RTX Corporation Patente

An abradable coating/thermal barrier coating (235) suitable for use with jet engine CMC components is described which comprises a material selected from hafnon, mixtures of hafnon and zircon, and rare earth disilicates (RE₂Si₂O₇), wherein RE is Sc, Y, La, Ce, Pr, Nd, Pm, Sm, Eu, Gd, Tb, Dy, Ho, Er, Tm, Yb, or Lu. The coating (235) has a porosity gradient wherein the porosity decreases in a radial direction. The porosity gradient provides a progressively increasing wear resistance to slow the rate of rub interaction. The porosity gradient also reduces the thermal gradient through the coating (235) thereby reducing the formation of thermal stresses within the coating (235) and any underlying CMC component (200).

A method of forming a multilayer environmental barrier coating, comprising: providing a ceramic matrix composite; depositing a bond coat layer on the ceramic matrix composite; depositing a top coat layer on the bond coat layer; depositing an over coat layer on the top coat layer, wherein the over coat layer includes no more than approximately 50 percent by volume of porosity.

An assembly is provided for an aircraft powerplant (20). This assembly includes a vane structure (88) and a heat exchanger (90). The vane structure (88) extends longitudinally to a leading edge (108) of the vane structure (88). The vane structure (88) includes a translating body (114), a stationary body (112) and an inlet passage (150). The translating body (114) is configured to translate longitudinally between a first position and a second position. The translating body (114) is configured to form the leading edge (108) of the vane structure (88) and close an inlet (156) into the inlet passage (150) when in the first position. The translating body (114) is configured to open the inlet (156) into the inlet passage (150) when in the second position. The heat exchanger (90) is disposed within the stationary body (112). The inlet passage (150) projects into the stationary body (112) from the inlet (156) into the inlet passage (150) to the heat exchanger (90).

A gas turbine engine component (200), comprising a substrate (210) containing a ceramic matrix composite material, the ceramic matrix composite material includes at least one environmental barrier coating (230) comprising at least hafnon and hafnium titanate.

An apparatus, comprising an optical probe (102) having optical components therein. A gas path router (104) receives an airflow from a gas path (106) of a gas turbine engine. The gas path router (104) is configured to route the airflow from the gas path (106) past the optical components of the optical probe (102) to purge debris from the optical components.

A constrained linear spring (20) can increase a spring rate of a linear spring. A constrained linear spring (20) includes a linear spring element (100) having a first end (102) and extends in a first direction to a second end (104), with one or more undulations extending in a second direction transverse to the first direction. The linear spring element (100) has a first spring rate when compressed in the second direction with the first and second ends (102, 104) unconstrained. The constrained linear spring (20) also includes a constraining element (200; 210; 220) dimensioned to contact the first and second ends (102, 104) and at least one undulation of the linear spring element (100) when the linear spring element (100) is uncompressed. The constrained linear spring element (100) of the constrained linear spring (20) has a second spring rate higher than the first spring rate when compressed in the second direction that can be useful to counteract a forward overturning moment (50) in an airfoil (45) of a vane (40).

Acoustic structures (122; 222; 322) for a gas turbine engine (100) are disclosed. In various embodiments, the acoustic structures (122; 222; 322) comprise a panel (118; 218) in the form of a cellular structure (126; 226; 326) having a plurality of cells (128; 228; 328); and a housing (240) configured to house the cellular structure (126; 226; 326), the housing (240) including a first rail (242) configured for attachment to a static structure within the gas turbine engine (100) and a second rail (244) configured for attachment to the static structure. In various embodiments, the acoustic structures (122; 222; 322) comprise a liner (116; 316) in the form of a non-segmented or two-piece structure configured to extend about the inner barrel of a nacelle structure (110). The disclosed panels and liners are configured to maximize the operational area of the acoustic structures (122; 222; 322) by eliminating obstructions from adjacent flow paths.

A powerplant (20) for an aircraft includes a first turbine engine (22), a second turbine engine (24) and a fuel cell system (26). The second turbine engine (24) includes a second engine flowpath (100), a second engine compressor section (96), a second engine combustor section (97 and a second engine turbine section (98). The second engine flowpath (100) extends from a second engine flowpath inlet (114) to a second engine flowpath outlet (116). The second engine flowpath inlet (114) and the second engine flowpath outlet (116) are each fluidly coupled with a flowpath (80) of the first turbine engine (22). The fuel cell system (26) includes a fuel cell, a fuel circuit (142) and an air circuit (144). The fuel circuit (142) extends through the fuel cell and is fluidly coupled with and upstream of a first fuel injector in the second engine combustor section (97). The air circuit (144) extends through the fuel cell and is fluidly coupled with and downstream of a bleed (160) from the second engine flowpath (100).

A gas turbine engine (100) includes a fan shaft (110) rotatable about an axis (A), a fan (112) connected to the fan shaft (110), and an outer housing (114) surrounding the fan (112) to define a bypass passage (B). A compressor section (24) has both a low pressure compressor (106) that is fixed to rotate with the fan shaft (116) and a high pressure compressor (124). A turbine section (28) has a low pressure turbine (104) driving a low speed spool (116) and a high pressure turbine (122) driving the high pressure compressor (124). The low pressure turbine (104) includes rotating blades and a static structure (130) aft of the rotating blades. A fan drivetrain (120) includes a geared architecture (102) connecting the low speed spool (116) to the fan shaft (110) such that the fan (112) and low pressure compressor (106) rotate at a lower speed than the low pressure turbine (104). A thrust bearing (132) is axially aft of the geared architecture (102) and axially forward of the high pressure compressor (122). A tower shaft (140) is rotatably driven by the low speed spool (116).

An aerodynamic component (501, 801) of a gas turbine engine (20) is provided and is fittable to a shell (510, 810) having a shell shape (512, 812). The aerodynamic component includes a body (520, 820) having a component shape (522, 822) initially deviating from the shell shape prior to an assembly operation in which the aerodynamic component is to be fit to the shell. Deviation of the component shape from the shell shape aids in an establishment of a final desired shape (530, 830) of the aerodynamic component following the assembly operation.

A gas turbine engine (10) includes a fan section (12), a compressor section (14), and a turbine section (18). The fan section (12) has a plurality of vane assemblies (64) spaced circumferentially about an engine axis (A) and each including an airfoil (66) extending between a leading edge (66a) and a trailing edge (66b), a control rod (68) extending through the airfoil (66), and a mechanism driven by the control rod (68) to change the shape of the airfoil (66).

A method is provided for repairing a turbine engine component (20). This method includes: machining away a worn land surface of an annular seal land (46A: 46B) to provide the annular seal land (46A; 46B) with a machined base surface (66A; 66B), the turbine engine component (20) including the annular seal land (46A; 46B), the annular seal land (46A; 46B) extending axially along an axis (32) to the machined base surface (66A; 66B), and the machined base surface (66A; 66B) extending circumferentially around the axis (32); brazing a repair ring (60A; 60B) to the machined base surface (66A; 66B) to provide a seal land preform; and machining the repair ring (60A; 60B) of the seal land preform to provide the turbine engine component (20) with a repaired seal land (46A; 46B) having a repaired land surface (78A; 78B), the repaired seal land (46A; 46B) extending axially along the axis (32) to the repaired land surface (78A; 78B), and the repaired land surface (78A; 78B) extending circumferentially around the axis (32).

A propulsion system (20) for an aircraft includes a first open propulsor rotor (40), a second open propulsor rotor (42), a turbine engine (32), a first drivetrain (36) and a second drivetrain (38). The first drivetrain (36) operatively couples the turbine engine (32) to the first open propulsor rotor (40). The turbine engine (32) is configured to drive rotation of the first open propulsor rotor (40) through the first drivetrain during a first mode and a second mode. The second drivetrain (38) is configured to operatively couple the turbine engine (32) to the second open propulsor rotor (42) during the first mode and to operatively decouple the turbine engine (32) from the second open propulsor rotor (42) during the second mode. The second drivetrain (38) is configured as or otherwise includes a tower shaft. The turbine engine (32) is configured to drive rotation of the second open propulsor rotor (42) through the tower shaft during the first mode.

An aircraft propulsion system (20) includes a thrust reverser assembly (100) that includes a reverser door (102) corresponding to a cooling air outlet duct (86) for directing a cooling air flow (112) from a condenser (88) forward.

An assembly (20) for an aerial vehicle includes a vehicle body (24), a gas turbine engine (22) and a support structure (26). The vehicle body (24) includes a body section (111). The gas turbine engine (22) includes a stationary structure (66). The gas turbine engine (22) is housed within the vehicle body (24). The support structure (26) extends between and is connected to the body section (111) and the stationary structure (66). The support structure (26) supports the gas turbine engine (22) within the vehicle body (24). The body section (111), the stationary structure (66) and the support structure (26) are included in a monolithic body (70).

A process for eliminating residual stress (14) in a forging (10) for rotors including providing the forging (10) for the rotor; and removing a bulk material from an area located near airfoils (20) of a finish part (18) formed from the forging prior to process steps selected from the group consisting of a rough turn, a finish mill of blades, a vibratory polish of the blades and a finish turn.

A propulsion system (20) for an aircraft includes a first open propulsor rotor (40), a second open propulsor rotor (42), a turbine engine (32) and a drive system (34). The second open propulsor rotor (42) is next to and downstream of the first open propulsor rotor (40). The turbine engine (32) includes a flowpath (72), a compressor section (67), a combustor section (68), a turbine section (69) and a rotating assembly (90). The rotating assembly (90) includes a turbine rotor (81) in the turbine section (69). The rotating assembly (90) is configured to drive rotation of the first open propulsor rotor (40) during a first mode and a second mode. The drive system (34) is discrete from the rotating assembly (90). The drive system (34) is configured to drive rotation of the second open propulsor rotor (42) during the first mode. The second open propulsor rotor (42) is rotationally fixed during the second mode.

An open propulsor rotor (58) is configured to rotate about a first axis (62). An open guide vane structure (60) includes a plurality of open guide vanes (78) arranged circumferentially about a second axis (112) offset from the first axis (62). The open guide vanes (78) include a first open guide vane. A turbine engine (54) includes a flowpath (96), a compressor section (91), a combustor section (92), a turbine section (93) and a rotating assembly. The flowpath (96) extends through the compressor section (91), the combustor section (92) and the turbine section (93) from an airflow inlet (98) into the flowpath to a combustion products exhaust (100) from the flowpath (96). The airflow inlet (98) is axially aligned with or located forward of the first open guide vane along the second axis (112). The rotating assembly includes a turbine rotor disposed in the turbine section (93). The rotating assembly is configured to drive rotation of the open propulsor rotor (58) through an offset geartrain (122).

A propulsion system (20) for an aircraft includes a fan (22), a core engine (24) configured for generating a gas flow (36) utilized to generate shaft power for driving the fan (22), a nacelle (48) that surrounds the core engine (24) and the fan (22), a condenser (58) where water in the gas flow (36) is condensed into a liquid form, an exhaust duct assembly (52) where the gas flow (36) is directed to the condenser (58), an ejector duct (60) where a portion of a bypass airflow (38) is thermally communicated with the condenser (58) to cool the gas flow (36), and an evaporator assembly (54), that is in thermal communication with the exhaust duct (60), where water recovered by the condenser (58) is heated to generate a steam flow (56) that is subsequently communicated to the core engine (24).

An aircraft propulsion system (20) includes a propulsion section (28) and a turbine engine (30). The propulsion section (28) includes an open propulsor rotor (32) and an open guide vane structure (34). The turbine engine (30) is configured to drive rotation of the open propulsor rotor (32). The turbine engine (30) includes an engine core (68), an inlet section (60), an exhaust section (65) and a flowpath (70). The engine core (68) includes a compressor section (61), a combustor section (62) and a turbine section (63) with the turbine section (63) disposed axially between the compressor section (61) and the open propulsor rotor (32). The inlet section (60) includes a flowpath inlet (72). The exhaust section (65) includes a flowpath exhaust (74). At least a portion of the exhaust section (65) is disposed axially between the open propulsor rotor (32) and the open guide vane structure (34). The flowpath (70) extends from the flowpath inlet (72), through the inlet section (60), the compressor section (61), the combustor section (62), the turbine section (63) and the exhaust section (65), to the flowpath exhaust (74).

A gas turbine engine (22) is provided that includes a compressor section (26), a combustor section (28), a turbine section (30), and an inspection port (74). The turbine section (30) includes a rotor stage (66A,B) and a stator vane stage (64A.B). A core gas path (36) extends through the compressor section (26), the combustor section (28), and the turbine section (30). The inspection port (74) is disposed within the turbine section (30). The inspection port is configured to provide access through a portion of the core gas path (36) that extends within the turbine section (30) and to a component disposed radially inside of the portion of the core gas path (36) that extends within the turbine section (30).

An aircraft propulsion system (20) includes a propulsion section (28) and a turbine engine (30). The propulsion section (28) includes an open propulsor rotor (32) and an open guide vane structure (34) disposed next to the open propulsor rotor (32). The turbine engine (30) is configured to drive rotation of the open propulsor rotor (32). The turbine engine (30) includes an engine core (68), an inlet section (60), an exhaust section (65) and a flowpath. The engine core (68) includes a compressor section (61), a combustor section (62) and a turbine section (63). The inlet section (60) includes a flowpath inlet (72). The exhaust section (65) includes a flowpath exhaust (74) and a mixer at the flowpath exhaust (74). The flowpath extends from the flowpath inlet (72), through the inlet section (60), the compressor section (61), the combustor section (62), the turbine section (63) and the exhaust section (65), to the flowpath exhaust (74). The mixer is configured to mix combustion products exhausted from the flowpath through the flowpath exhaust (74) with ambient air outside of the propulsion system.

A method and apparatus for applying a controlled thickness patch of repair material to a damaged surface adjacent a filleted surface (R) disposed between a case surface (15) and a strut (10) is disclosed. A spacer is positioned on the case surface (15) adjacent the filleted surface (R), with an outermost surface of the spacer (40) dimensioned to intersect the filleted surface (R) at a predetermined drop height (h) from the case surface (15). Repair material is applied to the damaged surface by screeding with a straight edge (50). The straight edge (50) is positioned with a first end against the outermost surface of the spacer (40) and the filleted surface (R) at the drop height (h), and a second end positioned against the strut (10) or a guide mask disposed thereon so as to form a controlled thickness patch of repair material to the strut (10) and filleted surface (R) when the repair material is screeded.

An engine system comprises a gas turbine engine (20) that has a spool (30,32) that includes a compressor rotor and a turbine, and a combustor (56) connected to receive compressed air from the compressor rotor. There is an electric engine motor (64) electrically connected to a battery (60) and coupled to drive the spool (30,32). A lubrication circuit (66) includes a bearing compartment (68), a scavenger line (76a) connected with the bearing compartment (68), and a mechanically-driven scavenger pump (74) operable to suction the bearing compartment (68) via the scavenger line (76a). A bearing compartment pump (80) is located in the scavenger line (76a) between the bearing compartment (68) and the scavenger pump (74). The bearing compartment pump (80) is electrically connected to the battery (60) and is operable to reduce pressure in the bearing compartment (68) independently of operation of the mechanically-driven scavenger pump (74).

An inspection method is provided during which a probe assembly (28) is arranged with a component (22). The probe assembly includes a conduit (34) and an inspection probe (36). The inspection probe includes a probe head (60) located within a bore (50) of the conduit. The probe head includes an ultrasonic transducer (70). The arranging of the probe assembly includes: moving the conduit to sealingly engage a distal end (46) of the conduit against a surface (52) of the component; and moving the inspection probe to abut the probe head against the surface of the component. An acoustic coupling media (40) is disposed within the bore contacting the probe head and the surface of the component. A characteristic of the component is determined using the probe assembly. The determining of the characteristic includes: generating an ultrasonic signal using the ultrasonic transducer; and directing the ultrasonic signal from the probe head, through the acoustic coupling media, into the component.

A method of and system (20) for inspecting a component (30) for an anomaly includes: performing a through-transmission ultrasonic (TTUT) inspection of a component that produces response signals; producing a C-scan map based on the response signals; performing an initial quality assessment of the C-scan map to determine if the C-scan map is acceptable or unacceptable, and if acceptable, performing a secondary quality assessment of the C-scan map that includes registering the C-scan map with a CAD file representing the component; producing a registered C-scan image using the C-scan map and the registered CAD file; analyzing the registered C-scan image to determine the presence of a potential anomaly in the component;for a potential anomaly determined as being present, classifying the potential anomaly as relevant or irrelevant; and reporting any relevant anomaly present in the component.

A system for providing cooling air within a hybrid electric gas turbine engine (100) includes at least one cooling air tube (128) configured to provide high pressure turbine cooling air from a first location (116) to a second location (120) within the hybrid electric gas turbine engine (100). At least one throttling valve (102) each located on the at least one cooling air tube (128) configured to limit a flow of the high pressure turbine cooling air from the first location to the second location. At least one electromechanical actuator (130) each associated with the at least one throttling valve (102) configured to actuate the throttling valve (102) to a first flow level when the hybrid electric gas turbine engine (100) is in a first condition and to a second flow level when the hybrid electric gas turbine engine (100) is in a second condition responsive to control signals from an external source.

A method and system (20) of inspecting a component (30) for an anomaly is provided that includes: performing a through-transmission ultrasonic (TTUT) inspection of a component, the TTUT inspection producing response signals; producing a C-scan map based on the response signals; performing a quality assessment of the C-scan map to determine if the C-scan map is acceptable or unacceptable; producing a registered C-scan image using a C-scan map determined to be acceptable; analyzing the registered C-scan image to determine the presence of a potential anomaly in the component; for a potential anomaly determined as being present, classifying the potential anomaly as a relevant anomaly or an irrelevant anomaly; and reporting any relevant anomaly present in the component.

A repair system repairs a part by pre-heating a surface of the part (22) before beginning the repair process. An inspection system coupled to the repair system identifies a region of interest having damage on the part (22). A pre-heating system (1600) is used to heat the surface of the part (22) in the region of interest using a uniform heating induction coil (1608) and a localized heating induction coil (1610) to generate heat. The surface of the part (20) is heated to a temperature. A metallic powder (1604) is deposited on the pre-heated surface and worked using a deep rolling process to repair the part (22).

A modulated air flow heat exchanger system in a bifurcation duct including an air oil cooler located proximate the bifurcation duct; at least one door positioned upstream of the air oil cooler proximate the bifurcation duct, wherein the at least one door controls a flow of a fan discharge airflow through the air oil cooler; and an actuator in operative communication with the at least one door.

A heat exchanger (10; 110; 210) includes a hollow fairing (14; 114; 214), a heat exchanger core (12; 112; 212) disposed in the hollow fairing (14; 114; 214) and comprising a plurality of fluid conduits (15; 115; 215), at least one external manifold (16, 18; 116, 118; 216, 218) disposed external to the hollow fairing (14; 114; 214) and fluidly coupled to the plurality of fluid conduits (15; 115; 215), and at least one compliant support structure (42, 44; 142, 144; 242; 244) connected to the at least one external manifold (16...218). The at least one compliant support structure (42...244) includes a first support leg (22A-1, 22B-1, 24A-1, 24B-1; 122A-1, 124A-1; 222A, 224A), a second support leg (22A-2, 22B-2, 24A-2, 24B-2; 122A-2, 124A-2; 222B, 224B), and a compliant member (46, 50, 52; 146, 150; 246, 248). The first support leg (22A-1...224A) extends from the at least one external manifold (16...218). The second support leg (22A-2...224B) extends from the hollow fairing (14; 114; 214) or an external structure (20; 220) of the heat exchanger (10; 110; 210) toward the first support leg (22A-1...224A) and is separated from the first support leg (22A-1...224A) by a first gap. The compliant member (46...248) is connected to one or both of the first support leg (22A-1...224A) and the second support leg (22A-2...224B).

A gas turbine engine assembly includes a core engine (20) that includes a core flow path where a core airflow (25) is compressed in a compressor section (24), communicated to a combustor section (56), mixed with fuel and ignited to generate an exhaust gas flow (74, 75) that is expanded through a turbine section (28). The turbine section (28) is coupled to drive the compressor section (24) through an engine drive shaft (88). A tap (64) is at a location up stream of the combustor section (56) for drawing a bleed airflow (66, 65). A bleed air heat exchanger (72) places the bleed airflow (66, 65) in thermal communication with an auxiliary flow (68) for heating the bleed airflow (66, 65). An exhaust heat exchanger (76) is configured to transfer thermal energy from the exhaust gas flow (75) into the bleed airflow (66, 65).

An annular flexible flange (14; 114; 214; 314; 414) is provided for connecting components (10, 12) of a gas turbine engine having different rates of thermal response to transient thermal events. The annular flexible flange (14...414) is disposed about an axis (A) and extends radially from an annular body. The annular flexible flange (14...414) includes a plurality of fastener flanges (18; 118; 218; 318; 418) spaced circumferentially about the flexible flange (14...414) and a plurality of flexible arms (42; 142; 242; 342; 442) connected to the annular body and connected to the plurality of fastener flanges (18...418). The plurality of flexible arms (42...442) are configured to flex in a radial direction. The plurality of fastener flanges (18...418) are separated from the annular body by a gap.

An apparatus is provided for a turbine engine (20). This turbine engine apparatus includes an airfoil (84), and the airfoil (84) includes a first end (88), a second end (90), a leading edge (92), a trailing edge (94), a first side (98), a second side (100) and a plurality of cooling holes (86). The leading edge (92), the trailing edge (94), the first side (98) and the second side (100) extend spanwise from the first end (88) to the second end (90). The first side (98) and the second side (100) extend longitudinally between and meet at the leading edge (92) and the trailing edge (94). The cooling holes (86) are located in the airfoil (84) according to a set of Cartesian coordinates of Table 1, and the set of Cartesian coordinates of Table 1 describe distances from a point of origin (104) on the airfoil (84) to the cooling holes (86).

An apparatus is provided for a turbine engine (20). This turbine engine apparatus includes an airfoil (84), and the airfoil (84) includes a first end (88), a second end (90), a leading edge (92), a trailing edge (94), a first side (98), a second side (100) and a plurality of cooling holes (86). The leading edge (92), the trailing edge (94), the first side (98) and the second side (100) extend spanwise from the first end (88) to the second end (90). The first side (98) and the second side (100) extend longitudinally between and meet at the leading edge (92) and the trailing edge (94). The cooling holes (86) are located in the airfoil (84) according to a set of Cartesian coordinates of Table 1, and the set of Cartesian coordinates of Table 1 describe distances from a point of origin (104) on the airfoil (84) to the cooling holes (86).

A repair system for a part (1340) for a gas turbine engine includes a deposition system (1308) that implements a friction stir additive manufacturing (FSAM) process. An inspection system coupled to the repair system identifies a region of interest having damage on the part (1340). A process for deposition of metallic deposition material onto the region of interest performed using the deposition system (1308). A rod (1324) of the metallic deposition material is moved using pressure exerted within the deposition system (1308) into a deposition zone (1328) of the region of interest. The region of interest is pre-heated prior to the pressure being exerted. Frictional heat is generated when the rod (1324) contacts the deposition zone (1328). Parameters to control the components within the deposition system (1308) are determined using a function and a depth of the layers (1346) of the metallic deposition material determined for enabling the repair process.

A cooling passage (18) forming process including providing an article comprising at least one wall (26) having an exterior surface (32); forming at least one cooling passage through the at least one wall extending through the exterior surface; coating the exterior surface of the article with a coating system (68); and removing (60) the coating system located within the cooling passage.

A gas turbine engine case structure includes the unitary combination of: a case wall having an inner surface and an outer surface; at least one fuel inlet (264); circumferentially-distributed fuel injectors protruding inward from the case wall and having an outlet; and a fuel plenum fluidically between the at least one fuel inlet (264) and the fuel injectors and configured so that each inlet (264) of the at least one inlet (264) is coupled to feed multiple of the fuel injectors. The fuel plenum alternatingly passes forward and aft.

A ceramic matric composite (CMC) component (100) for a gas turbine engine, such as blade outer air seal (BOAS) (101), may be shielded from thermal stress. The CMC component (100) includes a first surface (110) configured for exposure to a hot gas stream (112), a second surface (120) configured for exposure to a cold gas stream (122), an edge surface (130) of the CMC component (100) disposed between and connecting the first (110) and second (120) surfaces. At least one coating layer (140) is disposed on the first surface (110) and wrapped over the edge surface (130) and the second surface (120), wherein the at least one coating layer (140) disposed on the second surface (120) has a lower thermal conductivity than the CMC component (100) so as to reduce a thermal gradient between the cooled first surface (110) and the heated second surface (120).

A method for repairing a gas turbine engine component (1000), comprising: providing at least one gas turbine engine component (1000) comprising a surface area having a portion comprising a first size and corrosion; removing the corrosion from the portion to form a portion free of corrosion and comprising a second size; treating the portion free of corrosion and comprising a second size to form a portion free of corrosion; disposing an amount of at least one anti-corrosion agent on the treated portion free of corrosion to form an anti-corrosion agent-coated treated portion free of corrosion; and removing at least a portion of the anti-corrosion agent from the anti-corrosion agent-coated treated portion free of corrosion to form a portion free of corrosion comprising the first size and a residual amount of the anti-corrosion agent.