AIRFOIL FOR A ROTARY MACHINE INCLUDING A PROPELLOR ASSEMBLY

An airfoil includes a proximal end and a distal end opposite the proximal end. The airfoil also includes a distal portion extending adjacent the distal end and an edge extending between the proximal end and the distal end. The airfoil further includes a surface extending between the proximal end and the distal end. The edge and the surface define a sweep and a cahedral through the distal portion and at least a portion of the intermediate portion.

BACKGROUND

The field of the disclosure relates generally to airfoils and, more particularly, to airfoils for rotary machines that include a propeller assembly.

At least some known rotary machines are gas turbine engines used for propelling an aircraft in flight. Some known gas turbine engines include a combustor, a compressor coupled upstream from the combustor, a turbine, a rotor assembly rotatably coupled between the compressor and the turbine. At least some known aircraft engines include at least one airfoil that is moved through fluid to produce an aerodynamic force. For example, some known rotary machines include a propeller assembly drivingly coupled to the rotor assembly. The rotor assembly rotates to drive the propeller assembly and, thereby, move air to propel the aircraft. Rotation of the propeller assembly also generates tip flow structures, which cause noise and aerodynamic inefficiencies.

To increase an operating efficiency, some known rotary machines include winglets on blades of the propeller assemblies to inhibit air flowing over the tip of the blade and, thereby reduce the generation of flow structures. The winglets project from each blade at or very near the tip of the blade and form a change in direction along a surface of the blade. However, the change in direction sometimes generates flow patterns that increase a noise level of the rotary machine during operation.

Some known aircraft engines include two or more propeller assemblies that interact to move air and, thereby, propel an aircraft. The interaction of the propeller assemblies generates flow structures, which cause operating inefficiencies. Accordingly, some aircraft engines include fences attached to the blades of the forward propeller assembly to reduce the operating inefficiencies caused by the interaction of the propeller assemblies. However, the two or more propeller assemblies are not configured for use in rotary machines having a single propeller assembly, such as unducted single fan propulsion systems and turboprop engines.

BRIEF DESCRIPTION

In one aspect, an airfoil includes a proximal end and a distal end opposite the proximal end. The airfoil also includes a distal portion extending adjacent the distal end and an edge extending between the proximal end and the distal end. The airfoil further includes a surface extending between the proximal end and the distal end. The edge and the surface define a sweep and a cahedral through the distal portion and at least a portion of the intermediate portion.

In another aspect, a rotary machine includes at least one rotatable member and a casing extending at least partly circumferentially around the at least one rotatable element. A propeller assembly is adjacent the casing and drivingly coupled to the at least one rotatable member. The propeller assembly includes a hub and at least one blade extending radially from the hub. The at least one blade includes a proximal end adjacent the hub and a distal end opposite the proximal end. An intermediate portion, an edge, and a surface extend between the proximal end and the distal end. A distal portion extends between the intermediate portion and the distal end. The edge and the surface define a sweep and a cahedral through the distal portion and at least a portion of the intermediate portion.

In yet another aspect, a method of operating a rotary machine is provided. The rotary machine includes at least one rotatable member and a propeller assembly drivingly coupled to the at least one rotatable member. The propeller assembly includes a hub and at least one blade extending radially from the hub. The method includes rotating the at least one rotatable member such that the propeller assembly rotates. The method further includes contacting air with the at least one blade. The at least one blade includes a leading edge and a surface. The leading edge and the surface define a sweep and a cahedral of the at least one blade. The method further includes directing air along the leading edge and the surface defining the sweep and the cahedral such that a flow pattern is generated adjacent the tip. The cahedral and the sweep of the at least one blade are configured to reduce noise generated by the flow pattern.

DETAILED DESCRIPTION

As used herein, the terms “axial” and “axially” refer to directions and orientations extending substantially parallel to a longitudinal axis of a rotary machine. The terms “radial” and “radially” refer to directions and orientations extending substantially perpendicular to the longitudinal axis of the rotary machine. In addition, as used herein, the terms “circumferential” and “circumferentially” refer to directions and orientations extending arcuately about the longitudinal axis of the rotary machine. Moreover, “upstream” refers to a forward end of the rotary machine, and “downstream” refers to an aft end of the rotary machine.

As used herein, the term “cahedral” refers to an angle of a portion of a blade in relation to the radial axis of the blade. The term “cahedral” is a generic term used to refer to both dihedral and anhedral. As used herein, the term “dihedral” refers to an angle of a portion of a blade in relation to the radial axis where the blade portion extends on a low pressure side of the blade. The term “anhedral” refers to an angle of a portion of a blade in relation to the radial axis where the blade portion extends on a high pressure side of the blade. In addition, as used herein, the term “sweep” refers to an angle of a curve of a blade in relation to the radial axis.

The rotary machine systems described herein decrease sound generated by the rotary machine systems by including an airfoil having a cahedral and a sweep. The cahedral and the sweep reduce noise generated by the rotary machine systems as the airfoil moves through air. In some embodiments, the cahedral and the sweep extend throughout a distal portion of each airfoil such that the rotary machine systems generate less noise during operation than known rotary machine systems. Moreover, some embodiments described herein provide for increased efficiency of the rotary machine systems.

Although generally described herein with respect to an aircraft gas turbine engine, the methods and systems described are applicable to any systems that include an airfoil, for example, without limitation, turbojets, turbofans, propellers, unducted fans driven by reciprocating engines, electric motors, wind turbines, and other systems that require an airfoil. In some embodiments, the systems described herein include, without limitation, any of the following airfoils: an outlet guide vane, a fan blade, a rotor blade, a stator vane, a ducted fan blade, an unducted fan blade, a strut, a nacelle inlet, a wind turbine blade, a propeller, an impeller, a diffuser vane, a return channel vane, flap leading edges, wing leading edges, landing gear fairings, a marine propeller, and a pylon.

FIG. 1is a sectional schematic view of an exemplary rotary machine. In the exemplary embodiment, the rotary machine is a gas turbine engine, indicated generally by the reference number100. Alternatively, the rotary machine is any other turbine engine and/or rotary machine, including, without limitation, a steam turbine engine, a centrifugal compressor, and a turbocharger. In the exemplary embodiment, gas turbine engine100includes a drive102and a drive shaft104extending from drive102. During operation of gas turbine engine100, drive102rotates drive shaft104to cause movement of a load connected to drive shaft104. In some embodiments, drive102includes, in serial flow relationship, a compressor (not shown), a combustor (not shown) downstream from compressor, and a turbine (not shown) downstream from combustor. In alternative embodiments, drive102includes any components that enable gas turbine engine100to operate as described herein. A casing112at least partly surrounds drive102and at least partly defines an airway for air116to flow through gas turbine engine100. Casing112includes an inlet118and an outlet (not shown) downstream from inlet118. In the exemplary embodiment, gas turbine engine100is generally disposed about an engine centerline122.

In the exemplary embodiment, gas turbine engine100further includes a propeller assembly124. Propeller assembly124includes a hub126and a plurality of blades128extending radially from hub126. Hub126is coupled to a propeller shaft130, which is drivingly coupled to drive shaft104by a drive gear131such that rotation of drive shaft104induces rotation of propeller assembly124. Each of blades128have a proximal end132adjacent and coupled to hub126and a distal end134, i.e., tip, opposite proximal end132. Propeller assembly124has a propeller diameter137that is measured substantially perpendicular to engine centerline122between distal ends134of blades128. As will be described in more detail below, each blade128includes a cahedral136defined through a distal portion138of each blade128. Each blade128further includes a sweep140(shown inFIG. 3) defined through at least distal portion138. As a result, blade128reduces noise generated during operation of gas turbine engine100. For example, in some embodiments, flow structures190in air116along blades128are dispersed into a flow pattern that reduces noise generated by gas turbine engine100.

A high-pressure surface154and an opposed low-pressure surface152extend between proximal end132and distal end134. A leading edge156and a trailing edge158also extend between proximal end132and distal end134. A thickness of blade128is defined between high-pressure surface154and low-pressure surface152. In alternative embodiments, blade128has any surfaces and edges that enable propeller assembly124to operate as described herein.

During operation, propeller assembly124rotates such that air116flows in a generally axial direction of propeller assembly124during operation of gas turbine engine100. However, a portion of air116flowing along blade128forms flow structures190. The shape and configuration of blade128affects flow structures190generated in air116flowing along blade128. In the exemplary embodiment, blade128has cahedral136and sweep140configured to reduce noise generated by blade128. In alternative embodiments, blade128has any configuration that enables blade128to operate as described herein.

FIG. 2Ais a stacking line129of blade128.FIG. 2Bis an alternative stacking line131of blade128. Stacking line129and stacking line131extend in the radial direction along the span of blade128. Blade128has distal portion138which includes portions of leading edge156, trailing edge158, high-pressure surface154(shown inFIG. 4), and low-pressure surface152(shown inFIG. 4). Distal portion138extends from an intermediate portion166of blade128to distal end134. Distal portion138is configured, at least in part, to reduce noise generated in air116flowing along blade128. For example, in the exemplary embodiment, distal portion138defines cahedral136of blade128. In the exemplary embodiment, high-pressure surface154(shown inFIG. 4) and low-pressure surface152(shown inFIG. 4) slope away from a radial axis168to define cahedral136. In addition, leading edge156(shown inFIG. 1) and trailing edge158(shown inFIG. 1) curve to at least partially define cahedral136. In alternative embodiments, cahedral136is defined by any surfaces and edges of blade128that enable propeller assembly124to operate as described herein.

In the exemplary embodiment, high-pressure surface154(shown inFIG. 4) and low-pressure surface152(shown inFIG. 4) extend on low-pressure side162of blade128such that cahedral136is a dihedral170. In alternative embodiments, any surfaces of blade128extend on any sides of blade128that enable blade128to operate as described herein. For example, in some embodiments, at least one of high-pressure surface154and low-pressure surface152(shown inFIG. 4) slopes away from radial axis168and extends on high-pressure side164to form an anhedral171. In further embodiments, any portions of blade128have any cahedral that enable blade128to operate as described herein. In the exemplary embodiment, a cahedral axis172extends through distal portion138and defines cahedral136in relation to radial axis168. In some embodiments, cahedral136is in a range extending from about 1° to about 180°. In alternative embodiments, cahedral136is any measurement that enables the blade128to operate as described herein. For example, in some embodiments, cahedral136is in a range extending from about −1° to about −180°.

In some embodiments, blade128has a varying cahedral136throughout the length of blade128. For example, in some embodiments, blade128has a dihedral170through at least a portion of distal portion138and anhedral171adjacent hub126to balance the loading on blade128. Accordingly, at least one of high-pressure surface154(shown inFIG. 4) and low-pressure surface152(shown inFIG. 4) crosses radial axis168to form opposed dihedral170and anhedral171. In alternative embodiments, blade128has any constant or varying cahedral136that enables blade128to operate as described herein. In further embodiments, the twist and/or camber of blade128are configured to balance the loading of blade128. In particular, the twist and/or camber are designed to compensate for the aerodynamic loading profile changes introduced by the changes to cahedral136and sweep140. InFIG. 3, blade128has an aft sweep140. In alternative embodiments, blade128has any sweep140that enables blade128to operate as described. For example, in some embodiments, blade128has an at least partially forward sweep140.

The camber is adjusted in span and chord directions of blade128and the twist is adjusted along the span of blade128. As a result, blade128is configured, for example, to reduce adverse effects of a high flow acceleration around leading edge156(shown inFIG. 1). In alternative embodiments, blade128has any camber and twist that enable propeller assembly124to operate as described herein.

In some embodiments, cahedral136and sweep140through distal portion138extends for a portion of the span of blade128. Substantially the remaining span of blade128is configured to balance cahedral136and sweep140. In particular, in some embodiments, the remaining span is shifted relative to radial axis168, i.e., restacked. For example, in some embodiments, distal portion138having cahedral136and sweep140extends approximately 30% of the span of blade128. The remaining portion, covering 70% of blade128, is restacked to offset the mechanical loading stresses associated with cahedral136and sweep140shifting blade128out of balance relative to a baseline blade design. In alternative embodiments, blade128has any restacking that enable propeller assembly124to operate as described herein. In some embodiments, the camber of blade128is adjusted through a portion of the span, i.e., blade128is recambered. In further embodiments, the twist of blade is adjusted through a portion of the span, i.e., blade128is retwisted.

Also, in the exemplary embodiment, high-pressure surface154and low-pressure surface152are angled throughout distal portion138. In alternative embodiments, high-pressure surface154and low-pressure surface152are angled through any portions of blade128that enable blade128to operate as described herein. In the exemplary embodiment, high-pressure surface154and low-pressure surface152are substantially parallel throughout distal portion138such that the thickness of blade128is substantially constant throughout distal portion138. The slope of high-pressure surface154in relation to radial axis168increases from a lesser slope at intermediate portion166to a greater slope at distal end134. Likewise, the slope of low-pressure surface152in relation to radial axis168increases from a lesser slope at intermediate portion166to a greater slope at distal end134. In alternative embodiments, high-pressure surface154and low-pressure surface152have any slopes that enable blade128to operate as described herein. For example, in some embodiments, at least one of high-pressure surface154and low-pressure surface152has a varying slope throughout distal portion138. In further embodiments, at least one of high-pressure surface154and low-pressure surface152has a portion that is substantially parallel to radial axis168.

In the exemplary embodiment, blade128has a length174defined between distal end134and proximal end132of blade128along radial axis168. Length174is any measurement that enables blade128to operate as described herein. In the exemplary embodiment, distal portion138extends a percentage of length174of blade128that facilitates the reduction of flow structures190generated in air116flowing along blade128. In some embodiments, distal portion138extends a percentage of length174greater than approximately 1%. In alternative embodiments, distal portion138extends any percentage of length174that enables blade128to operate as described herein.

FIG. 3is a side view of propeller blade128. High-pressure surface154, low-pressure surface152, leading edge156, and trailing edge158define sweep140of blade128. In particular, leading edge156defines a leading edge sweep and trailing edge158defines a trailing edge sweep. In the exemplary embodiment, blade128has a sweep140that is greater than the sweep of at least some known blades. Accordingly, sweep140facilitates blade128reducing sound and increasing aerodynamic efficiency during operation of gas turbine engine100in comparison to known systems. In the exemplary embodiment, sweep140is an aft sweep. In alternative embodiments, blade128has any sweep140that enables propeller assembly124to operate as described herein. For example, in some embodiments, sweep140is a forward sweep. In the exemplary embodiment, sweep140varies along leading edge156and trailing edge158. In alternative embodiments, leading edge156and trailing edge158define any varying and constant sweeps140that enable blade128to operate as described herein. In the exemplary embodiment, leading edge156and trailing edge158define sweeps140that increase throughout distal portion138. A leading edge sweep axis180is tangential to leading edge156at distal end134and makes a leading edge sweep angle182along the flow direction with radial axis168. In the exemplary embodiment, a trailing edge sweep axis184is tangential to trailing edge158and makes a trailing edge sweep angle186along the flow direction with radial axis168.

In the exemplary embodiment, cahedral136and sweep140of blade128work in conjunction to increase operating efficiency and reduce noise generated by propeller assembly124. In particular, cahedral136and sweep140affect loading on blade128and the camber and twist of blade128is redesigned to counteract changes to loading distribution to maximize noise benefits of sweep140and cahedral136. Otherwise, improperly designed blades including cahedral and sweep can increase noise generated during operation of propeller assemblies. However, as described herein, blades128including cahedral136and sweep140reduce noise generated during operation of propeller assembly124. In some embodiments, cahedral136is a suction side dihedral170and sweep140is an aft sweep to facilitate cahedral136and sweep140working in conjunction. In further embodiments, cahedral136is a pressure side anhedral171and sweep140is a forward sweep to facilitate cahedral136and sweep140working in conjunction. In alternative embodiments, blade128includes any combinations of cahedral136and sweep140that enable propeller assembly124to operate as described herein.

FIG. 4is a sectional view of blade128. Blade128further includes a chord line192, a mean camber line194, a first axis196, and a second axis198. Chord line192and mean camber line194extend between leading edge156and trailing edge158. Along the span of blade128, cahedral136(shown inFIG. 2A) is defined in a first direction197perpendicular to chord line192and sweep140(shown inFIG. 3) is defined in a direction along chord line192. In alternative embodiments, blade128has any chord line192and mean camber line194that enable propeller assembly124to operate as described herein.

FIG. 5is a sectional schematic view of an embodiment of a rotary machine in the form of a gas turbine engine, indicated generally by the reference number200. In the exemplary embodiment ofFIG. 5, gas turbine engine200is an unducted single fan turbine engine. Unducted single fan turbine engines are described in detail in U.S. Patent Application Publication No. 2015/0284070, which is incorporated by reference herein in its entirety. In the exemplary embodiment, gas turbine engine200includes a drive202and a propeller assembly208. A housing210at least partially encloses drive202. Propeller assembly208includes a shaft218and a plurality of blades220coupled to, and extending radially from, shaft218. Blades220extend at least partially exterior of housing210. A stationary vane204is coupled to housing210and extends on exterior of housing210downstream from blades220. Drive202is drivingly coupled to shaft218by a drive shaft222. During operation, drive202transfers rotational power to propeller assembly208via shaft218and drive shaft222. Propeller assembly208directs air224to move in a flow direction. As air224is moved by propeller assembly208, air224interacts with blades220and stationary vane204, generating noise. In some embodiments, blades220are configured to reduce noise generated. In particular, sweep and cahedral of blades220are configured as described in more detail below to cause less propeller self-noise and interaction noise with stationary objects, such as stationary vane204. In some embodiments, stationary vane204includes sweep and cahedral as described herein to further facilitate reduction of generated noise and increase the operating efficiency of gas turbine engine200. In alternative embodiments, gas turbine engine200includes any vanes that enable gas turbine engine200to operate as described herein. For example, in some embodiments, stationary vane204is disposed upstream of blades220.

In the exemplary embodiment, blade220has a length240defined between distal end232and proximal end230. Length240is any measurement that enables blade220to operate as described herein. In the exemplary embodiment, distal portion238extends a percentage of length240to reduce noise generated in air116(shown inFIG. 1) flowing along distal portion238. In some embodiments, distal portion238extends a percentage of length240greater than approximately 1%. In the exemplary embodiment, distal portion238extends approximately 20% of length240. In alternative embodiments, distal portion238extends any percentage of length240that enables blade220to operate as described herein.

In the exemplary embodiment, leading edge234and trailing edge236define a sweep of blade220. In the exemplary embodiment, sweep is an aft sweep. In alternative embodiments, blade220has any sweep that enables propeller assembly208to operate as described herein. In the exemplary embodiment, leading edge234and trailing edge236define sweeps that decreases to the intermediate portion and increase from the intermediate portion to distal end232. A sweep axis242is parallel to both leading edge234and trailing edge236at distal end232and defines a sweep angle244with a radial axis246of blade220. In some embodiments, sweep angle244is in a range extending from about 1° to about 90°. In further embodiments, sweep angle244is in a range extending from about 5° to about 70°. In alternative embodiments, sweep angle244is any measurement that enables the blade220to operate as described herein.

FIG. 7is a perspective view of an airfoil300.FIG. 8is a sectional view of airfoil300. Airfoil300is superimposed on a baseline airfoil302. In the exemplary embodiment, airfoil300is rotated relative to a longitudinal axis304of baseline airfoil302. In some embodiments, airfoil300rotated by any angle that enables airfoil300to operate as described herein. In the exemplary embodiment, tip loading of airfoil300from airflow306is reduced at least partly due to the shift of airfoil300.

FIG. 9is a perspective view of an airfoil400.FIG. 10is a sectional view of airfoil400. Airfoil400includes a mean camber line402extending through a leading edge404and a trailing edge406. Airfoil400is superimposed on a baseline airfoil408, which has a mean camber line410. Airfoil400is shifted in relation to baseline airfoil408such that mean camber line402and mean camber line410make an angle θ. In some embodiments, angle θ is any angle that enables airfoil400to operate as described herein. In the exemplary embodiment, acceleration of airflow412around leading edge404is reduced at least partly due to the shift of airfoil400, which reduces the drag of airfoil400and noise generated by airflow412. In some embodiments, airfoil400includes sweep140(shown inFIG. 3) and dihedral170(shown inFIG. 2A) to further reduce noise generated during operation of airfoil400. In further embodiments, a chord line of airfoil400is adjusted in relation to a chord line of baseline airfoil408. For example, in some embodiments, airfoil400has a chord length that is greater than the chord length of baseline airfoil408. In alternative embodiments, airfoil400defines any chord that enables airfoil400to operate as described herein.

FIG. 11is a sectional schematic of cahedral and sweep of an airfoil500. Arrow502indicates a direction of cahedral of airfoil500. In particular, a dihedral (not shown) of airfoil500is in the direction of arrow502. In contrast, an anhedral (not shown) is in the opposite direction of arrow502. Arrow504indicates a direction of sweep of airfoil500. In particular, an aft sweep (not shown) of airfoil500is in the direction of arrow504. A forward sweep (not shown) is in the opposite direction of arrow504.

In reference back toFIG. 1, air116along blade128generates a flow structure190adjacent distal portion138of blade128. For comparison, airflow moving along a standard blade operating at a standard tip speed generates a compact flow structure adjacent a tip of the blade. As tip speed is reduced to lower noise, the blade loading increases such that the generated flow structures increase in size and energy, limiting the amount of noise reduction obtainable. For example, highly loaded blades at conditions such as takeoff have a thickened boundary layer and potentially flow separation which may lead to increased noise generation. In contrast, air116moving along blade128generates a relatively dispersed, low energy flow structure190adjacent distal portion138of blade128. Dispersing the flow structure by means of the blade design results in lower radiated noise during periods of aircraft operation such as during takeoff.

In a transonic flow regime, a shock impulse of the flow pattern is decreased in strength by blade128. In particular, when blade128reaches higher Mach numbers at the tip, such as during cruise, the shock impulse propagating from blade128has a decreased strength and is smoother (in the nearfield) compared to shock impulses from blades without cahedral136and sweep140. As a result, blade128reduces noise generated by systems using blade128.

In addition, gas turbine engine100with blade128including sweep140(shown inFIG. 3) and dihedral170(shown inFIG. 2A) generates less noise than both a rotary machine including a baseline blade and a rotary machine including a blade including only dihedral. When matched with thrust and accounting for the 3-dimensional airfoil design (including twist, camber, chord, and thickness) for corrections to aerodynamic and mechanical loading changes near the tip, gas turbine engine100including blade128generates less noise than the rotary machine including a baseline blade.

In some embodiments, gas turbine engine100includes a ducted and/or shrouded fan and blade128reduces the self and interaction noise generated by gas turbine engine100. For example, in some embodiments, gas turbine engine100is any of the following: an aerial vehicle, a turbofan, an air-handling fan, a lift fan, and a pump. Accordingly, during operation of gas turbine engine100, cahedral136and sweep140decreases noise generated by interaction between blade128and the shroud and/or duct. In addition, blade128can increase operability of gas turbine engine100. Furthermore, cahedral and sweep on stationary vane204may provide similar operability and performance enhancements for gas turbine engine200.

In reference toFIGS. 1-3, an exemplary method of operating a gas turbine engine100includes rotating drive shaft104to drive rotation of propeller assembly124. Blade128contacts air116imparting a force. In some embodiments, distal portion138contacts air116such that air116flows along distal portion138. A flow pattern is generated in air116adjacent the tip. In some embodiments, the flow pattern includes a shock impulse. In the exemplary method, dihedral170and sweep140of blade128are configured such that the shock impulse of the flow pattern is spread smoothly. In some embodiments of the method, air116is directed towards inlet118of casing112after air116is contacted by blade128.

The above-described rotary machine systems decrease the sound generated by the rotary machine systems by including a propeller assembly with blades having a cahedral and a sweep. The cahedral and the sweep reduce noise generated by the rotary machine during rotation of the propeller assembly. In some embodiments, the cahedral and the sweep extend throughout a distal portion of each blade of the propeller assembly to facilitate the rotary machine systems generating less noise during operation than known rotary machine systems. Moreover, the above-described rotary machine systems provide for increased efficiency of the rotary machine systems.

An exemplary technical effect of the embodiments described herein includes at least one of: (a) reducing noise generated by propeller assemblies and fan systems during operation; (b) increasing the operating efficiency of gas turbine engines, and (c) spreading force dispersed from blades to the air.

Exemplary embodiments of blades and propeller assemblies for use in gas turbine engines are described above in detail. The apparatus and systems described herein are not limited to the specific embodiments described, but rather, components of the apparatus and systems may be utilized independently and separately from other components described herein. For example, the blades may also be used in combination with other systems including rotating blades, and are not limited to practice with only the systems described herein. Rather, the exemplary embodiments can be implemented and utilized in connection with many machine system applications.