Patent Description:
Gas turbine engines can include a fan for propulsion air and to cool components. The fan also delivers air into a core engine where it is compressed. The compressed air is then delivered into a combustion section, where it is mixed with fuel and ignited. The combustion gas expands downstream over and drives turbine blades. Static vanes are positioned adjacent to the turbine blades to control the flow of the products of combustion.

Some fans include hollow fan blades made of a metallic or composite material. Various techniques can be utilized to construct hollow fan blades, including attaching a cover skin to an airfoil body.

<CIT> discloses a method of making hollow compressor blades, <CIT> discloses processes for manufacturing a hollow turbomachine blade, and <CIT> discloses a method of manufacturing a turbine component (fan blade) according to the preamble of claim <NUM>.

A method of forming a gas turbine engine component according to the invention is stated in claim <NUM> and includes attaching a cover skin to an airfoil body, the airfoil body and the cover skin cooperating to define pressure and suction sides of an airfoil, and moving the airfoil in a forming line including a plurality of stations. The plurality of stations include a set of heating stations, a deforming station and a set of cool down stations. The moving step includes positioning the airfoil in the set of heating stations to progressively increase a temperature of the airfoil, then positioning the airfoil in the deforming station including causing the airfoil to deform between first and second dies, and then positioning the airfoil in the set of cool down stations to progressively decrease the temperature of the airfoil.

Advantageously, the airfoil is metallic.

Preferably, each heating station of the set of heating stations includes one or more infrared heating elements.

Preferably, the set of heating stations includes two or more heating stations arranged in series.

Preferably, the temperature of the airfoil in at least one heating station of the set of heating stations is greater than a temperature of the first and second dies.

Preferably, the forming line includes a loading station upstream of the set of heating stations, and the method further includes positioning the airfoil in the loading station at a loading temperature between <NUM> degrees and <NUM> degrees Celsius (<NUM> degrees and <NUM> degrees Fahrenheit).

Preferably, the step of positioning the airfoil in the deforming station includes moving the first and second dies towards and into abutment with respective ones of the pressure and suction sides.

Preferably, the step of positioning the airfoil in the deforming station includes heating the first and second dies to a temperature of at least <NUM> degrees Celsius (<NUM> degrees Fahrenheit).

Preferably, the airfoil body extends from a root section to a tip portion. The tip portion defines a stagger angle relative to the root section, and the stagger angle is greater than or equal to <NUM> degrees, absolute, prior to the attaching step.

Advantageously, the step of causing the airfoil to deform occurs such that a change in the stagger angle of the airfoil presented to the deforming station is no more than <NUM> degrees, absolute.

Preferably, the attaching step includes welding at least a perimeter of the cover skin to the airfoil body.

Preferably, the cover skin is dimensioned to enclose at least one internal cavity in the airfoil body, and the attaching step includes trapping an inert gas in the at least one internal cavity.

More preferably, the inert gas comprises argon.

Preferably, the airfoil includes a plurality of airfoils, and the moving step includes moving the plurality of airfoils together as an airfoil set in each of the plurality of stations.

Preferably, includes positioning the airfoil set in a common support fixture, and wherein the moving step includes moving the common support fixture together with the airfoil set in the plurality of stations.

The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description.

The fan section <NUM> of the engine <NUM> is designed for a particular flight condition -- typically cruise at about <NUM> Mach and <NUM>,<NUM> meters (<NUM>,<NUM> feet). The flight condition of <NUM> Mach and <NUM>,<NUM> meters (<NUM>,<NUM> feet), with the engine at its best fuel consumption - also known as "bucket cruise Thrust Specific Fuel Consumption ('TSFC')" - is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. The "Low corrected fan tip speed" as disclosed herein according to one non-limiting embodiment is less than about <NUM> meters/second (<NUM> ft/second).

<FIG> illustrates a gas turbine engine component <NUM> according to an example. The component <NUM> can be incorporated in the gas turbine engine <NUM> of <FIG>, for example. In the illustrated example of <FIG>, the component <NUM> is an airfoil <NUM>. The airfoil <NUM> can be a fan blade 42A for the fan <NUM> of <FIG>, for example. Other types of airfoils, including blades, vanes and struts in the fan, compressor and turbine sections <NUM>, <NUM>, <NUM>, mid-turbine frame <NUM> and turbine exhaust case (TEC) <NUM> (<FIG>) may benefit from the examples disclosed herein which are not limited to the design shown. Other parts of the gas turbine engine <NUM> may benefit from the examples disclosed herein, including industrial turbines.

The airfoil <NUM> includes an airfoil section <NUM> extending in a spanwise or radial direction R from a root section <NUM>. The root section <NUM> is a shape that is configured to mount the fan blade 42A in the engine <NUM>, such as a dovetail shape. Generally, one side of the airfoil section <NUM> is a suction side SS and the other side is a pressure side PS (<FIG>) separated in a thickness direction T. The pressure side PS has a generally concave profile, and the suction side SS has a generally convex profile. The airfoil section <NUM> extends in the thickness direction T between the pressure and suction sides PS, SS to define an aerodynamic surface contour of the airfoil section <NUM>, as illustrated in <FIG>. The airfoil <NUM> is rotatable about an axis of rotation RR. The axis of rotation RR can be collinear or parallel to the engine axis A (<FIG>). The airfoil section <NUM> includes a first skin or airfoil body <NUM> that extends in the radial direction R from the root section <NUM> to a tip portion <NUM> (<FIG>). The tip portion <NUM> is a terminal end of the airfoil <NUM>. The airfoil body <NUM> extends in a chordwise direction X between a leading edge LE and a trailing edge TE. The airfoil body <NUM> defines at least one of the pressure and suction sides PS, SS. In the illustrated example of <FIG> and <FIG>, the airfoil body <NUM> defines both the pressure and suction sides PS, SS.

The airfoil <NUM> includes a cover (or second) skin <NUM> disposed on a surface of the airfoil body <NUM> and is arranged to provide a continuous surface with the suction side SS of the airfoil <NUM>, as illustrated by <FIG>. In another example, the cover skin <NUM> is disposed on the pressure side PS of the airfoil <NUM>. The cover skin <NUM> is shown in an uninstalled position in <FIG> for illustrative purposes. The component <NUM> can include two or more cover skins along each of the pressure and/or suction sides PS, SS of the airfoil section <NUM>.

The airfoil body <NUM> and cover skin <NUM> can be made out of metallic materials such as titanium or aluminum. Other materials for the airfoil body <NUM> and cover skin <NUM> can be utilized, including metals or alloys and metal matrix composites.

Referring to <FIG> with continuing reference to <FIG>, the airfoil <NUM> includes at least one internal cavity <NUM> defined in the airfoil section <NUM>. In other examples, the internal cavities <NUM> are omitted such that the airfoil section <NUM> is substantially or completely solid. In the illustrative example of <FIG>, the airfoil body <NUM> includes one or more ribs <NUM> that define a plurality of internal cavities <NUM>. The airfoil <NUM> can include fewer or more than three internal cavities <NUM>, such as only one internal cavity <NUM>. Each internal cavity <NUM> can be defined having different dimensions, shapes and at other orientations than illustrated by <FIG> and <FIG>. The internal cavities <NUM> can substantially or completely free of any material such that the airfoil section <NUM> is hollow.

In the illustrated example of <FIG>, ribs 74A have a generally circular or otherwise elliptical geometry, ribs 74B have generally elongated, oblong or racetrack shaped geometry, and ribs 74C are generally linear or curvilinear. Ribs 74A, 74B and 74C have a thickness TA, TB and TC, respectively. In examples, thicknesses TA, TB are greater than or equal to about <NUM> (<NUM> inches) and less than or equal to about <NUM> (<NUM> inches) or more narrowly between <NUM> to <NUM> (<NUM> and <NUM> inches). Thickness TC can be greater than thicknesses TA, TB, such as between <NUM> to <NUM> (<NUM> and <NUM> inches), for example. Ribs 74A, 74B can be attached to the cover skin <NUM> utilizing any of the techniques disclosed herein, including laser or electron beam welding, brazing, diffusion bonding or other fastening techniques. At least some of the ribs <NUM> can be spaced apart from the cover skin <NUM> to define a gap GG when in an assembled position, as illustrated by rib 74C of <FIG>.

Walls <NUM> of the component <NUM> bound the internal cavities <NUM>. The walls <NUM> can be internal or external walls of the component <NUM>. The airfoil body <NUM> and cover skin <NUM> define one or more of the walls <NUM>. The cover skin <NUM> is attached to the airfoil body <NUM> to enclose or otherwise bound the internal cavities <NUM>, with the airfoil body <NUM> and cover skin <NUM> cooperating to define the pressure and suction sides PS, SS of the airfoil section <NUM>.

Referring to <FIG>, span positions of the airfoil section <NUM> are schematically illustrated from <NUM>% to <NUM>% in <NUM>% increments to define a plurality of sections <NUM>. Each section <NUM> at a given span position is provided by a conical cut that corresponds to the shape of segments a flowpath (e.g., bypass flowpath B or core flow path C of <FIG>), as shown by the large dashed lines. In the case of an airfoil <NUM> such as with an integral platform <NUM>, the <NUM>% span position corresponds to the radially innermost location where the airfoil section <NUM> meets the fillet joining the airfoil <NUM> to the platform <NUM> (see also <FIG> illustrating platform <NUM>). In the case of an airfoil <NUM> without an integral platform, the <NUM>% span position corresponds to the radially innermost location where the discrete platform <NUM> meets the exterior surface of the airfoil section <NUM>. A <NUM>% span position corresponds to a section of the airfoil section <NUM> at the tip portion <NUM>.

Referring to <FIG> with continuing reference to <FIG>, the airfoil section <NUM> is sectioned at a radial position between the root section <NUM> and tip portion <NUM>. In examples, each airfoil section <NUM> is specifically twisted about a spanwise axis in the radial direction R with a corresponding stagger angle α at each span position. Chord CD, which is a length between the leading and trailing edges LE, TE, forms stagger angle α relative to the chordwise direction X or a plane parallel to the axis or rotation RR. The stagger angle α can vary along the span of the airfoil section <NUM> to define a twist. For example, the tip portion <NUM> can define a stagger angle α relative to the root section <NUM> that is greater than or equal to <NUM> degrees or <NUM> degrees, absolute. In some examples, the stagger angle α at the tip portion <NUM> relative to the root section <NUM> is between <NUM>-<NUM> degrees, absolute, or more narrowly between <NUM>-<NUM> degrees, absolute, such that the airfoil section <NUM> is twisted about a spanwise axis as illustrated by the airfoil <NUM> of <FIG> and <FIG>. The airfoil section <NUM> can be three-dimensionally twisted about the spanwise axis.

<FIG> illustrates a process of constructing or forming a gas turbine engine component in a flow chart <NUM>. The process can be utilized to form the component <NUM> of <FIG> and <FIG>, including an airfoil <NUM> such as fan blade 42A, another hollow airfoil, or a solid airfoil, for example. In this invention, 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. Reference is made to component <NUM> of <FIG> for illustrative purposes, which disclose exemplary conditions or states of the component <NUM> in the process <NUM>. In the illustrative example of <FIG>, the component <NUM> is a hollow airfoil <NUM> or fan blade including a metallic airfoil section <NUM>. The techniques disclosed herein can be utilized to form a new component or to repair a previously fielded component.

Referring to <FIG> and <FIG>, one or more portions of the component <NUM> can be prepared or otherwise provided at steps 176A-176E (shown in dashed lines). The component <NUM> includes a first skin or airfoil/main body <NUM> and a cover (or second) skin <NUM> that define one or more walls <NUM> of the component <NUM>.

At step 176A, airfoil body <NUM> is formed with respect to a predefined blade geometry, which can be defined with respect to one or more design criteria. The airfoil body <NUM> can be forged, cast, or produced by additive manufacturing from a metal or metal alloy, for example. At step 176B, internal and/or external surfaces of the airfoil body <NUM> are machined with respect to the predefined blade geometry. At step 176C, cover skin <NUM> is hot formed with respect to a predefined cover geometry. The cover skin <NUM> can be formed from sheet metal, for example. Airfoil body <NUM> and cover skin <NUM> can be machined and hot formed, respectively, to a configuration that compensates for weld distortion such that upon entering deform machine <NUM> no more than a <NUM> or <NUM> degree change in the stagger angle α is required or otherwise occurs. At step 176D, the cover skin <NUM> is chemically milled with respect to the predefined cover geometry. At step 176E, the cover skin <NUM> is cleaned to remove surface contaminants using a laser cleaning technique, for example.

One or more internal cavities <NUM> are formed in the airfoil body <NUM> and/or the cover skin <NUM> (internal cavities <NUM>' defined in cover skin <NUM> shown in dashed lines for illustrative purposes). Ribs <NUM> can be arranged to define various geometries of the internal cavities <NUM>, including any of the geometries of ribs <NUM> of <FIG>.

Various techniques can be utilized to form the internal cavities <NUM>, including casting, machining or additive manufacturing techniques. The internal cavities <NUM> can be defined in the airfoil body <NUM> and/or cover skin <NUM> during steps 176A-176C, for example. The cover skin <NUM> is dimensioned to enclose at least one, or more than one, internal cavity <NUM> in the airfoil body <NUM> when in an installed position.

At step 176F, cover skin <NUM>' is positioned relative to the airfoil body <NUM>. Cover skin <NUM>' is shown in dashed lines in <FIG> at a distance away from the airfoil body <NUM> for illustrative purposes. The positioning can include moving the cover skin <NUM>' in a direction DA and into abutment with ribs <NUM> of the airfoil body <NUM> to define a pre-finished state of the airfoil section <NUM>, as illustrated by cover skin <NUM>.

At step <NUM>, the cover skin <NUM> is attached to the airfoil body <NUM> to define the airfoil <NUM>. In examples, a perimeter P (see also <FIG>) of the cover skin <NUM> and/or locations of the cover skin <NUM> abutting the ribs <NUM> are attached to the airfoil body <NUM> to enclose or otherwise bound the internal cavities <NUM>. Various techniques can be utilized to attach the cover skin <NUM> to the airfoil body <NUM>, including laser or electron beam welding, brazing, diffusion bonding or other fastening techniques. The predefined blade and cover geometries prior can be set with respect to an expected distortion in the airfoil <NUM> caused by attachment of the airfoil body <NUM> and cover skin <NUM> during the attaching step <NUM>.

The airfoil body <NUM> extends from a root section to a tip portion (e.g., root section <NUM> and tip portion <NUM> of <FIG>) to define a stagger angle α (<FIG>) such that the airfoil body <NUM> is twisted. The stagger angle α of the airfoil section <NUM> can include any of the stagger angles α disclosed herein, such as being greater than or equal to <NUM> degrees, absolute, at the airfoil tip relative to the root section prior to attaching the cover skin <NUM> at step <NUM>.

Attaching the cover skin <NUM> can include trapping an inert gas in each internal cavity <NUM>. In the illustrated example of <FIG>, the component <NUM> can be situated in a controlled environment E (shown in dashed lines) prior to and during the attaching step <NUM>. A fluid source FS (shown in dashed lines) is operable to convey an amount of fluid F to the environment E. Example fluids F include inert gases such as argon or helium. The fluid F circulates in the environment E and is communicated to the internal cavities <NUM>. Attaching the cover skin <NUM> to the airfoil body <NUM> can cause an amount of the fluid F to be trapped in the internal cavities <NUM>. In other examples, fluid F is communicated to the internal cavities via passages in the root section (see, e.g., root section <NUM>, cavities <NUM> and fluid source FS of <FIG>). Walls of the ribs <NUM> can include one or more vent holes <NUM> (shown in dashed lines in <FIG>) at approximately mid-point within the rib <NUM>, for example, to permit equalization of pressure of the trapped inert gases between adjacent internal cavities <NUM> during attaching step <NUM>.

<FIG> illustrates a continuous flow automated forming line <NUM> that can be utilized with the exemplary process <NUM> to form the component(s) <NUM> with respect to a predefined geometry, such as the airfoil section <NUM> of <FIG>, for example. However, other components can be utilized according to the teachings disclosed herein, such as vanes and endwalls.

The forming line <NUM> can be a continuous flow automated line including a plurality of stations <NUM> (shown in dashed lines). The forming line <NUM> can include a controller CONT (shown in dashed lines) that is operable to index, move or position the component(s) <NUM> in and through each of the stations <NUM> according to a predefined schedule. One would understand how to configure the controller CONT with logic to execute the predefined schedule according to the teachings disclosed herein. It should be appreciated that other techniques for moving the components <NUM> can be utilized including, but not limited to, manually positioning the components <NUM> in each station <NUM>. Each of the stations <NUM> can be separate and distinct from one another.

The stations <NUM> can include at least a loading station 180A, a set of heating stations 180B, a deforming station 180C, a set of cool down stations 180D, and an unloading station 180E. The loading station 180A is upstream of the heating stations 180B, which are upstream of the deforming station 180C. The deforming station 180C is upstream of the cool down stations 180D, which are upstream of the unloading station 180E. The forming line <NUM> can include a conveyor <NUM> moveable in a direction FD to move or position the components <NUM> in each of the stations <NUM>.

Referring to <FIG> with continuing to reference to <FIG>, at step <NUM> at least one component <NUM> such as airfoil <NUM> is positioned in the forming line <NUM> at the loading station <NUM> subsequent to attaching step <NUM>. The airfoil <NUM> can be positioned in a root upward orientation as illustrated by the airfoils <NUM> and respective root sections <NUM> of <FIG>, for example. The positioning step <NUM> can include positioning a plurality of airfoils <NUM>, such as two airfoils <NUM>, at the loading station 180A to move the airfoils <NUM> together as an airfoil set <NUM> in each of the stations <NUM>. It should be appreciated that fewer or more than two airfoils <NUM> can be moved at a time in each of the stations <NUM>. The airfoil set <NUM> can be positioned in a common support fixture <NUM>. The conveyor <NUM> can be an overhead conveyer operable to move the airfoil set <NUM> together with the common support fixture <NUM> in each of the stations <NUM>.

The components <NUM> can be enclosed in each of the respective heating, deforming and/or cool down stations 180B, 180C, 180D. Each station <NUM> can be an "open air" or non-vacuum environment, which can reduce a complexity of forming the components <NUM>. The open air environment can exclude any furnace or protective environment such as a vacuum or argon environment.

The components <NUM> are moved or positioned in the loading station 180A at a loading temperature. The loading temperature can be above <NUM> degrees Celsius (°C) (<NUM> degrees Fahrenheit (F)) such as between <NUM> and <NUM> (<NUM> and <NUM> degrees Fahrenheit (F)). In some examples, the loading temperature can be set to approximately room temperature. For the purposes of this disclosure, the term "room temperature" means a temperature between <NUM> and <NUM> (between <NUM> degrees and <NUM> degrees Fahrenheit (F)) and the term "approximately" means ±<NUM>% of the value unless otherwise disclosed.

At step 176I, from the loading station 180A each component <NUM> enters a controlled heating phase including preheating the components <NUM> to a predetermined temperature prior to positioning the components <NUM> in the forming station 180C. Step 176I includes positioning the components <NUM> in the set of heating stations 180B to progressively heat or increase a temperature of the components <NUM> to a first predetermined temperature threshold.

The set of heating stations 180B can include two or more heating stations 180B arranged in a series. In the illustrated example of <FIG>, the set of heating stations <NUM> includes four separate and distinct heating stations 180B-<NUM> in 180B-<NUM> arranged in series. In other examples, the forming line <NUM> includes only one heating station 180B.

In the illustrated example of <FIG>, the first heating station 180B-<NUM> is set to a temperature of approximately <NUM> degrees Fahrenheit (F) (<NUM>). The second heating station 180B-<NUM> is set to a temperature of approximately <NUM> (<NUM> degrees Fahrenheit (F)). The third heating station 180B-<NUM> is set to a temperature of approximately <NUM> (<NUM> degrees Fahrenheit (F)) and the fourth heating station 180B-<NUM> is set to a temperature of approximately <NUM> (<NUM> degrees Fahrenheit (F)). The components <NUM> together with the common support fixture <NUM> are moved or advanced in each of the heating stations 180B in approximately <NUM>-<NUM> minute increments, for example. It should be appreciated that other temperatures and heating durations can be utilized with the teachings disclosed herein.

Various techniques can be utilized to heat the components <NUM> in each of the heating stations 180B. In the illustrated example of <FIG>, each heating station 180B includes one or more infrared heating elements <NUM>, such as an array of quartz lamps. Each quartz lamp or other heating element <NUM> can be independently controllable in its intensity so as to provided uniform temperature over the entire component <NUM> at each station of forming line <NUM>. The heating elements <NUM> are oriented towards the components <NUM> when the components <NUM> are positioned in the respective heating stations 180B. In some examples, the components <NUM> together with the support fixture <NUM> remain substantially stationary in the respective station <NUM> for a predetermined duration to ensure substantially uniform heating and/or cooling of the components <NUM>. Preheating the component <NUM> prior to positioning the components <NUM> in the deforming station 180C can relax or otherwise reduce residual stresses in the component <NUM> due to attachment of the cover skin <NUM> to the airfoil body <NUM> during step <NUM>. For example, approximately <NUM>% relaxation or movement of the airfoil section <NUM> toward a target aerodynamic profile can occur in response to the component <NUM> being positioned in the third heating station 180B-<NUM> for the predetermined duration.

Referring to <FIG> with continuing reference to <FIG> and <FIG>, at step 176J the components <NUM> are moved from the last heating station 180B-<NUM> to the deforming station 180C. The deforming station 180C includes a deforming machine <NUM>. The deforming machine <NUM> includes a base 190A mounted to a static structure. The machine <NUM> includes a pair of supports 190B extending from the base 190A. The airfoils <NUM> can be suspended or otherwise supported by respective root sections <NUM> in the common support fixture <NUM> (shown in dashed lines) residing above deforming machine <NUM> such that the airfoils <NUM> are oriented substantially vertically between the dies <NUM>, <NUM> with tip portions <NUM> positioned downward or otherwise below respective root section <NUM> as illustrated by <FIG>. Vertically orienting the airfoils <NUM> by hanging or suspending the airfoils <NUM> by the respective root sections <NUM> can reduce spanwise distortions such as buckling during heating and cooling and relation of the airfoils <NUM>.

The common support fixture <NUM> is moved in direction FD to position each airfoil section <NUM> between respective a set of first and second dies (or die halves) <NUM>, <NUM>. The dies <NUM>, <NUM> are contoured to mate with pressure and suction sides PS, SS of the respective airfoil section <NUM> of the airfoil <NUM>. In the illustrative example of <FIG>, each of the dies is a split die <NUM>'/<NUM>' to facilitate replacement of portions of the die <NUM>'/<NUM>'.

The deforming station 180C is operable to cause the airfoil section <NUM> of each airfoil <NUM> to deform or resize between the first and second dies <NUM>, <NUM>. The machine <NUM> includes one or more actuators 190C that are operable to move the dies <NUM>, <NUM> in response to signal(s) from controller CONT (shown in dashed lines). Movement of the dies <NUM>, <NUM> includes exerting a pressure on surfaces of the airfoil section <NUM> sufficient to cause a predetermined amount of deformation to occur.

The dies <NUM>, <NUM> are operable to heat the components <NUM> to a second predetermined temperature threshold prior to and during holding the components <NUM> under compression by applying pressure from the actuators 190C. For example, the dies <NUM>, <NUM> can be heated to and continuously operating at a temperature of at least <NUM> (<NUM> degrees Fahrenheit (F)), or more narrowly between approximately <NUM> and <NUM> (<NUM> and <NUM> degrees Fahrenheit (F)). In some examples, the second temperature threshold or range is equal to or greater than the first temperature threshold or range of the set of heating stations 180B.

Each component <NUM> is preheated to the first temperature threshold subsequent to the attaching step <NUM> and prior to abutment with the dies <NUM>, <NUM>. For example, the temperature of the component <NUM> can be at least <NUM>% of a surface temperature of the dies <NUM>, <NUM>, or more narrowly between <NUM>% and <NUM>% of the surface temperature, when the component <NUM> is initially positioned in the deforming station 180C between, but prior to abutment with, the dies <NUM>, <NUM>.

In the illustrative example of <FIG>, the temperature of the components <NUM> in at least one of the heating stations 180B is greater than a temperature of the first and second dies <NUM>, <NUM>. The second temperature threshold is less than the first temperature threshold or range such that a temperature of the components <NUM> is reduced subsequent to moving from the set of heat stations 180B. A reduction of temperature may occur due to transit between the heating station 180B-<NUM> and the deforming station 180C.

For example, the dies <NUM>, <NUM> are operable to heat the respective component <NUM> to a temperature of approximately <NUM> (<NUM> degrees Fahrenheit (F)). The fourth heating station 180B-<NUM> can be set to a temperature of <NUM> (<NUM> degrees Fahrenheit (F)) In other examples, the temperature of the component <NUM> exiting the fourth heating station 180B-<NUM> is less than the temperature of the first and second dies <NUM>, <NUM> when the respective component <NUM> is moved to the deforming station 180C.

<FIG> illustrates a backside of a refined version one of the dies <NUM>/<NUM>. Each of the die <NUM>/<NUM> can include one or more heating elements <NUM> that are positioned in a backside cavity of the die <NUM>/<NUM>. Each die <NUM>/<NUM> can be made of metal or a metal alloy, such as a cast nickel alloy which can improve the ability of continuously operating the dies <NUM>, <NUM> at the second predefined temperature threshold.

Each heating element <NUM> can be a heating coil that is coupled to an energy source ES (shown in dashed lines). The energy source ES can be a power supply operable to communicate electrical current to the heating element <NUM> in response to controller CONT to heat the respective die <NUM>/<NUM> to the second temperature threshold. The controller CONT can be coupled to at least one sensor <NUM> (shown in dashed lines), such as a thermocouple, to monitor surface temperatures of the respective die <NUM>/<NUM>. The controller CONT is operable to adjust the temperature of the die <NUM>/<NUM> to maintain or otherwise approach the second temperature threshold.

A non-metallic heat conductive layer <NUM> such as cloth can be situated between the heating elements <NUM> and surfaces of the die <NUM>/<NUM> to reduce a likelihood of arcing. At least one coating <NUM> can be deposited on surfaces of the die <NUM>/<NUM>. Example coatings include diffused aluminide which can provide oxidation protection.

Referring to <FIG> and <FIG>, at step 176J the component <NUM> undergoes permanent deformation to vary a geometry of the walls <NUM> of the airfoil body <NUM> and/or cover skin <NUM> (<FIG>). Step 176J includes moving the first and second dies <NUM>, <NUM> in opposed directions D1, D2 (<FIG>) towards and into abutment with respective ones of the pressure and suction sides PS, SS of the respective airfoil section <NUM>, as illustrated by <FIG> and <FIG>.

Each airfoil section <NUM> is clamped or held in compression between respective dies <NUM>, <NUM> at or approximately the second temperature threshold for a predetermined duration, such as approximately <NUM>-<NUM> minutes, to cause the airfoil section <NUM> to permanently deform between the dies <NUM>, <NUM>. The predetermined duration can be set to cause the airfoil section <NUM> to undergo creep deformation or hot sizing, to minimize or otherwise reduce the residual stresses in the component <NUM> that may be caused during the attaching step <NUM>, and to allow the walls <NUM> to conform to the surface profile of the dies <NUM>, <NUM>. In some examples, the deformation of the airfoil section <NUM> occurs such that a change in the stagger angle α (see <FIG>) of the of airfoil <NUM> that is presented to the deforming station 180C is no more than <NUM> or <NUM> degrees, absolute, at the tip portion relative to the root section. The deformation due to hot sizing the component <NUM> can be less than <NUM> (<NUM> inches) for example.

The dies <NUM>, <NUM> can serve as "gas sizing" dies that are utilized to cause at least a portion of the component <NUM> to undergo deformation. Creep deformation, hot sizing and gas sizing are generally known. However, utilization of such techniques to form the components in situ as disclosed herein are not known. For example, heating of the fluid F trapped in the internal cavities <NUM> (<FIG>) of the component <NUM> during the attaching step <NUM> causes the internal cavities <NUM> to pressurize and the walls <NUM> of the airfoil section <NUM> to move outwardly or otherwise deform during the deforming step 176J.

Referring to <FIG> and <FIG>, at step <NUM> each component <NUM> enters a controlled cool down phase subsequent to the deforming step 176J. The components <NUM> are moved in the set of cool down stations 180D to progressively decrease the temperature of each component <NUM>. The set of cool down stations 180D can include two or more cool down stations 180D arranged in a series. In the illustrated example of <FIG>, the set of cool down stations <NUM> includes four separate and distinct cool down stations 180D-<NUM> through 180D-<NUM> arranged in series. In other examples, the forming line <NUM> includes only one cool down station 180D. The controlled cool down phase can reduce residual stresses that may otherwise be reintroduced in the component <NUM> were the component <NUM> to otherwise be moved directly from the deforming station 180C to the unloading station 180E at room temperature.

Various techniques can be utilized to cool down the components <NUM> in each of the cool down stations 180D. In the illustrated example of <FIG>, each cool down station 180D includes one or more infrared heating elements <NUM> oriented towards the components <NUM> when the components <NUM> are positioned in the cool down stations 180D.

In the illustrated example of <FIG>, the first cool down station 180D-<NUM> is set to a temperature of approximately <NUM> (<NUM> degrees Fahrenheit (F)) The second cool down station 180D-<NUM> is set to a temperature of <NUM> (<NUM> degrees Fahrenheit (F)). The third cool down station 180D-<NUM> is set to a temperature of approximately <NUM> (<NUM> degrees Fahrenheit (F)), and the fourth cool down station 180D-<NUM> is set to a temperature of between <NUM> and <NUM> (<NUM> degrees and <NUM> degrees Fahrenheit (F)) such as approximately room temperature. The components <NUM> together with the common support fixture <NUM> are moved in each of the cool down stations 180D in approximately <NUM>-<NUM> minute increments, for example. It should be appreciated that other temperatures and cool down durations can be utilized with the teachings disclosed herein. At step <NUM>, the components <NUM> are unloaded from the forming line <NUM> at an unloading temperature. The unloading temperature can be above °C (<NUM> degrees Fahrenheit (F)) such as between <NUM> and <NUM> (<NUM> degrees and <NUM> degrees Fahrenheit (F)), or more narrowly approximately room temperature, for example. The components <NUM> can be loaded into the support fixture <NUM> at the loading station 180A and unloaded from the support fixture <NUM> at the unloading station 180E utilizing various techniques, such as by a robot or manually.

One or more finishing steps can be performed subsequent to unloading the components <NUM> at step <NUM>. For example, an interior inspection of the component <NUM> can occur at step <NUM>. One or more final machining operations of the component <NUM> can occur at step 176N. A final inspection of the component <NUM> can occur at step 176O.

The process and forming line disclosed herein can be utilized to rapidly dimensionally correct the components subsequent to welding or otherwise attaching the various components. The process can be performed in an open air environment, which can reduce complexity.

Claim 1:
A method of forming a gas turbine engine component (<NUM>; <NUM>) comprising:
attaching a cover skin (<NUM>; <NUM>) to an airfoil body (<NUM>; <NUM>), the airfoil body (<NUM>; <NUM>) and the cover skin (<NUM>; <NUM>) cooperating to define pressure and suction sides (PS, SS) of an airfoil (<NUM>; <NUM>); and characterized by
moving the airfoil (<NUM>; <NUM>) in a forming line (<NUM>) including a plurality of stations (<NUM>), the plurality of stations (<NUM>) including a set of heating stations (180B), a deforming station (180C) and a set of cool down stations (180D) wherein the moving step includes positioning the airfoil (<NUM>; <NUM>) in the set of heating stations (180B) to progressively increase a temperature of the airfoil (<NUM>; <NUM>), then positioning the airfoil (<NUM>; <NUM>) in the deforming station (180C) including causing the airfoil (<NUM>; <NUM>) to deform between first and second dies (<NUM>, <NUM>), and then positioning the airfoil (<NUM>; <NUM>) in the set of cool down stations (180D) to progressively decrease the temperature of the airfoil (<NUM>; <NUM>).