RTX Corporation Patente

A heat exchanger plate 44 for provides heat transfer between a first flow 910 along a first flowpath 900 and a second flow 912 along a second flowpath 902. The heat exchanger plate 44 comprised a body 52 having: a first face 62 and a second face 64 opposite the first face 62; a leading edge 54 along the second flowpath 902 and a trailing edge 56 along the second flowpath 902; a proximal edge 60 having at least one inlet port 22 along the first flowpath 900 and at least one outlet port 24 along the first flowpath 900; and at least one passageway along the first flowpath 900. Along a proximal portion 302, the first face 62 and the second face 64 converge at a first angle. Along a distal portion 304, the first face 62 and the second face 64 converge at a second angle less than the first angle.

A gas turbine engine (20) includes a fan (42) delivering air into a bypass duct (13) defined between a nacelle (301) and an inner core housing (15). The inner core housing (15) receives a compressor section (24), a turbine section (28) and a combustor (56). The nacelle (301) has an inner periphery (141) with a forwardmost point, and receives an acoustic structure on the inner periphery (141) adjacent the forwardmost point. The acoustic structure is defined by a three dimensional array (100) of interconnected resonators (102), with the interconnected resonators (102) extending in a radial direction (R), a circumferential direction (C) and an axial direction (A) all defined about a rotational axis (A) of the engine (20), with the interconnected resonators (102) having a larger cross-sectional area central body (104), and six members (106, 108, 110, 112, 143; 198) connecting the central body (104) of the resonators (102) to adjacent resonators (102) at respective central bodies (104). A perforated face sheet (118) is inward of the three dimensional array (100) of interconnected resonators and an anti-icing system.

A fan exit guide vane with acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; a radially inner attachment region opposite spanwise from a radially outer attachment region; a span dimension extending between the radially inner attachment region and the radially outer attachment region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the fan exit guide vane; and an acoustic zone integrally formed within the fan exit guide vane.A gas turbine engine includes the fan exit guide vane with acoustic treatment. A process forms the fan exit guide vane with acoustic treatment.

A gas turbine engine includes a a fan case (108) surrounding a fan rotor, and a core engine (232) including a low pressure compressor. A rigid connection between the fan case (108) and the core engine (232) includes three triangular-frame connecting members (116) rigidly connected to the fan case (108) at a fan case connection point, and to the core engine (232) at a core engine connection point. The triangular-frame connecting members (116) each are defined by two rigid legs (118, 120) which extend between the fan case (108) and to the core engine (232), along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other. A plurality of non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise and are rigidly mounted to at the fan case (108) and/or the core engine (232).

A swirl recovery vane with acoustic treatment including a leading edge and a trailing edge opposite chordwise from the leading edge; an attachment region opposite spanwise from a tip region; a span dimension extending between the attachment region and the tip region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the swirl recovery vane; an acoustic zone integrally formed within the swirl recovery vane. A gas turbine engine includes the swirl recovery vane with acoustic treatment. A process forms the swirl recovery vane with acoustic treatment.

A gas turbine engine (20) has: a compressor (24); a combustor (56); a turbine (28); a gaspath (C) through the compressor (24), combustor (56) and turbine (28); a bleed flowpath (920) from the gaspath (C); an air-air heat exchanger (200) having a heat donor leg along the bleed flowpath (920) ; and a Ranque-Hilsch vortex tube (220) having an inlet (222) along the bleed flowpath (920) and having a cool air outlet (228) and hot air outlet (226). The cool air outlet (228) is connected to a cold air load.

A manufacturing method is provided. During this method, a preform component (60') for a turbine engine is provided that includes a substrate (74). A meter section (102) of a cooling aperture (64) is formed in the substrate (74'). An external coating (76',78') is applied over the substrate. At least a portion of the substrate (74') and the external coating (76',78') is scanned with an imaging system (156) to provide scan data indicative of an internal structure of the portion of the substrate (74') and the external coating (76',78'). A diffuser section of the cooling aperture is formed in the external coating (76',78') and the substrate (74') based on the scan data.

A component retaining assembly (65) includes a housing (62) including a circumferential slot (76). A component (64) is mated to the housing (62). The component (64) includes a first face (86) including at least two tabs (78). The at least two tabs (78) extending at outward from the first face (86) at least partially past a portion of the circumferential slot (76). A retaining ring (66) is disposed within the circumferential slot (76) and abuts the first face (86) of the component (64). The at least two tabs (78) overlap a portion of the retaining ring (66). A gas turbine engine (20) and a method are also disclosed.

A swirl recovery vane with structurally capable acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; an attachment region opposite spanwise from a tip region; a span dimension extending between the attachment region and the tip region; a chord dimension extending between the leading edge and trailing edge; a pressure side opposite a suction side of the swirl recovery vane; a structural feature formed within the swirl recovery vane extending at least one of spanwise through the swirl recovery vane between the attachment region and tip region and/or chordwise between the leading edge and the trailing edge; an acoustic receiver formed within the structural feature extending at least one of spanwise through the swirl recovery vane between the attachment region and the tip region or chordwise between the leading edge and the trailing edge; and an acoustic panel inserted into the acoustic receiver. A gas turbine engine includes the swirl recovery vane with structurally capable acoustic treatment. A process forms the gas turbine engine with the swirl recovery vane with structurally capable acoustic treatment.

A fan exit guide vane with acoustic treatment includes a leading edge and a trailing edge opposite chordwise from the leading edge; a radially inner attachment region opposite spanwise from a radially outer attachment region; a span dimension extending between the radially inner attachment region and the radially outer attachment region; a chord dimension extending between the leading edge and the trailing edge; a pressure side opposite a suction side of the fan exit guide vane; an acoustic panel receiver formed within the fan exit guide vane extending at least one of spanwise through the fan exit guide vane between the radially inner attachment region and the radially outer attachment region or chordwise between the leading edge and the trailing edge; and an acoustic panel inserted into the acoustic panel receiver. A fan exit guide vane system with acoustic treatment for a gas turbine engine includes fan exit guide vane with acoustic treatment. A gas turbine engine includes the fan exit guide vane system with acoustic treatment for a gas turbine engine. A process forms the fan exit guide vane with acoustic treatment.

A method of assembling a tube assembly (100) includes the steps of providing a saddle clamp (104) having a first clamp half (109) extending for less than 50 percent of a circumference of a tube (102), and a second clamp half (108) extending for more than 50 percent of the circumference of the tube (102), placing a tube (102) on the first clamp half (109), resting on a static structure (106), and applying a lubricant (122) to a laterally inner surface (117) of the second clamp half (108) and then bringing the second clamp half (108) onto the tube (102), and then driving a fastener (110) to secure the first and second clamp halves (108, 109). A gas turbine engine (20) is also disclosed.

A method of making a metallic component (114) includes mounting a polymeric scaffold (102) onto a mandrel (108), wherein the mandrel (108) includes one or more chucks (108a) configured to secure the polymeric scaffold (102) onto the mandrel (108) and wherein the mandrel (108) is configured to rotate the polymeric scaffold (102); depositing, with a second additive manufacturing technique using a material applicator (110), a component material (112) onto the polymeric scaffold (102) to form the metallic component (114) on the polymeric scaffold (102), wherein the component (114) has a shape that is the same as the polymeric scaffold (102); removing the polymeric scaffold (102); and the metallic component (114) deposited on the polymeric scaffold (102) from the mandrel (108); and removing the polymeric scaffold (102) from the metallic component (114).

A three-dimensional (3-D) composite structure includes a 3-D lattice structure (32) having a plurality of electrically insulative struts, a matrix phase (38) surrounding the 3-D lattice structure (32), first and second electrically conductive face sheets (34, 35) positioned on two faces of the 3-D lattice structure (32), and a plurality of electrically insulative containment sheets positioned on all faces of the 3-D lattice structure (32) that do not include the first and second face sheets (34, 35). The matrix phase (38) includes an electrorheological material. The first and second face sheets (34, 35) are positioned such that an electric potential applied between the first and second face sheets (34, 35) creates an electric field in the matrix phase (38) that causes a desired reversible alteration to the viscosity of the matrix phase (38). The first and second face sheets (34, 35) and the plurality of containment sheets are collectively configured to contain the matrix phase (38) within the 3-D lattice structure (32).

A method of operation is provided during which hydrogen fuel is produced using a reversible fuel cell system (32) onboard a vehicle while the vehicle is stationary and/or docked. The reversible fuel cell system (32) receives water and input electricity to produce the hydrogen fuel. The hydrogen fuel is stored onboard the vehicle. Output electricity is generated using the reversible fuel cell system (32) while the vehicle is moving. The reversible fuel cell system (32) receives the hydrogen fuel stored onboard the vehicle and air to generate the output electricity.

An assembly for an aircraft propulsion system (20) includes a propulsor rotor (42) and a guide vane structure (88). The guide vane structure (88) includes a plurality of guide vanes (90). The guide vane structure (88) is disposed axially next to the propulsor rotor (42). The guide vanes (90) include a first guide vane. The first guide vane includes a camber line (98), a leading edge (100), a trailing edge (102), an inner end (94), an outer end (96), at least one leading edge section (106) and a trailing edge section (108). The first guide vane extends longitudinally along the camber line (98) from the leading edge (100) to the trailing edge (102). The leading edge section (106) forms a respective portion of the leading edge (100) and is arranged radially along the trailing edge section (108) between the inner end (94) and the outer end (96). The trailing edge section (108) forms the trailing edge (102). The leading edge section (106) is configured to translate longitudinally along the camber line (98).

An assembly for an aircraft includes a first propulsion system (24B) and a second propulsion system (24A). A first open propulsor rotor (60) is configured to rotate in a first rotational direction about a first axis (46). A first engine core (102) includes a first rotating structure (116) with a first turbine rotor (107). The first rotating structure (116) is configured to rotate in the first rotational direction about the first axis (46) and drive rotation of the first open propulsor rotor (60) through a first geartrain (58). A second open propulsor rotor is configured to rotate in a second rotational direction about a second axis that is opposite the first rotational direction. A second engine core includes a second rotating structure with a second turbine rotor. A second rotating structure is configured to rotate in the first rotational direction about the second axis and drive rotation of the second open propulsor rotor through a second geartrain.

A space filler (210; 230; 240; 260) for forming a fibrous preform (200) may comprise an additively manufactured ceramic material. The additively manufactured ceramic material may define a plurality of pores (214; 234; 244; 264). A shape of the additively manufactured ceramic material may complement a shape of a void (204) formed by fibrous regions (202) of the fibrous preform (200).

A propulsion system (20) for an aircraft includes an open propulsor rotor (38), a turbine engine (32) and a thermal management system (36). The turbine engine (32) includes an engine core (82) and an engine flowpath (76). The engine core (82) includes a compressor section (71), a combustor section (72) and a turbine section (73). The engine flowpath (76) extends longitudinally though the compressor section (71), the combustor section (72) and the turbine section (73) from an airflow inlet (78) into the engine flowpath (76) to a combustion products exhaust (80) from the engine flowpath (76). The thermal management system (36) is configured to manage heat energy generated by the propulsion system (20) during operation of the turbine engine (32). The thermal management system (36) includes an electric boost compressor (114), a heat exchanger (116) and a system flowpath (110) disposed outside of the engine core (82). The system flowpath (110) extends longitudinally through the electric boost compressor (114) and the heat exchanger (116) from an airflow inlet (122) into the system flowpath (110) to an airflow exhaust (124) from the system flowpath (110).

The incorporation of non-structural ceramic plies to a CMC component (10) or preform is used to provide one or more passages (25) allows for cooling air (45) to pass between segregated cooling cavities. In various embodiments, this may include preforming additional layers of non-structural plies (20) to allow for machining of one or more metered passages (25), or the incorporation of a pre-formed passage or passages (25) into non-structural ply layups. The one or more passages (25) may be machined and/or preformed and may include cylindrical, racetrack, slot, conical or similar geometries. The one or more passage (25) may be used to pass cooling air (45) between cavities separated by sealing features (W-seal, feather seal, etc.). Bypassing discrete sealing features is beneficial to control and/or meter cooling air flow (45) between high-pressure and lower-pressure cavities.

Zusammenfassung wird geladen …

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).

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).

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 plastic encased mandrel includes an open cell carbonaceous foam that is shaped to correspond to a core of a ceramic matrix composite (CMC) part. A polyvinyl butyral (PVB) plastic is injection molded into the carbonaceous foam, before or after shaping, to encase the carbonaceous foam core and form an encased mandrel. The open cell carbonaceous foam may be silicon carbide (SiC) foam or reticulated vitreous carbon (RVC) foam. The foam core may be incorporated into the CMC part when SiC foam is used for better distribution of thermal and structural loads or the foam core may be removed from the CMC part when RVC foam is used, such as when a CMC vane incorporates a metal spar.

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 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.

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.

A propulsion system (20) for an aircraft includes a propulsor rotor (38), a powerplant (24), a nacelle (36) and a lockout system (100). The powerplant (24) is coupled to and configured to drive rotation of the propulsor rotor (38). An electrical system (86) for the powerplant (24) is configured to receive electrical power from an electrical power source (90). The nacelle (36) includes an access structure (92A) configured to move between a first position and a second position. The access structure (92A) at least partially covers the electrical system (86) when in the first position. The access structure (92A) is arranged in the first position for propulsion system operation. The access structure (92A) is operable to be arranged in the second position for propulsion system maintenance and/or propulsion system inspection. The lockout system (100) is configured to automatically trigger an operation to cut off the electrical power to the electrical system (86) from the electrical power source (90) when the access structure (92A) moves away from the first position.

The disclosure describes a ceramic matrix composite (CMC) component having a cooling cavity (130). The CMC component has a base (105) having a bottom inner surface (120) and a top outer surface (110), in which the base (105) is made from a plurality of ceramic fiber plies within a matrix. A cooling cavity (130) is provided within the base (105) in which the cooling cavity (130; 235) has an opening in the top outer surface (110) of the base (105). A cover plate (140) made from one or more ceramic fiber plies within a matrix is provided to close the opening of the cooling cavity (130). The cover plate (140) is bonded to the base (105) by simultaneous densification of the base (105) and cooling cavity (130).

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.

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 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.

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.

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).

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 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 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).

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.