Shroud for splitter and rotor airfoils of a fan for a gas turbine engine

A fan for a gas turbine engine is provided. The fan includes a rotor including at least one rotor stage having a rotatable disk defining a flowpath surface and an array of rotor airfoils extending outward from the flowpath surface; an array of splitter airfoils extending outward from the flowpath surface, wherein the splitter airfoils and the rotor airfoils are disposed in a sequential arrangement; and a shroud extending between the rotor airfoils and the splitter airfoils.

FIELD

The present subject matter relates generally to a gas turbine engine, or more particularly to a gas turbine engine having a fan assembly and a shroud for splitter and rotor airfoils.

BACKGROUND

A gas turbine engine generally includes a fan and a core arranged in flow communication with one another. Additionally, the core of the gas turbine engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.

It is desirable to improve gas turbine engine performance, e.g., turbofan engine performance, over wider operating ranges than is presently possible. Multi-stage fans in turbofan engines can be relatively complex, heavy, and require additional moving parts in the fan module. Thus, it is desirable to improve rotor aerodynamic performance and reduce the weight of such fans.

BRIEF DESCRIPTION

In one exemplary embodiment of the present disclosure, a fan for a gas turbine engine is provided. The fan includes a rotor including at least one rotor stage having a rotatable disk defining a flowpath surface and an array of rotor airfoils extending outward from the flowpath surface; an array of splitter airfoils extending outward from the flowpath surface, wherein the splitter airfoils and the rotor airfoils are disposed in a sequential arrangement; and a shroud extending between the rotor airfoils and the splitter airfoils.

In certain exemplary embodiments the shroud is connected to the rotor airfoils at a position radially inward of a tip of each of the rotor airfoils and radially outward of a root of each of the rotor airfoils.

In certain exemplary embodiments the shroud is connected to the splitter airfoils at a position radially inward of a tip of each of the splitter airfoils and radially outward of a root of each of the splitter airfoils.

In certain exemplary embodiments the shroud is connected to a portion of a tip of each of the splitter airfoils.

In certain exemplary embodiments a first portion of the shroud is connected to a first splitter airfoil at a position radially inward of a tip of the first splitter airfoil and radially outward of a root of the first splitter airfoil, and wherein a second portion of the shroud is connected to a second splitter airfoil at a tip of the second splitter airfoil.

In certain exemplary embodiments a third portion of the shroud is disposed relative to a third splitter airfoil at a position adjacent to a tip of the third splitter airfoil.

In certain exemplary embodiments the shroud includes a plurality of linked shroud sections.

In certain exemplary embodiments a chord dimension of the splitter airfoil is less than a chord dimension of the rotor airfoil.

In certain exemplary embodiments a span dimension of the splitter airfoil is less than a span dimension of the rotor airfoil.

In certain exemplary embodiments the splitter airfoils and the rotor airfoils are disposed in a staggered and alternating arrangement.

In certain exemplary embodiments each splitter airfoil is located approximately midway between two adjacent rotor airfoils.

In another exemplary embodiment of the present disclosure, a gas turbine engine including a turbomachinery core operable to produce a flow of combustion gases and a turbine configured to extract energy from the combustion gases so as to drive a fan is provided. The fan includes a rotor including at least one rotor stage having a rotatable disk defining a flowpath surface and an array of rotor airfoils extending outward from the flowpath surface; an array of splitter airfoils extending outward from the flowpath surface, wherein the splitter airfoils and the rotor airfoils are disposed in a sequential arrangement; and a shroud extending between the rotor airfoils and the splitter airfoils.

In certain exemplary embodiments the shroud is connected to the rotor airfoils at a position radially inward of a tip of each of the rotor airfoils and radially outward of a root of each of the rotor airfoils.

In certain exemplary embodiments the shroud is connected to the splitter airfoils at a position radially inward of a tip of each of the splitter airfoils and radially outward of a root of each of the splitter airfoils.

In certain exemplary embodiments the shroud is connected to a portion of a tip of each of the splitter airfoils.

In certain exemplary embodiments a first portion of the shroud is connected to a first splitter airfoil at a position radially inward of a tip of the first splitter airfoil and radially outward of a root of the first splitter airfoil, and wherein a second portion of the shroud is connected to a second splitter airfoil at a tip of the second splitter airfoil.

In certain exemplary embodiments a third portion of the shroud is disposed relative to a third splitter airfoil at a position adjacent to a tip of the third splitter airfoil.

In certain exemplary embodiments the shroud includes a plurality of linked shroud sections.

In certain exemplary embodiments a chord dimension of the splitter airfoil is less than a chord dimension of the rotor airfoil.

In certain exemplary embodiments a span dimension of the splitter airfoil is less than a span dimension of the rotor airfoil.

DETAILED DESCRIPTION

The terms “forward” and “aft” refer to relative positions within a gas turbine engine, with forward referring to a position closer to an engine inlet and aft referring to a position closer to an engine nozzle or exhaust.

A multi-stage fan splittered rotor of the present disclosure includes an array of rotor airfoils, an array of splitter airfoils, and a shroud extending between the rotor airfoils and the splitter airfoils as described and shown herein. A shroud of the present disclosure improves rotor aerodynamic performance and simultaneously reduces weight of the fan design. For example, the shroud of the present disclosure improves rotor performance, e.g., partial tip shroud allowing corner clearance flow benefits, and reduces full airfoil hub thickness requirements for frequency placement by using the shroud to set modal frequencies to thereby improve performance, hub pressure ratio, weight, and sensitivity to clearances.

Synergistically incorporating the splitter airfoils with the rotor airfoils using the shroud of the present disclosure solves the above-described technical problems and enables (1) a reduced weight design, (2) improved aerodynamic performance and stall range by enabling thinner airfoils and more aggressive levels of rotor leading edge forward sweep, (3) elimination of splitter blade tip vortex, which is another source of total pressure loss, further improving splittered rotor aerodynamic performance, (4) reduced sensitivity to rotor tip clearance by way of splittered rotor airfoil count reduction, and/or (5) supercharging of the fan hub airflow to higher total pressure levels before it travels into the core engine high pressure compressor inlet thereby helping to increase cycle operating pressure ratios.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,FIG. 1is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment ofFIG. 1, the gas turbine engine is a high-bypass turbofan jet engine10, referred to herein as “turbofan engine10.” As shown inFIG. 1, the turbofan engine10defines an axial direction A (extending parallel to a longitudinal centerline or axis12provided for reference) and a radial direction R. In general, the turbofan10includes a fan section14and a turbomachine16disposed downstream from the fan section14.

For the embodiment depicted, the fan section14includes a variable pitch fan38having a plurality of fan blades40coupled to a disk42in a spaced apart manner. As depicted, the fan blades40extend outwardly from disk42generally along the radial direction R. Each fan blade40is rotatable relative to the disk42about a pitch axis P by virtue of the fan blades40being operatively coupled to a suitable actuation member44configured to collectively vary the pitch of the fan blades40in unison. The fan blades40, disk42, and actuation member44are together rotatable about the longitudinal axis12by LP shaft36across a power gear box46. The power gear box46includes a plurality of gears for stepping down the rotational speed of the LP shaft36to a more efficient rotational fan speed.

Referring still to the exemplary embodiment ofFIG. 1, the disk42is covered by rotatable front nacelle48aerodynamically contoured to promote an airflow through the plurality of fan blades40. Additionally, the exemplary fan section14includes an annular fan casing or outer nacelle50that circumferentially surrounds the fan38and/or at least a portion of the turbomachine16. The nacelle50is, for the embodiment depicted, supported relative to the turbomachine16by a plurality of circumferentially-spaced outlet guide vanes52. Additionally, a downstream section54of the nacelle50extends over an outer portion of the turbomachine16so as to define a bypass airflow passage56therebetween.

In an exemplary embodiment of the present disclosure, the fan14may include a number of rotor stages, each of which includes a row of fan blades or rotor airfoils83mounted to a rotor81having a rotatable disk87(FIGS. 2 and 3). The fan14may also include at least one stator stage including a row of stationary or stator airfoils that serve to turn the airflow passing therethrough. As used herein, the term “fan” refers to any apparatus in a turbine engine having a rotor with airfoils operable to produce a fluid flow.

It is contemplated that the principles of the present invention are equally applicable to multi-stage fans, single-stage fans, and other fan configurations; as well as with low-bypass turbofan engines, high-bypass turbofan engines, and other engine configurations.

It should be appreciated, however, that the exemplary turbofan engine10depicted inFIG. 1is by way of example only, and that in other exemplary embodiments, the turbofan engine10may have any other suitable configuration. For example, in other exemplary embodiments, the turbofan engine10may be a direct drive turbofan engine (i.e., not including the power gearbox46), may include a fixed pitch fan38, etc. Additionally, or alternatively, aspects of the present disclosure may be incorporated into any other suitable gas turbine engine, such as a turboshaft engine, turboprop engine, turbojet engine, a land-based gas turbine engine for power generation, an aeroderivative gas turbine engine, etc.

FIGS. 2-13illustrate exemplary embodiments of the present disclosure. Referring toFIGS. 2-5 and 14-16, in an exemplary embodiment, a portion of an exemplary rotor stage80of the present disclosure that is suitable for incorporation in a fan14for an engine10is illustrated. The fan14may include a number of rotor stages, each of which includes a row of fan blades or rotor airfoils83mounted to a rotor81having a rotatable disk87. The rotor stage80of fan14includes a multi-stage fan splittered rotor having an array of rotor airfoils83, an array of splitter airfoils144, and a shroud200extending between the rotor airfoils83and the splitter airfoils144as described in more detail below. In at least certain exemplary embodiments, the exemplary multi-stage fan splittered rotor depicted inFIGS. 2-3may be incorporated into, e.g., the exemplary engine10described above with reference toFIG. 1.

A shroud200of the present disclosure improves rotor aerodynamic performance and simultaneously reduces weight of the fan design. For example, the shroud200of the present disclosure improves rotor performance, e.g., partial tip shroud allowing corner clearance flow benefits, and reduces full airfoil hub thickness requirements for frequency placement by using the shroud200to set modal frequencies to thereby improve performance, hub pressure ratio, weight, and sensitivity to clearances.

Synergistically incorporating the splitter airfoils144with the rotor airfoils83using the shroud200solves the above-described technical problems and enables (1) a reduced weight design, (2) improved aerodynamic performance and stall range by enabling thinner airfoils and more aggressive levels of rotor leading edge forward sweep, (3) elimination of splitter blade tip vortex, which is another source of total pressure loss, further improving splittered rotor aerodynamic performance, (4) reduced sensitivity to rotor tip clearance by way of splittered rotor airfoil count reduction, and/or (5) supercharging of the fan hub airflow to higher total pressure levels before it travels into the core engine high pressure compressor inlet thereby helping to increase cycle operating pressure ratios.

The shroud200of the present disclosure extends between and connects the rotor airfoils83and the splitter airfoils144. In this manner, the shroud200enables a lower maximum thickness-to-chord ratio in the hub region of the main blade or airfoil83, which both reduces the weight of the “bladed disk” or “blisk” and improves the aerodynamic performance of the main blade or airfoil83. The shroud200also improves splitter blade or airfoil performance by the removal of the tip vortex that is present without the shroud200. In an exemplary embodiment, the airfoil performance of the present disclosure may include a rotor tip stagger of approximately 60 degrees to approximately 65 degrees. In an exemplary embodiment, the airfoil performance of the present disclosure may include a flow coefficient of approximately 0.3 to approximately 0.6. In one embodiment, the airfoil performance of the present disclosure may include a tip radius of approximately 19 inches to approximately 22 inches. In an exemplary embodiment, the airfoil performance of the present disclosure may include a tip speed of approximately 1400 ft/sec to approximately 1650 ft/sec.

Referring toFIGS. 2-5 and 14-16, in one embodiment, an exemplary rotor stage80of a fan14may include a number of rotor stages, each of which includes a row of fan blades or rotor airfoils83mounted to a rotor81having a rotatable disk87. The fan14may also include at least one stator stage including a row of stationary or stator airfoils that serve to turn the airflow passing therethrough.

The rotor81defines an annular flowpath surface82, a portion of which is shown inFIG. 2. Referring toFIGS. 2 and 6, in one embodiment, the flowpath surface82may have an axisymmetric surface profile. Referring toFIG. 14, in other embodiments, the flowpath surface may have a non-axisymmetric surface profile182. For example, referring toFIG. 14, a non-axisymmetric surface profile may be contoured with a concave curve or “scallop”184between each adjacent pair of rotor airfoils83. For comparison purposes, the dashed lines inFIG. 14illustrate a hypothetical axisymmetric cylindrical surface with a radius passing through the roots84of the rotor airfoils83. It can be seen that the non-axisymmetric flowpath surface182curvature has its maximum radius (or minimum radial depth “d” of the scallop184) at the roots84of the rotor airfoils83, and has its minimum radius (or maximum radial depth “d” of the scallop184) at a position approximately midway between adjacent rotor airfoils83, e.g., at the root154of a splitter airfoil144.

In steady state or transient operation, this scalloped configuration (FIG. 14) is effective to reduce the magnitude of mechanical and thermal hoop stress concentration at the airfoil hub intersections on the rim96of the rotor81along the flowpath surface182. This contributes to the goal of achieving acceptably-long component life of the disk87and the rotor81.

Referring toFIGS. 2, 3, 6, 7, and 14-16, the rotor81includes a plurality or array of rotor airfoils83extending outward from the flowpath surface82. Each fan blade or rotor airfoil83extends from a root84at the flowpath surface82to a tip86and includes a concave pressure side88joined to a convex suction side90at a leading edge92and a trailing edge94. Referring toFIG. 4, each rotor airfoil83has a span, or span dimension, “S1” defined as the radial distance from the root84to the tip86, and a chord, or chord dimension, “C1” defined as the length of an imaginary straight line connecting the leading edge92and the trailing edge94. Depending on the specific design of the rotor airfoil83, its chord C1may be different at different locations along the span S1. In one embodiment, a relevant measurement is the chord C1at the root84of the rotor airfoil83.

Referring toFIGS. 2, 3, 6, 7, and 14-16, in one embodiment, the rotor airfoils83are uniformly spaced apart around the periphery of the flowpath surface82. Referring toFIG. 15, in one embodiment, a mean circumferential spacing “s” between adjacent rotor airfoils83is defined as s=2πr/Z, where “r” is a designated radius of the rotor airfoils83(for example at the root84) and “Z” is the number of rotor airfoils83. A non-dimensional parameter called “solidity” is defined as c/s, where “c” is equal to the blade chord as described above. In embodiments of the present disclosure, the rotor airfoils83may have a spacing which is significantly greater than a spacing that would be expected in the prior art, resulting in a blade solidity significantly less than would be expected in the prior art. This reduced solidity can minimize efficiency losses of the fan14and this reduced blade solidity can also have the effect of reducing weight and simplifying manufacturing by minimizing the total number of fan airfoils used in a given rotor.

An aerodynamically adverse side effect of reduced blade solidity is to increase the rotor passage flow area between adjacent rotor airfoils83. This increase in rotor passage through flow area increases the aerodynamic loading level and in turn tends to cause undesirable flow separation on the suction side90of the rotor airfoil83, at the inboard portion near the root84, also referred to as “hub flow separation”.

To reduce or prevent hub flow separation, the rotor stage80may be provided with splitters, or in a “splittered” configuration. Referring toFIGS. 2, 3, 6, 7, and14-16, a plurality or an array of splitter blades or airfoils144extend outward from the flowpath surface82. In an exemplary embodiment, one splitter airfoil144is disposed between each pair of adjacent rotor airfoils83. For example, each splitter airfoil144may be located approximately midway between two adjacent rotor airfoils83. In some embodiments, the splitter airfoils144and the rotor airfoils83are disposed in a sequential arrangement. In some exemplary embodiments, the splitter airfoils144and the rotor airfoils83are disposed in a staggered and alternating arrangement. It is contemplated that the splitter airfoils144and the rotor airfoils83may be arranged in other sequential configurations for a variety of different applications.

Referring toFIGS. 2, 3, 6, 7, and 14-16, in the circumferential direction, the splitter airfoils144may be located halfway or circumferentially biased between two adjacent rotor airfoils83. In other words, the rotor airfoils83and splitter airfoils144alternate around the periphery of the flowpath surface82.

Referring toFIGS. 2, 3, 5, and 14-16, each splitter airfoil144extends from a root154at the flowpath surface82to a tip156, and includes a concave pressure side158joined to a convex suction side160at a leading edge162and a trailing edge164. Referring toFIG. 5, in one embodiment, each splitter airfoil144has a span, or span dimension, “S2” defined as the radial distance from the root154to the tip156, and a chord, or chord dimension, “C2” defined as the length of an imaginary straight line connecting the leading edge162and the trailing edge164. Depending on the specific design of the splitter airfoil144, its chord C2may be different at different locations along the span S2. In one embodiment, a relevant measurement is the chord C2at the root154of the splitter airfoil144.

Referring toFIGS. 4 and 5, in an exemplary embodiment, a chord dimension C2of the splitter airfoil144is less than a chord dimension C1of the rotor airfoil83. Furthermore, in an exemplary embodiment, a span dimension S2of the splitter airfoil144is less than a span dimension S1of the rotor airfoil83. For example, in some embodiments, the span S2and/or the chord C2of the splitter airfoils144may be some fraction less than the corresponding span S1and chord C1of the rotor airfoils83. These may be referred to as “part-span” and/or “part-chord” splitter airfoils. For example, in some embodiments, the span S2may be equal to or less than the span S1. In some embodiments, for reducing frictional losses, the span S2is 50% or less of the span S1. In other embodiments, for the least frictional losses, the span S2is 30% or less of the span S1. In some exemplary embodiments, the chord C2may be equal to or less than the chord C1. In other exemplary embodiments, for the least frictional losses, the chord C2is 80% or less of the chord C1. In one embodiment, the length of the chord C2of the splitter airfoil144is approximately 25% to approximately 75% of the length of the chord C1of the rotor airfoil83. In one embodiment, the height of span S2of the splitter airfoil144is approximately 20% to approximately 50% of the span S1of the rotor airfoil83.

The splitter airfoils144function to locally increase the hub solidity of the rotor stage80and thereby prevent the above-mentioned flow separation from the rotor airfoils83. A similar effect could be obtained by simply increasing the number of rotor airfoils83, and therefore reducing the blade-to-blade spacing. This, however, has the undesirable side effect of increasing aerodynamic surface area frictional losses which would manifest as reduced aerodynamic efficiency and increased rotor weight. Therefore, the dimensions of the splitter airfoils144and their position may be selected to prevent flow separation while minimizing their surface area.

Referring toFIGS. 7, 9, and 16, in one embodiment, the splitter airfoils144are positioned so that their trailing edges164are at approximately the same axial position as the trailing edges94of the rotor airfoils83, relative to the rotor81. It is contemplated that the splitter airfoils144and the rotor airfoils83may be arranged in other configurations for a variety of different applications.

Advantageously, referring toFIGS. 2, 3, and 6-16, a fan rotor81of the present disclosure includes a shroud200extending between the rotor airfoils83and the splitter airfoils144. In this manner, the shroud200of the present disclosure improves rotor aerodynamic performance and simultaneously reduces weight of the fan design. For example, the shroud200of the present disclosure improves rotor performance, e.g., partial tip shroud allowing corner clearance flow benefits, and reduces full airfoil hub thickness requirements for frequency placement by using the shroud200to set modal frequencies to thereby improve performance, hub pressure ratio, weight, and sensitivity to clearances.

Synergistically incorporating the splitter airfoils144with the rotor airfoils83using the shroud200solves the above-described technical problems and enables (1) a reduced weight design, (2) improved aerodynamic performance and stall range by enabling thinner airfoils and more aggressive levels of rotor leading edge forward sweep, (3) elimination of splitter blade tip vortex, which is another source of total pressure loss, further improving splittered rotor aerodynamic performance, (4) reduced sensitivity to rotor tip clearance by way of splittered rotor airfoil count reduction, and/or (5) supercharging of the fan hub airflow to higher total pressure levels before it travels into the core engine high pressure compressor inlet thereby helping to increase cycle operating pressure ratios.

In an exemplary embodiment, the shroud200of the present disclosure extends between and connects the rotor airfoils83and the splitter airfoils144. In this manner, the shroud200enables a lower maximum thickness-to-chord ratio in the hub region of the main blade or airfoil83, which both reduces the weight of the “bladed disk” or “blisk” and improves the aerodynamic performance of the main blade or airfoil83. The shroud200also improves splitter blade or airfoil performance by the removal of the tip vortex that is present without the shroud200.

Referring toFIGS. 2-3 and 6-10, in an exemplary embodiment, a shroud200of the present disclosure includes a leading edge212, a trailing edge214, a shroud superior surface216extending between the leading edge212and the trailing edge214, and an opposing shroud inferior surface218extending between the leading edge212and the trailing edge214. The shroud superior surface216and the shroud inferior surface218provide airfoil surfaces for guiding and controlling a fluid flow.

Referring toFIGS. 2-3 and 6-10, in one embodiment, the shroud200extends between and connects the rotor airfoils83and the splitter airfoils144. For example, the shroud200is connected to the rotor airfoils83at a position radially inward of a tip86of each of the rotor airfoils83and radially outward of a root84of each of the rotor airfoils83. Furthermore, in this embodiment, the shroud200is connected to a portion of a tip156of each of the splitter airfoils144.

In an exemplary embodiment, by incorporating the splitter airfoils144under a shroud200disclosed herein, a system of the present disclosure enables supercharging the fan hub flow before traveling down the core into the high pressure compressor; reduces and/or eliminates the splitter tip vortex (another source of loss) thereby improving splitter performance; improving rotor aerodynamic performance; and simultaneously reducing the weight.

Referring toFIGS. 2, 3, 6, and 7, in an exemplary embodiment, the shroud200includes a plurality of linked shroud sections. For example, the shroud200may include a first shroud section202and a second shroud section204.

Referring toFIG. 11, in another exemplary embodiment, the shroud200extends between and connects the rotor airfoils83and the splitter airfoils144. For example, the shroud200is connected to the rotor airfoils83at a position radially inward of a tip86(FIG. 4) of each of the rotor airfoils83and radially outward of a root84(FIG. 4) of each of the rotor airfoils83. Furthermore, in this embodiment, the shroud200is connected to the splitter airfoils144at a position radially inward of a tip156(FIG. 5) of each of the splitter airfoils144and radially outward of a root154(FIG. 5) of each of the splitter airfoils144. In the embodiment shown inFIG. 11, the shroud extends inward of the tip156of the splitter airfoil144. It is contemplated that in other embodiments, the shroud200may extend further inward of the tip156of the splitter airfoil144towards the root154of the splitter airfoil144for other desired applications.

Referring toFIG. 13, in another exemplary embodiment, the shroud200extends between and connects the rotor airfoils83and the splitter airfoils144. For example, the shroud200is connected to the rotor airfoils83at a position radially inward of a tip86of each of the rotor airfoils83and radially outward of a root84of each of the rotor airfoils83. Furthermore, in other exemplary embodiments, different portions of the shroud200may be connected to different portions of the respective splitter airfoils144. For example, a first portion206of the shroud200may be connected to a first splitter airfoil170in a first configuration, e.g., at a position radially inward of a tip156of the first splitter airfoil170and radially outward of a root154of the first splitter airfoil170, as shown inFIG. 11. Furthermore, in an exemplary embodiment, a second portion208of the shroud200may be connected to a second splitter airfoil172in a second configuration, e.g., at a tip156of the second splitter airfoil172, as shown inFIGS. 2-3 and 6-7. In other words, different sections of the shroud200may be connected to different portions and/or in different configurations with the respective splitter airfoils144,170,172. Referring toFIGS. 12 and 13, in some embodiments, a third portion210of the shroud200may be in a third configuration, e.g., the shroud200may be disposed relative to a third splitter airfoil174at a position adjacent to a tip156of the third splitter airfoil174.

In an exemplary embodiment, components of the rotor81of the fan14, e.g., the disk87, the rotor airfoils83, and the splitter airfoils144, may be constructed from any material capable of withstanding the anticipated stresses and environmental conditions in operation. Non-limiting examples of known suitable alloys include iron, nickel, and titanium alloys. In some embodiments, the disk87, the rotor airfoils83, and the splitter airfoils144are depicted as an integral, unitary, or monolithic whole. This type of structure may be referred to as a “bladed disk” or “blisk”. The principles of the present invention are equally applicable to a rotor built up from separate components.

1. A fan for a gas turbine engine, the fan comprising: a rotor comprising at least one rotor stage including a rotatable disk defining a flowpath surface and an array of rotor airfoils extending outward from the flowpath surface; an array of splitter airfoils extending outward from the flowpath surface, wherein the splitter airfoils and the rotor airfoils are disposed in a sequential arrangement; and a shroud extending between the rotor airfoils and the splitter airfoils.

2. The fan of any preceding clause, wherein the shroud is connected to the rotor airfoils at a position radially inward of a tip of each of the rotor airfoils and radially outward of a root of each of the rotor airfoils.

3. The fan of any preceding clause, wherein the shroud is connected to the splitter airfoils at a position radially inward of a tip of each of the splitter airfoils and radially outward of a root of each of the splitter airfoils.

4. The fan of any preceding clause, wherein the shroud is connected to a portion of a tip of each of the splitter airfoils.

5. The fan of any preceding clause, wherein a first portion of the shroud is connected to a first splitter airfoil at a position radially inward of a tip of the first splitter airfoil and radially outward of a root of the first splitter airfoil, and wherein a second portion of the shroud is connected to a second splitter airfoil at a tip of the second splitter airfoil.

6. The fan of any preceding clause, wherein a third portion of the shroud is disposed relative to a third splitter airfoil at a position adjacent to a tip of the third splitter airfoil.

7. The fan of any preceding clause, wherein the shroud includes a plurality of linked shroud sections.

8. The fan of any preceding clause, wherein a chord dimension of the splitter airfoil is less than a chord dimension of the rotor airfoil.

9. The fan of any preceding clause, wherein a span dimension of the splitter airfoil is less than a span dimension of the rotor airfoil.

10. The fan of any preceding clause, wherein the splitter airfoils and the rotor airfoils are disposed in a staggered and alternating arrangement.

11. The fan of any preceding clause, wherein each splitter airfoil is located approximately midway between two adjacent rotor airfoils.

12. A gas turbine engine, comprising: a turbomachinery core operable to produce a flow of combustion gases; a turbine configured to extract energy from the combustion gases so as to drive a fan, wherein the fan includes: a rotor comprising at least one rotor stage including a rotatable disk defining a flowpath surface and an array of rotor airfoils extending outward from the flowpath surface; an array of splitter airfoils extending outward from the flowpath surface, wherein the splitter airfoils and the rotor airfoils are disposed in a sequential arrangement; and a shroud extending between the rotor airfoils and the splitter airfoils.

13. The gas turbine engine of any preceding clause, wherein the shroud is connected to the rotor airfoils at a position radially inward of a tip of each of the rotor airfoils and radially outward of a root of each of the rotor airfoils.

14. The gas turbine engine of any preceding clause, wherein the shroud is connected to the splitter airfoils at a position radially inward of a tip of each of the splitter airfoils and radially outward of a root of each of the splitter airfoils.

15. The gas turbine engine of any preceding clause, wherein the shroud is connected to a portion of a tip of each of the splitter airfoils.

16. The gas turbine engine of any preceding clause, wherein a first portion of the shroud is connected to a first splitter airfoil at a position radially inward of a tip of the first splitter airfoil and radially outward of a root of the first splitter airfoil, and wherein a second portion of the shroud is connected to a second splitter airfoil at a tip of the second splitter airfoil.

17. The gas turbine engine of any preceding clause, wherein a third portion of the shroud is disposed relative to a third splitter airfoil at a position adjacent to a tip of the third splitter airfoil.

18. The gas turbine engine of any preceding clause, wherein the shroud includes a plurality of linked shroud sections.

19. The gas turbine engine of any preceding clause, wherein a chord dimension of the splitter airfoil is less than a chord dimension of the rotor airfoil.

20. The gas turbine engine of any preceding clause, wherein a span dimension of the splitter airfoil is less than a span dimension of the rotor airfoil.