43 shown herein are concave and convex, respectively. The airfoil 12 extends along the chord direction C between the leading edge LE and the trailing edge TE which are spaced apart in the chord direction. The thickness T of the airfoil portion 12 varies in both the chord direction C and the blade length direction S, and extends between the pressure side surface 41 and the suction side surface 43 of the blade 10. Pressure side 41 and suction side 43 are also referred to as convex and concave surfaces of the blade or airfoil. The airfoil 12 is attached to the hub and may be integrated with the hub to form a blade-integrated rotor (IBR) or a blisk configuration integrated with the disk.

  Layer 40 is generally all made of a unidirectional fiber filament layer material, preferably tape, as often referred to. Layers 40 are arranged in approximately length order and are used to form a composite airfoil 12 as shown in FIG. As shown in FIGS. 1 and 3, the layer 40 is essentially the layer that forms the airfoil 12 and the root 20 of the blade 10.

  The composite fan blade 10 is comprised of a filament reinforced laminate 30 formed from a composite layup 36 of layers 40 of different filament reinforced airfoils. The blade 10 uses a filament reinforced laminate or layer having filament orientations of 0 degrees, + P degrees, and -P degrees as shown in FIG. The angle P is a pre-determined angle measured from 0 degrees, which corresponds to an airfoil centerline that can be the centerline of the airfoil, or a line that is stacked, that extends approximately radially. It is usually about 45 degrees. An exemplary arrangement is specifically pointed out and described in US Pat. No. 4,022,547 by inventor Stanley.

  With reference to FIGS. 1 to 4, the composite fan blade 10 includes a core portion 50 of a composite pseudo-isotropic layer 52. The pressure side girder 54 and the suction side girder 56 are arranged in the airfoil 12 substantially near each of the pressure side surface 41 and the suction side surface 43 or substantially along each of the pressure side surface 41 and the suction side surface 43. The chord direction extension portion 58 of the core portion 50 made of the composite pseudo isotropic layer 52 is sandwiched. The chord extension 58 of the core 50 extends partially in the chord direction through the airfoil 12. The chord direction extension part 58 of the core part 50 is located in the airfoil part 12 substantially at the center in the chord direction. The exemplary embodiment of the chordal extension 58 shown herein has a chord extending about one third through the airfoil 12 and in the middle of the airfoil 12, the chord. Is located almost in the center of. The chordal extension 58 of the composite quasi-isotropic layer of the core 50 is preferably limited to the periphery of the thicker section of the airfoil 12 as shown in FIG. It is most effective to center on the maximum thickness Tmax position 61 of the airfoil 12. The Tmax position 61 occupies approximately one third of the center of the airfoil in the chord direction C between the leading edge LE and the trailing edge TE of the exemplary airfoil shown herein. The pressure side spar 54 and the suction side spar 56 are made of a laminate 62 of unidirectional tape layers 63 (see FIG. 5) having a fiber orientation of 0 degrees with respect to the length S and having a fiber orientation of 0 degrees.

  Referring to FIGS. 3 and 4, the pressure side spar 54 and the suction side spar 56 (and the unidirectional tape layer forming the same) pass through the root 20 of the fan blade and through the portion 53 of the airfoil 12. It extends in the blade length direction S up to the beam tip 57. The pressure side girder 54 and the suction side girder 56 have a height H in the blade length direction measured from the root 20 of the fan blade to the girder tip 57, and this height H is larger than the length S of the airfoil portion. small. In one embodiment of the composite fan blade 10 shown herein, the pressure side spar 54 and the suction side spar 56 (and the unidirectional tape layer forming the same) extend over the entire length of the root 20 including the dovetail 28. ing.

  The core portion 50 of the quasi-isotropic layer generally includes layers in which tapes having different fiber orientations of + P, 0, and −P are alternately arranged. The pressure side spar 54 and the suction side spar 56 include a unidirectional tape layer having a fiber orientation of primarily 0 degrees. An exemplary blade layer layup is disclosed in US Pat. No. 5,375,978, issued December 27, 1994, entitled “Foreign Object Damage Resistant Composite Blade and Manufacture” by Inventor Evans. Which is assigned to the same assignee as the present invention and is incorporated herein by reference. In the layer layup disclosed in US Pat. No. 5,375,978, fiber orientations of 0 degrees, +45 degrees, 0 degrees, and −45 degrees are standard pseudo-isotropic in layers having multiple layer shapes. Referred to as sex layup order.

  The spar laminate 62 includes a unidirectional tape layer having primarily a 0 degree fiber orientation. Some layers may have different fiber orientations. As an example, a laminated body has eight layers in total of four layers of fiber orientation of 0 degree on both sides of two layers of +30 degrees and -30 degrees. The layup of this layer can be expressed as 0, 0, 0, 0, +30, -30, 0, 0, 0, 0.

  1-3, the spar is designed to increase the radial or wing length stiffness of the airfoil 12 without increasing the weight of the blade. It has a chord width W and a girder thickness TS. The spar is also designed or adjusted to avoid bending airfoil modes, such as the first and second bending airfoil modes 1F and 2F. The height H in the wing length direction and the thickness TS of the spar are designed or adjusted to avoid bending airfoil modes, such as the first and second bending airfoil modes 1F and 2F. The unidirectional tape layer girder, which has a fiber orientation of predominantly 0 degrees, makes it possible to increase the rigidity of the blade without adding thickness and without adding weight and performance penalties. An exemplary embodiment of the composite blade shown herein is a fan blade, but a girder made from a core portion of a quasi-isotropic layer and a laminate 62 of zero degree unidirectional tape layers 63. The composite blade with can also be used for blades of other gas turbine engines such as compressor blades.

  The exemplary embodiment of the composite blade 10 shown herein includes one or more skin layers 66 around the core portion 50 of a composite quasi-isotropic layer, and a pressure side spar 54 and a suction side spar. 56. A leading edge metal shield 68 is joined around the leading edge LE. This shield is often called a metal sheath.

  Referring to FIGS. 6 and 7, another spar design of the composite fan blade 10 includes a composite quasi-isotropic layer core 50 and two sets of pressure side and suction side spar. The two sets include upstream and downstream pressure side girders 74, 76 spaced in the chord direction and upstream and downstream suction side girders 78, 80 spaced in the chord direction. A chordwise extension 58 of the core portion 50 made of a composite quasi-isotropic layer substantially near each of the suction side surfaces 43 or substantially along each of the pressure side surfaces 41 and suction side surfaces 43. It is sandwiched.

  The invention has been described in an illustrative manner. It is to be understood that the terminology used is intended to be illustrative rather than limiting. While what has been considered to be preferred and exemplary embodiments of the invention has been described herein, other modifications of the invention will be apparent to those skilled in the art from the teachings herein, and thus the present All such modifications that fall within the true spirit and scope of the invention are desired to be protected by the appended claims.

  Accordingly, what is desired to be secured by patent is an invention that is defined and specified in the accompanying claims.

10 Blade 12 Airfoil 20 Root 28 Dovetail 30 Filament Reinforced Laminate 36 Composite Material Lay-up 40 Layer 41 Pressure Side 43 Pressure Side 47 Tip 50 Core 52 Composite Pseudoisotropic Layer 53 Part of Airfoil 54 Pressure side girder 56 Vacuum side girder 57 Girder tip 58 Chord direction extension 61 Tmax position 62 Laminate 63 Unidirectional tape layer 66 Outer skin layer 68 Leading edge metal shield 74 Upstream pressure side girder 76 Downstream pressure side girder 78 Upstream Suction side beam 80 downstream suction side beam 1F first bending airfoil mode 2F second bending airfoil mode C chord direction H height LE leading edge P angle S length in airfoil length T thickness TE Trailing edge TS Girder thickness W Width

Claims (26)

A gas turbine engine composite blade (10) comprising:
A pressure side surface (41) and a negative side surface extending outwardly in the blade length direction (S) from the blade root (20) of the blade (10) to the blade tip (47) along the length (S). An airfoil (12) having a pressure side (43);
A composite quasi-isotropic layer (52) extending outward in the blade length direction through the blade (10) having the airfoil portion (12) toward the root (20) and the tip (47). Including the core (50) of the blade (10),
One or more girders (54, 56) comprising a laminate (62) of unidirectional tape layers (63) having a fiber orientation of predominantly 0 degrees relative to the length (S);
The one or more girders (54, 56) pass through the root (20) toward the tip (47) and through a part (53) of the airfoil (12) in the wing length direction. Extending outwards,
Gas turbine engine composite blade (10).   The one or more airfoil portions (12) including a pressure side girder (54) and a suction side girder (56) sandwiching the chord direction extension portion (58) of the core portion (50). The blade (10) of claim 1, further comprising:   The core portion (50) disposed near each of the pressure side surface (41) and the suction side surface (43) or along each of the pressure side surface (41) and the suction side surface (43). The blade (10) of claim 2, further comprising the chordal extension (58).   The blade (10) according to claim 1, further comprising the chordal extension (58) of the core (50) about the maximum thickness position (61) of the airfoil (12). .   The airfoil portion (12) includes the pressure side girder (54) and the suction side girder (56) sandwiching the chord direction extension (58) of the core portion (50). The blade (10) of claim 4, further comprising a spar.   The core portion (50) disposed near each of the pressure side surface (41) and the suction side surface (43) or along each of the pressure side surface (41) and the suction side surface (43). The blade (10) of claim 5, further comprising the chordal extension (58) of   The spar (54, 56) further comprising a wing length height (H), a chord width (W), and a spar thickness (TS) to avoid a bent airfoil mode. The blade (10) according to item 1.   The blade (10) of claim 7, further comprising the bent airfoil mode, including first and second bent airfoil modes.   The one or more airfoil portions (12) including a pressure side girder (54) and a suction side girder (56) sandwiching a chord direction extension (58) of the core portion (50). The blade (10) of claim 8, further comprising:   The core portion (50) disposed near each of the pressure side surface (41) and the suction side surface (43) or along each of the pressure side surface (41) and the suction side surface (43). The blade (10) of claim 9, further comprising a chordal extension (58) of   The blade (10) of claim 8, further comprising the chordal extension (58) of the core (50) about the maximum thickness position (61) of the airfoil (12). .   The airfoil portion (12) includes the pressure side girder (54) and the suction side girder (56) sandwiching the chord direction extension (58) of the core portion (50). The blade (10) of claim 11, further comprising a spar.   The core portion (50) disposed near each of the pressure side surface (41) and the suction side surface (43) or along each of the pressure side surface (41) and the suction side surface (43). The blade (10) of claim 12, further comprising the chordal extension (58) of   In the airfoil portion (12), upstream and downstream pressure side girders (74, 76) spaced apart in the chord direction sandwiching the chord direction extension portion (58) of the core portion (50), and a blade The blade (10) of claim 1, further comprising the one or more girders, including upstream and downstream suction side girders (78, 80) spaced in a chordal direction.   The core portion (50) disposed near each of the pressure side surface (41) and the suction side surface (43) or along each of the pressure side surface (41) and the suction side surface (43). The blade (10) of claim 14, further comprising a chordal extension (58) of   The blade (10) according to claim 15, further comprising the chordal extension (58) of the core (50) about the maximum thickness position (61) of the airfoil (12). .   The upstream and downstream pressure side girders (74, 74) having a wing length (H), a chord width (W), and a spar thickness (TS) to avoid bending airfoil modes. The blade (10) of claim 16, further comprising: 76) and upstream and downstream suction side girders (78, 80) spaced apart in the chord direction.   The blade (10) of claim 17, further comprising the bent airfoil mode including first and second bent airfoil modes.   The core portion (50) disposed near each of the pressure side surface (41) and the suction side surface (43) or along each of the pressure side surface (41) and the suction side surface (43). 18. The blade (10) of claim 17, further comprising the chordal extension (58).   The blade (10) of claim 19, further comprising the bent airfoil mode including first and second bent airfoil modes. Said root (20) including an integral dovetail (28);
One or more skin layers (66) around the core (50);
A leading edge metal shield (68) joined around the leading edge (LE);
The blade (10) according to claim 1.   The spar (54, 56) further comprising a wing length height (H), a chord width (W), and a spar thickness (TS) to avoid a bent airfoil mode. Item 24. The blade according to Item 21.   23. The blade (10) of claim 22, further comprising the bent airfoil mode, including first and second bent airfoil modes.   The airfoil portion (12) includes the pressure side girder (54) and the suction side girder (56) sandwiching the chord direction extension (58) of the core portion (50). The blade (10) of claim 23, further comprising a spar.   The blade (10) according to claim 24, further comprising the chord extension (58) of the core (50) about the maximum thickness position (61) of the airfoil (12). .   The core portion (50) disposed near each of the pressure side surface (41) and the suction side surface (43) or along each of the pressure side surface (41) and the suction side surface (43). 26. The blade (10) of claim 25, further comprising the chordal extension (58).