RTX Corporation Patente

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.

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 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.

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).

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 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 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 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 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).

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).

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.

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).

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 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 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 gas turbine engine case structure includes the unitary combination of: a case wall (70) having an inner surface and an outer surface; a fuel injector (244) protruding inward from the case wall (70) and having an outlet (245); and a splash plate (270) having a first face and a second face. The fuel injector outlet (245) faces the second face. A support connects the splash plate (270) to the case wall (70) and the fuel injector (244). The support directs air to a gap between the splash plate (270) and the injector (244).

A mechanical system includes a plurality of wire harnesses (102, 104) each have a generally cylindrical outer periphery, and a protective outer coating (90) around at least one electrical wire (92). A clamp (106) includes at least two clamp halves (120, 122; 152) to form a cylindrical space for receiving and supporting each of the plurality of wire harnesses (102, 104). The clamp (106) has grommets (132) associated with each of the at least two clamp halves (120, 122). The grommets (132) are formed of a non-metallic material. The at least two clamp halves (120,122; 152) are secured together, and are secured to a static surface (110). There is an alignment feature (129, 130; 154, 160) to align the at least two clamp halves (120, 122; 152) relative to each other. A gas turbine engine (20) is also disclosed.

A fuel injector assembly for a gas turbine engine includes a swirler (64) and a fuel nozzle. The swirler (64) has an outer wall (80) having a SOW inner wall surface (140), an inner wall (82) having a SIW inner wall surface (120) and a SIW outer wall surface (122), an outer passage (86), and an inner passage (84). The SIW inner wall surface (120) includes first and second sections. The first section (120A) extends between a first airflow inlet (112) and the second section (120B), and the second section (120B) extends between the first section (120A) and an inner wall distal end (118). The first section (120A) tapers radially inward, and the second section (120B) extends parallel the center axis (70). The outer wall (80) circumscribes the inner wall (82) and extends axially along the center axis (70) to a distal outer wall end (138), and the distal end wall (118) is axially recessed within the swirler (64) from the distal outer wall end (138). The fuel nozzle projects into the inner passage (84).

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.

An assembly is provided for an aircraft propulsion system. This assembly includes a carrier (90), a first gear system (86), a second gear system (88) and a lock device (106). The first gear system (88) includes a sun gear (92), a first ring gear (94) and a plurality of first intermediate gears (96) between and meshed with the sun gear (92) and the first ring gear (94). Each of the first intermediate gears (96) is rotatably mounted to the carrier (90). The second gear system (88) includes a second ring gear (102) and a plurality of second intermediate gears (104) meshed with the second ring gear (102). Each of the second intermediate gears (104) is rotatably mounted to the carrier (90). The lock device (106) is configured to lock rotation of the carrier (90) about an axis (28) during a first mode. The lock device (106) is configured to unlock rotation of the carrier (90) about the axis (28) during a second mode.

A fluid injection system (200; 300; 400) for a gas turbine engine (110) may comprise a fluid injector (160) configured to inject a fluid into an exhaust flow exiting a turbine section (28) of the gas turbine engine (110). The fluid injector (160) may be coupled to a turbine exit guide vane (150) located at a forward end of an exhaust system of the gas turbine engine (110). The fluid may decrease a temperature of the exhaust flow exiting the turbine section (28) and/or increase a thrust of the gas turbine engine (110).

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.

An electronics pressurization system (100) includes a compressed air source (125) providing pressurized air, a manifold (115), and a case (105a; 105b) containing a first high-voltage electrical system, the case (105a; 105b) comprising a pressurizable vessel and a vent hole (106a, 106b, 106c, 106d; 106e). The compressed air source (125) is connected to the manifold (115), and the manifold (115) is connected to the case (105a; 105b). The compressed air source (125) provides a flow of pressurized air, via the manifold (115) to the case (105a; 105b) to maintain an air pressure in the case (105a; 105b) at a higher-than-ambient pressure.

A nut strip assembly is provided that includes a carrier and a plurality of nuts. The carrier includes a base, a plurality of first tabs and a plurality of second tabs. The base is configured with a plurality of apertures. Each of the first tabs is connected to and project out from a first side of the base. Each of the second tabs is connected to and projects out from a second side of the base. The nuts are secured to the carrier. Each of the nuts are aligned with a respective one of the apertures. Each of the nuts is captured between the base and a respective one of the first tabs. Each of the nuts captured between the base and a respective one of the second tabs.

An artificial diffraction inspection sample including a sample component having a component body with an exterior surface; and a complimentary flat bottom hole formed in the component body, the complimentary flat bottom hole comprising a complimentary bottom surface with a complimentary incidence angle, the complimentary bottom surface being rotated around an acoustic path vector; wherein the complimentary bottom of the complimentary flat bottom hole represents an axial-radial crack.

A system (20) for inspecting a gas turbine engine component (23) includes an inspection environment (22) operable to access feature information associated with a component design (36). The feature information includes a unique identifier (40) assigned to a respective geometric feature (38) of the component design (36). The environment (22) is operable to access one or more inspection criterion (47) associated with the unique identifier (40). The environment (22) is operable to access process information (45) obtained during manufacturing of a physical instance of the respective geometric feature (38) according to one or more manufacturing parameters (48). The environment (22) is operable to generate, using a machine learning model (50), a prediction of whether the physical instance of the respective geometric feature (38) meets the one or more inspection criterion based (47) on the process information (45) . A method for inspecting a component is also disclosed.

A gas turbine engine is provided and includes a first fan blade (201₁) including a suction surface, a second fan blade (201₂) comprising a pressure surface (215) and neighboring the first fan blade (201₁) and a throat region (220) interposed between the suction surface (216) of the first fan blade (201₁) and the pressure surface (215) of the second fan blade. The throat region (220) includes a passage throat (221) located at a minimum distance between the pressure and suction surfaces. The first and second fan blades are configured such that a pre-compression region is defined in the throat region (220) ahead of the passage throat. Each of the first and second fan blades includes a mean camber line (MCL) defining a flattened suction surface.

A coated substrate is described for use as part of a jet engine component which includes a substrate and a coating system. The substrate can be a ceramic matrix composite or a superalloy substrate. The coating system is applied to the substrate by thermal spraying such as APS. The coating system includes a layer containing as hafnium silicate (hafnon), zirconium silicate (zircon), a rare earth phosphate (REPO₄), rare earth oxides (RE₂O₃), alumina, an aluminosilicate, rare earth-stabilized zirconia, and HfO₂-SiO₂-rare earth (RE) oxide, hafnia (HfO₂) stabilized (partially or fully) by the addition of another component, zirconia (ZrO₂) coating stabilized (partially or fully) by the addition of another component, a rare earth zirconate (RE₂Zr₂O₇), a rare earth hafnate (RE₂Hf₂O₇), and combinations thereof, wherein RE is Sc, Y, La, Ce, Pr, Nd, Pm, Sm, Eu, Gd, Tb, Dy, Ho, Er, Tm, Yb, or Lu. To mitigate crack formation during amorphous-crystalline phase transformation, the layer further contains 0.1 to 10 wt.% of a metal oxide selected from Al₂O₃, SiO₂, Nb₂O₃, MgO, CaO, SrO, BaO, and combinations thereof, as a sintering agent.

A method of forming a heat exchanger includes the steps of forming an inlet header (102) having a hollow interior formed by a wall, forming heat transfer tubes (108) of a diameter smaller than a diameter of the inlet header (102) and extending away from the inner wall of the inlet header (102), the heat transfer tubes (108) having a hollow interior, with the hollow interior of the inlet header (102) being blocked from communication with the hollow interior of the heat transfer tubes (108), forming an outlet header having a hollow interior formed by a wall, the outlet header formed at an opposed end of the inlet header (102), and the hollow interior of the outlet header being in communication with the hollow interior of the heat transfer tubes (108), cutting access opening through the wall of the inlet header (102) at a location opposed to the heat transfer tubes (108), and cutting through an opposed side of the wall of the inlet header (102) to form orifices (140) to communicate the interior of the inlet header (102) to the interior of the heat transfer tubes (108), then closing the access openings. A heat exchanger arrangement is also disclosed.

A composite fan blade (58) includes a plurality of uniplies (74) stacked together to define the blade (58); and at least one braided (78) or woven (76) ply between two plies of the plurality of uniplies (74). Z-pins (96) can be incorporated into the fan blade (58. A method for making a composite fan blade is also disclosed.

A fuel nozzle thermal management system including a fan fluidly coupled with at least one of a fuel nozzle, a fuel manifold, an external fuel supply, and fuel drain lines proximate a combustor.

A heat exchanger fairing including a fairing wall comprising a flexible region configured to change at least one dimension responsive to thermal expansion or contraction of a heat exchanger.

A turbine engine rotor has a central shaft and a disk stack having a plurality of disks encircling the shaft. A seal (100) has a first member (102) having, in central axial cross-section: a rearwardly-open channel (240) receiving a portion (260) of one of the disks; a sleeve (232) extending rearward from the channel (240) and between the disk and the shaft; and a portion (230) extending radially inward from the sleeve (232) and having a forward surface (234). The seal (100) has a second member (101) having, in central axial cross-section: a forward-facing surface (116) contacting an aft-facing surface (224) of the shaft (130); and a rear-facing surface (109) contacting the forward surface (234) of said portion (230) extending radially inward from the sleeve (232).

A turbine engine rotor blade (76) is provided that has an airfoil (78) that includes a side wall (98) with exterior and interior wall surfaces. The airfoil (78) includes a first wall channel (104) that is defined by a base surface (110) that is a portion of the side wall interior surface and first and second interior wall surfaces (112, 114). A first interface surface (116) extends between the base surface (110) and the first interior wall surface (112). A second interface surface (118) extends between the base surface (110) and the second interior wall surface (114). A third interface surface (120) extends between the first interior wall surface (112) and the second interior wall surface (114). The base surface (110) includes base surface first and second segments (110A, 110B), and a peak interface surface (122) extending between the base surface first and second segments (110A, 110B). The first interface surface (116) is disposed closer to the side wall exterior surface (92) than both the second interface surface (118) and the peak interface surface (122).