Composite airfoil for gas turbine

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section extending between a leading edge and a trailing edge in a chordwise direction and extending between a tip and a root section in a spanwise direction. A composite core defines the root section and a portion of the airfoil section. First and second skins extend along opposed sides of the composite core. 12. A method of forming a composite airfoil is also disclosed.

BACKGROUND

This disclosure relates to composite articles for a gas turbine engine, and more particularly to composite airfoils and methods of making the same.

A gas turbine engine typically includes at least a compressor section, a combustor section and a turbine section. The compressor section pressurizes air into the combustion section where the air is mixed with fuel and ignited to generate an exhaust gas flow. The exhaust gas flow expands through the turbine section to drive the compressor section and, if the engine is designed for propulsion, a fan section.

The fan section includes an array of airfoils carried by a fan hub. Some airfoils are made of one or more layers of a composite material. During operation, the airfoils may be subjected to impact by foreign objects, such as during a bird strike event. An impact may cause the layers to delaminate, which can result in loss of structural capability and liberation of plies.

SUMMARY

An airfoil for a gas turbine engine according to an example of the present disclosure includes an airfoil section extending between a leading edge and a trailing edge in a chordwise direction and extending between a tip and a root section in a spanwise direction. The airfoil section defines a suction side and a pressure side separated in a thickness direction. A composite core defines the root section and a portion of the airfoil section such that the composite core is spaced apart from the tip, and the composite core has a three-dimensional network of woven fibers. First and second skins extend along opposed sides of the composite core for at least some span positions of the airfoil section and join together to define the tip, and each of the first and second skins is a composite skin include a two-dimensional network of fibers.

In a further embodiment of any of the foregoing embodiments, the airfoil section is free of the core for at least a majority of span positions of the airfoil section.

In a further embodiment of any of the foregoing embodiments, the three-dimensional network of woven fibers is formed from a dry fiber preform.

In a further embodiment of any of the foregoing embodiments, the composite core and the first and second skins are formed together by resin transfer molding or by resin pressure molding to define the airfoil section and the root section.

In a further embodiment of any of the foregoing embodiments, the two-dimensional network of fibers is formed from a pre-impregnated fabric or a pre-impregnated tape.

In a further embodiment of any of the foregoing embodiments, the composite core terminates in the spanwise direction along the airfoil section prior to about 25% span.

In a further embodiment of any of the foregoing embodiments, the core is skewed toward the leading edge.

A gas turbine engine according to an example of the present disclosure includes a fan section that has a fan hub that carries a plurality of fan blades. The fan hub is rotatable about an engine longitudinal axis, a compressor section, and a turbine section that drives the compressor section and the fan section. Each of the plurality of fan blades includes a root section received in a respective slot defined by the fan hub, and an airfoil section extending between a leading edge and a trailing edge in a chordwise direction and extending between a tip and the root section in a spanwise direction. The airfoil section defines a suction side and a pressure side separated in a thickness direction. A composite core defines the root section and a portion of the airfoil section such that the composite core is spaced apart from the tip. First and second skins extend along opposed sides of the composite core for at least some span positions of the airfoil section and join together to define the tip, and each of the first and second skins is a composite skin including a two-dimensional network of fibers.

In a further embodiment of any of the foregoing embodiments, the composite core includes a three-dimensional network of woven fibers formed from a dry fiber preform. The two-dimensional network of fibers is formed from a pre-impregnated fabric or a pre-impregnated tape, and the composite core and the first and second skins are formed together by resin pressure molding or by resin transfer molding to define the airfoil section and the root section.

In a further embodiment of any of the foregoing embodiments, the core terminates in the spanwise direction along the airfoil section prior to about 25% span.

In a further embodiment of any of the foregoing embodiments, the root section defines a dovetail geometry that interfaces with the respective slot, and the first and second skins extend along the dovetail geometry to space apart the core from walls of the respective slot.

A method of forming a composite airfoil for a gas turbine engine, according to an example of the present disclosure includes fabricating a core that has a three-dimensional network of woven fibers from a dry fiber preform, fabricating first and second skins each having a two-dimensional network of fibers, arranging the first and second skins relative to the core in a mold, and forming the core and the first and second skins together with resin in the mold to define a composite airfoil such that the core defines a root section of the composite airfoil and the first and second skins extend at least partially along the core to define an airfoil section of the composite airfoil.

In a further embodiment of any of the foregoing embodiments, the step of forming includes curing the core and the first and second skins in the resin subsequent to delivering the resin into the mold.

In a further embodiment of any of the foregoing embodiments, the step of forming is performed using a resin transfer molding process to define the airfoil section and the root section.

In a further embodiment of any of the foregoing embodiments, the step of fabricating the first and second skins is performed by an automated fiber placement process.

In a further embodiment of any of the foregoing embodiments, the method includes partially curing the core prior to the step of arranging.

In a further embodiment of any of the foregoing embodiments, the step of forming is performed using a resin pressure molding process to define the airfoil section and the root section.

In a further embodiment of any of the foregoing embodiments, the two-dimensional network of fibers is formed from a dry fiber tape.

In a further embodiment of any of the foregoing embodiments, the two-dimensional network of fibers is formed from a pre-impregnated fabric or pre-impregnated tape.

In a further embodiment of any of the foregoing embodiments, the airfoil section extends between a tip and the root section in a spanwise direction, and the step of arranging is performed such that the airfoil section is free of the core for at least a majority of span positions of the airfoil section.

The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of an embodiment. The drawings that accompany the detailed description can be briefly described as follows.

DETAILED DESCRIPTION

Referring toFIG. 2, the fan42includes a fan hub43that carries a plurality of fan blades or airfoils60(one shown for illustrative purposes) that are rotatable about the engine longitudinal axis A. The airfoil60includes a root section62and an airfoil section64. The fan hub43defines a plurality of circumferentially spaced apart slots45. The root section62of each airfoil60is dimensioned to be slideably received within a respective slot45to mount the airfoil60in the engine10. The root section62can have a generally dovetail geometry that mates with walls of the slot45.

FIGS. 3-5illustrate an example composite airfoil160. In this disclosure, like reference numerals designate like elements where appropriate and reference numerals with the addition of one-hundred or multiples thereof designate modified elements that are understood to incorporate the same features and benefits of the corresponding original elements. The airfoil160can be a fan blade incorporated into the fan42ofFIGS. 1 and 2, for example. Other portions of the engine can benefit from the teachings herein, such as fan exit guide vanes41or airfoils in the compressor or turbine sections24,28ofFIG. 1. Other systems can also benefit from the teachings disclosed herein, including ground-based power generation systems.

The airfoil160includes a root section162and an airfoil section164. The airfoil section164extends in a spanwise or radial direction R from the root section162to a tip166. The tip166is a terminal end of the airfoil160. The airfoil section164generally extends in a chordwise or axial direction X between a leading edge L/E and a trailing edge T/E. The airfoil section164defines a pressure side P and a suction side S separated in a thickness direction T. Generally, the airfoil section164provides an aerodynamic surface for guiding airflow to downstream portions of the engine in response to rotation of the airfoil160. The root section162of the airfoil160is mounted to a rotor, such as the fan hub43ofFIGS. 1 and 2.

Referring toFIG. 6, span positions of the airfoil section164are schematically illustrated from 0% to 100% span in 10% increments to define a plurality of sections167. Each section167at a given span position is provided by a conical cut that corresponds to the shape of segments the bypass flowpath B or the core flow path C (FIG. 1), as shown by the large dashed lines. The 0% span position corresponds to the radially innermost location of the airfoil section164that meets the root section162. A 100% span position corresponds to a section of the airfoil160at the tip166.

Referring back toFIGS. 3-5, the airfoil160includes a composite core168, a first skin170and a second skin172. The core168defines the root section162and at least a portion of the airfoil section164, with the core168terminating at an edge or boundary174(shown in dashed lines inFIG. 3). The core168can extend along a mean camber line171(shown in dashed lines inFIG. 4) that bisects a thickness of the airfoil160for at least some span positions of the airfoil section164.

The core168can be dimensioned to provide increased rigidity to localized portions of the root and airfoil sections162,164. In the illustrated example ofFIG. 3, the core168defines a portion of the airfoil section164such that the boundary174of the core168is spaced apart from the tip166. In the illustrated example ofFIG. 3, the airfoil section164is free of the core168for at least a majority of span positions of the airfoil section164. In some examples, the boundary174of the core168terminates in the radial direction R along the airfoil section164prior to about 25% span, or more narrowly between about 5% span and about 20% span. For the purposes of this disclosure, the term “about” means ±3% of the respective quantity unless otherwise stated. In other examples, the core168defines a portion of the tip166, as illustrated by boundary174′ (FIG. 3).

In some examples, core168′ is skewed toward the leading edge L/E of the airfoil160, as illustrated by boundary174″ ofFIG. 3. A reference plane175is established along the airfoil section164between 0% and 100% span. The reference plane175is equidistant between the leading and trailing edges L/E, T/E for each respective span position. Core168″ is situated such that a center of mass CM″ of the core168″ is defined along the mean camber line171(FIG. 4) at a location between the leading edge L/E and the reference plane175, whereas a center of mass CM of the core168may be situated along the reference plane175, for example. Situating the core168″ relatively closer to the leading edge L/E can provide improved localized impact resistance, while skins170,172can provide relatively greater in-plane strength and stiffness to regions adjacent to the trailing edge T/E, for example. Terminating the core168prior to the tip166can provide the airfoil160with better in-plane strength and stiffness in regions where the core168is not present, but may have less delamination resistance.

The skins170,172extend along the core168to define external surfaces of the airfoil160. In the illustrated example ofFIGS. 3 and 5, the skins170,172extend along opposed sides of the core168to define the root section162. The root section162can define a dovetail geometry that interfaces with a respective slot145of a rotor (shown in dashed lines inFIG. 5). The skins170,172extend along the dovetail geometry to space apart the core168from walls of the respective slot145. The skins170,172can be relatively more compliant than the core168, and can provide improved absorption of impacts and load transfer between the airfoil160and the fan hub43(FIG. 2) during engine operation, which can improve durability of the airfoil160. In alternative examples, the root section162is free of the skins170,172such that the core168abuts the walls of the respective slot145.

The skins170,172extend along opposed sides of the core168for at least some span positions of the airfoil section164and join together to define the tip166. The boundary174of the core168can be contoured or tapered to provide a relatively smooth interface between the skins170,172. The skins170,172are joined together along the leading and trailing edges L/E, T/E and define the pressure and suction sides P, S. In some examples, the core168defines at least a portion of the leading and/or trailing edges L/E, T/E of the airfoil160. In other examples, the core168is spaced apart from the leading and/or trailing edges L/E, T/E.

The core168and skins170,172can be made of various composite materials to define the airfoil160. The core168and skins170,172can be constructed from fibers arranged in various orientations and in one or more layers based on structural requirements. In some examples, the core168and/or skins170,172include carbon fibers. The core168and/or skins170,172can be constructed from other materials, including fiberglass, an aramid such as Kevlar®, a ceramic such as Nextel™, and a polyethylene such as Spectra®.

In the illustrated example ofFIGS. 3-5, the core168includes and is constructed from a three-dimensional network of fibers176(FIG. 7A), which can be woven or interlaced. The three-dimensional network of fibers176can be formed from a dry fiber preform, for example. Each of the skins170,172can be a composite skin that includes and is constructed from a two-dimensional network of fibers178(FIG. 7B), which can be woven or interlaced.

The two-dimensional network of fibers178can be formed from a pre-impregnated (“prepreg”) fabric or a pre-impregnated tape, for example. Pre-impregnating the fibers with resin can provide relatively greater strength and toughness to the composite article. Matrix (resin) materials used for prepreg can have greater toughness compared to matrix materials used with resin infusion methods, such that a part made from prepreg is relatively stronger, as measured by delamination resistance.

In other examples, the skins170,172include a plurality of relatively thin uni-tape plies having a plurality of fibers oriented in the same direction. In some examples, the core168and/or skins170,172can include a network of biaxial braids180(FIG. 7C). Other configurations for the core168and/or skins170,172can include a network of tri-axial braids182(FIG. 7D), for example. In other examples, the skins170,172include a network of stitched or non-crimped fabrics. The example fiber constructions are known in the art, but the incorporation of the fiber constructions into airfoil160to define the core168and skins170,172utilizing the teachings herein are not known.

The core168and/or skins170,172can include different fiber types in the fiber directions to tailor the strength and stiffness of the core168and/or skins170,172. For example, high modulus carbon fibers may be used in conjunction with low modulus carbon fibers. In yet another example, fiberglass or aramid fibers may be used in combination with carbon fibers.

Incorporating a three-dimensional network of fibers into the core168, including the three-dimensional network of woven or interlaced fibers176(FIG. 7A), provides reinforcement in localized portions of the airfoil160that may be more susceptible to degradation due to impact from foreign objects. Incorporating a two-dimensional network of fibers into the skins170,172, including the two-dimensional network of fibers178(FIG. 7B), can provide relatively greater in-plane strength and stiffness to other portions of the airfoil160.

Various techniques can be utilized to form a composite article for a gas turbine engine such as the airfoil160.FIG. 8illustrates an example process in a flow chart184for forming a composite article or airfoil, such as airfoil160. Reference is made to airfoil160for illustrative purposes. The techniques disclosed herein can improve structural properties of the airfoil160, which can improve durability, and can reduce manufacturing cost.

At step186, a core168is fabricated or otherwise prepared. The core168can be fabricated from a dry fiber preform to form or otherwise define a three-dimensional network of woven fibers. The core168may be fabricated on a loom using a weaving process, for example. A tackifier can be applied to the preform to prepare the core168. The core168can be dimensioned according to expected stresses or impacts that may be observed during use of the composite article.

At step188, first and second skins170,172are fabricated or otherwise prepared. The skins170,172can be prepared from a pre-impregnated fabric or pre-impregnated tape to form or otherwise define a two-dimensional network of fibers.

The core168and skins170,172can be arranged to establish a layup. The layup can be formed on a tool. Steps186and/or188can be performed by an automated fiber placement (AFP) process, or the steps can be performed manually by hand placement of a three-dimensional woven core and hand placing of prepreg plies to form the skins. AFP is generally known, and includes placement of narrow strips of unidirectional material or “tows” to build up the composite layers that constitute the article according to a predefined geometry. Forming the three-dimensional woven core168using the techniques disclosed herein can eliminate the need to form relatively small plies or layers that may otherwise be needed to fill a volume of the airfoil to define the airfoil geometry but that may not be practical to use in an AFP process.

At step190, the skins170,172are assembled or otherwise arranged relative to the core168in a mold. Surfaces of the mold can be dimensioned according to an external profile or contour of the airfoil160. In some examples, step186includes partially curing the core168at step187prior to step190and prior to positioning the core168in the mold, which may be referred to as a “b-staged” core. Step187can include infiltrating the core168with a resin that is chemically compatible with a resin used for subsequent injection. Once infiltration occurs, the core168is partially cured or “b-staged”. As known, prepreg is already b-staged.

Steps186,188and/or190can be performed such that the airfoil section164is free of the core168for at least a majority of span positions of the airfoil section164, including the core168terminating at any of the span positions disclosed herein.

At step192, a composite airfoil such as airfoil160is formed. Step192can include forming the core168and the skins170,172together with resin in the mold to define the airfoil160, and such that the core168defines the root section162of the airfoil160and the skins170,172extend at least partially along the core168to define at least a portion of the root section162and the airfoil section164of the airfoil160. Resin materials can include a thermoset epoxy, for example, and can infuse the dry fiber preform.

The process184, including step192, can be performed using a closed-molding process. For example, step192can be performed using a resin transfer molding (RTM) process or a resin pressure molding (RPM) process to define the airfoil160including forming the core168and skins170,172together to define the root and airfoil sections162,164.

Resin transfer molding (RTM) is generally known for manufacturing composite articles. RTM is a closed-molding process that typically includes fabricating a fiber preform by laying up plies of fiber sheets in a stack, placing the fiber preform in a closed mold, and then saturating the fiber preform with a liquid thermoset resin. The resin is typically mixed with a catalyst or hardener prior to being injected into the closed mold, or can be previously mixed together in a one-part resin system. One-part resin systems already have the catalyst mixed with the resin. The article is heated in the mold to a desired temperature to cure the article. The mold can be heated using a liquid heating system, for example. In some examples, the mold is heated by direct contact with heated platens such as in a compression press or free-standing in an oven. A variation of RTM is vacuum-assisted resin transfer molding (VARTM). In a VARTM process, a vacuum is used to draw the resin into the mold. The RTM process generally results in a part with a slightly lower volume percentage of fiber compared to a part made from prepreg and processed in an autoclave.

Resin pressure molding (RPM) is generally known for manufacturing composite articles. RPM can be considered a variation of an RTM process. RPM is a closed-molding process which includes delivering a liquid resin into a closed mold in which some, or all, of the fiber reinforcement has been pre-impregnated with a resin. Thereafter and similar to RTM, a combination of elevated heat and hydrostatic resin pressure are applied to the mold to cure the article.

Step192can include curing the core168and skins170,172in the resin at step198. Step198can occur subsequent to injecting or otherwise delivering the resin into the mold. Step192can include heating the mold to a predetermined temperature for a set period of time to at least partially or fully cure the core168and skins170,172. One would understand how to determine the temperature and time period to cure the core168and skins170,172utilizing the teachings herein. Step192can include applying resin pressure concurrently with heat to the core168, skins170,172and resin at step196, such as during an RTM or RPM process. The RPM process can include utilization of a b-staged core168and prepreg skins170,172, or a dry fiber core168and prepreg skins170,172, for example. Forming the airfoil160with an RPM process can close any gaps in the composite article and can yield relatively better dimensional control of the cured article. This technique can eliminate separate injection and curing steps for the core168.

Pre-impregnated fabric or tape can provide the ability to use resin systems with a higher toughness and impact resistance as compared to resin that may be used in a transfer molding (RTM) process to infuse the dry fiber tape. The dry fiber preform used to form the core168is infused with resin, but the relatively lower toughness resin that may used in the RTM process is adequate since the three-dimensional network of fibers used to form the core168can provide relatively greater toughness and impact resistance. In other examples, the skins170,172including a two-dimensional network of fibers is formed from a dry fiber tape. A binder material can be added to the dry fiber tape to hold the layers of the fiber tape together and promote adhesion.

One or more finishing procedures can be performed on the cured composite article defining the airfoil160at step194. Example finishing procedures can include one or more grinding operations to remove excess material at parting lines caused by the molding process, or final dimensioning of the airfoil160geometry.

In alternative examples, the core168is staged on a tool. The core168is integrated with the skins170,172. The core168and skins170,172are then cured together in an autoclave.