2001results about "Efficient propulsion technologies" patented technology

A power turbine blade for a gas turbine engine includes at least one of a third-stage blade, a fourth-stage blade, or a fifth-stage blade. The power turbine blade includes a platform and an airfoil that extends in a spanwise direction from the platform to a tip shroud. The airfoil has an external airfoil surface formed in conformance with multiple cross-section airfoil profiles described by a set of Cartesian coordinates.

The invention belongs to the technical field of aircraft design, and particularly relates to an integrated tilting power system integrated with hydrogen fuel cell motor electric control and an aircraft. Comprising a tilting mechanism connected with a machine body; the integrated motor and controller is fixedly connected with the tilting mechanism and located at the end, close to the upstream of the incoming flow, of the tilting mechanism; the tilting rotor wing is mounted at one end, close to the incoming flow upstream, of the integrated motor and controller; the hydrogen fuel cell is fixedly connected with the tilting mechanism and is positioned at one end, close to the downstream of the incoming flow, of the tilting mechanism; the nacelle shell wraps the tilting mechanism, the integrated motor, the controller and the hydrogen fuel cell; the nacelle outer heat dissipation fins wrap the nacelle in the circumferential direction; and the tilting mechanism, the integrated motor and controller and the liquid cooling radiating pipeline of the hydrogen fuel cell exchange heat with the radiating fins outside the nacelle.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircrafts driven by electric propulsion systems and in other types of vehicles. In some embodiments, a system of an aircraft is disclosed, configured to determine one or more desired commands for the aircraft, retrieve engine information for at least one electric propulsion unit (EPU) of a plurality of EPUs of the aircraft, wherein the engine information includes at least one time-based metric for temperature associated with the at least one EPU, generates control commands based on the received engine information, and controls effectors according to the generated control commands to meet the desired commands of the aircraft.

The invention belongs to the field of aero-engine anti-icing system design, and discloses an aero-engine anti-icing system and a design method. The system comprises an anti-icing component, a heat supply structure, a control assembly and a monitoring assembly, wherein the anti-icing component selects an air inlet support plate and a rectification cover hood; the heat supply structure introduces air from two air flows of the combustion chamber and conveys hot air through a pipeline and an air collection cavity; the control assembly adopts intermittent control, and system management and control are achieved through a pressure reducing valve and a digital electronic controller; the monitoring assembly comprises a pressure sensor and a thermocouple. The design method sequentially comprises the steps of anti-icing component and type determination, anti-icing structure design, control scheme design, simulation analysis, test verification and the like. A complete forward design path is provided, the anti-icing efficiency is improved, the structure is simplified, safe and stable operation of the engine is guaranteed, and the method can be widely applied to design and improvement of an anti-icing system of the aero-engine.

The invention provides a turbine blade with a cooling structure for a gas turbine, and relates to the technical field of gas turbine blade cooling, the turbine blade comprises an inner layer wall and an outer layer wall, the inner layer wall is a hollow pipe body extending out from the normal direction of a turbine rotating shaft, and the hollow part is used as a cold air flow channel and receives cooling air flow. The outer layer wall is arranged on the outer side of the inner layer wall in a sleeving mode at intervals, an annular runner is defined by the outer layer wall and the inner layer wall, a plurality of impact holes are formed in the inner layer wall, a plurality of air film holes are formed in the outer layer wall, and cooling air flow in the cold air flow channel enters the annular runner through the impact holes to impact and cool the outer layer wall and is attached to the outer layer wall after being jetted through the air film holes. And a protective layer is formed between the outer layer wall and high-temperature fuel gas to realize gas film cooling. The opposite surfaces of the inner layer wall and the outer layer wall are flow guide curved surfaces, the projection of the flow guide curved surfaces in the normal direction of the rotating shaft is a first curve, the projection of the flow guide curved surfaces in the tangential direction of the annular flow channel is a second curve, and the flow guide curved surfaces are modeled to restrain secondary contra-rotating vortexes generated when cooling airflow flows in the annular flow channel.

The invention is a flying object equipped with a loading part holding a delivery payload, and the loading part or the delivery payload is attached to an airframe so that the loading part or the delivery payload is tilted backward in the forward and backward direction in a landing state or hovering state. The loading part or delivery payload is attached to the airframe so that the loading part or delivery payload is substantially horizontal during cruising.

A cooled brush seal for use with a CMC component of a gas turbine engine having a hot side exposed to a hot gas path and an opposing side having a cavity for feeding a cooling flow to at least one film cooling hole. The cooled brush seal includes: an OD backing plate; an ID backing plate; a plurality of brush seal bristles sandwiched between the OD backing plate and the ID backing plate; a first hole extending through the OD backing plate, the brush seal bristles, and the ID backing plate; and a cover plate attached to an inner surface of the ID backing plate and including a second hole circumferentially offset from the first hole and forming a second passage into the cavity, wherein the ID backing plate and the cover plate form a circumferential cooling passage extending between the first hole and the second hole.

An aircraft propulsion system includes a gas turbine engine, a nacelle, and an electrical assembly. The gas turbine engine includes a rotational assembly and an inner fixed structure. The nacelle includes a nacelle body, a first bifurcation, and a second bifurcation. The nacelle body extends circumferentially about the gas turbine engine. The nacelle body forms an annular bypass duct between the nacelle body and the inner fixed structure. The first bifurcation and the second bifurcation extend between and connect the nacelle body and the inner fixed structure. The electrical assembly includes a motor-generator, a motor control unit, a motor-generator (MG) cooling system, and a motor control unit (MCU) cooling system. The motor-generator is coupled to the shaft. The motor control unit is electrically connected to the motor-generator. The MG cooling system includes a first heat exchanger disposed at the first bifurcation. The MCU cooling system includes a second heat exchanger disposed at the second bifurcation.

The invention relates to a method (400) for manufacturing a casing (102) of a turbine engine (10), the casing (102) comprising a skin (102a) having an annular shape about an axis (A) and comprising tubular vents (106) which project on an outer annular face (102e) of the skin (102a) and are oriented radially with respect to the axis, the method (400) comprising the following steps: - preparing (402) an annular skin made of a thermoplastic-based material, the skin comprising openings (102h) passing through in the radial direction, - preparing (404) tubular vents made of thermoplastic-based material, - attaching (406) the tubular vents to the skin, with the attachment step (406) comprising, for each vent (106): • welding (406a) a support (200) to the skin at one of the openings, and • welding (406b) a sleeve (202) to the support and / or the skin, with the assembly formed by the support and the sleeve forming the vent.

The invention discloses a thermal analysis method of a heterogeneous material combined component considering multi-scale contact thermal resistance, and belongs to the technical field of aero-engine thermal management. The method comprises the following steps: firstly, establishing a period calculation model of a CMC outer ring and metal casing combined member, secondly, setting anisotropy of CMC heat conduction according to geometric profile change of a gas outer ring, and finally, introducing an empirical relation expression of contact thermal resistance by setting equivalent heat conduction of a contact area between the CMC outer ring and the metal casing in CFD calculation. The difference between the two is explored by comparing the temperature field results of the combined component with or without the contact thermal resistance, and the influence of the contact thermal resistance on flowing heat exchange is explored. According to the method, by considering the multi-scale contact thermal resistance between different materials, the influence of the contact thermal resistance on the turbine flow heat exchange result is researched, meanwhile, the temperature field of the turbine component in CFD calculation is corrected, and a guidance basis is provided for improving the prediction precision of the temperature result of the gas turbine component under the complex gas thermal load.

Aspects of the present disclosure generally relate to systems and methods for flight control of aircrafts driven by electric propulsion systems and in other types of vehicles. In some embodiments, a flight control system of an aircraft is disclosed, configured to receive one or more signals to control movement of the aircraft, determine at least one flight condition of the aircraft, calculate at least one or more loads associated with the aircraft based on the determined at least one flight condition; determine a flight configuration to alleviate loads on one or more components of the aircraft based on the received one or more signals and the calculated one or more loads by predicting a distribution of loads associated with the aircraft, generate one or more effector commands based on the flight configuration; and actuate one or more aircraft effectors based on the one or more effector commands.

An electric aircraft includes: one or two or more electric propulsion systems each including a propeller or fan for propulsion disposed to contribute to a lift of a main wing and an electric motor that drives the propeller or fan; and a controller that adjusts the electric propulsion system on the basis of a relationship between a variable relating to an operating state of the electric propulsion system and an aerodynamic force generated on the main wing such that a total thrust by the electric propulsion systems or the aerodynamic force on the main wing has a predetermined value or falls within a predetermined range.

Aircraft capable of hovering comprising a fuselage at least one rotor that is rotatable with respect to the fuselage, electrical drive means adapted to rotate the rotor and electrically powered by a battery, and a cooling system of the battery. Said cooling system comprises, in turn, a first opening for the air to enter, second openings for the air to escape and a passage which places the first opening in fluidic communication with the second openings and within which the battery is placed. The cooling system further comprises a fan adapted to increase the kinetic energy of the air contained in the passage and which is operated when the forward speed of the aircraft with respect to the ground is lower than a speed threshold value and / or when the temperature of the battery exceeds a temperature threshold value.

The application discloses a kind of low thermal conductivity high reflectivity five-element co-doped ceramic laser protection coating and preparation method thereof, belong to high-energy laser protection technical field.The protective coating includes the adhesive layer and the reflective heat insulation layer of laminated arrangement, adhesive layer is arranged on the surface of base material, reflective heat insulation layer is arranged on the surface of adhesive layer;Wherein, adhesive layer is formed by spraying one of Ni-Al self-bonding powder, Ni-Mo-Al, NiCrAlY, CoCrAlY and NiCoCrAlY;Reflective heat insulation layer is the ceramic coating of five-element co-doped ZrO2, in ceramic coating, base element is ZrO2, doping component is:Y2O3, Gd2O3, Yb2O3, TiO2 And Ta2O5, after doping, material is single cubic phase structure at 25 ℃-1600 ℃.By multiple oxide doping to ZrO2, the coating can maintain cubic phase at room temperature while having lower thermal conductivity and higher reflectivity, can solve the problems of oxidation reflectivity reduction of metal material, mass loss of ablation material and insufficient heat insulation capacity of heat insulation material in laser protection field, prolong service life.

A device and related method that provides, but is not limited thereto, a two-phase heat transfer device with unique combination of enhanced evaporation and increased cooling capacity. An advantage associated with the device and method includes, but is not limited thereto, increased cooling capacity per unit area, controlled and optimized evaporation, prevention of boiling, and prevention of drying of the evaporator. An aspect associated with an approach may include, but is not limited thereto, using a non-wetting coating or structure to keep working fluid away from the spaces between elongated members of an evaporator and using a wetting coating or structure to form thin films of working fluid around the distal region of the elongated members. For example it can be used to cool a computer chip, a skin of a hypersonic flying object, parabolic solar collector, turbine or engine blade, or any other heat source that requires high heat flux.

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

The invention discloses a gradient coating preparation method combined with a gas phase aluminizing process, and relates to the technical field of thermal spraying, and the gradient coating preparation method comprises the following technological processes: S1, carrying out sand blasting treatment on a common gas turbine hot part material GH3536, and then carrying out transferred arc activation treatment on the sand-blasted surface by using vacuum plasma equipment; s2, Ni < 23 > Co < 20 > Cr < 12 > Al < 4 > Ta < 0.6 > Y powder is sprayed; s3, spraying of Ni < 23 > Co < 20 > Cr < 10 > Al < 0.6 > Y powder is conducted; s4, spraying of Ni20Cr10Al powder is carried out; s5, Ni5Al powder is sprayed; s6, a beta-NiAl layer containing an Al2O3 protective film is prepared on the surface of the coating and the surface of the part through the gas phase aluminizing technology; gradient coatings with different components are prepared through atmospheric plasma spraying, and finally a beta-NiAl layer containing an Al2O3 protective film is formed on the surfaces of the gradient coatings through gas phase aluminizing.

An aircraft includes an aircraft heat source; a propulsion system including an electric propulsion engine, the electric propulsion engine including an electric motor and a fan rotatable by the electric motor, the electric propulsion engine further defining a fan air flowpath; a thermal management system including a heat source exchanger in thermal communication with the aircraft heat source, a heat sink exchanger in thermal communication with the fan air flowpath of the electric propulsion engine, and a thermal distribution bus extending from the heat source exchanger to the heat sink exchanger; and a control system operably connected to the thermal management system for selectively thermally coupling the heat sink exchanger with the heat source exchanger.

The invention provides a method for prolonging the high-temperature service life of a thermal barrier coating of a hydrogen-doped gas turbine. The method comprises the steps that Ni, Co, Cr, Fe, Al, Hf, Si and Y are smelted and cast into master alloy according to the proportion; preparing the mother alloy into high-entropy alloy powder through an inert gas atomization process; high-entropy alloy powder is deposited on the surface of the high-temperature alloy substrate by adopting a high velocity oxy-fuel spraying process, and a bonding layer of the thermal barrier coating is obtained; and preparing a thermal barrier coating on the bonding layer by adopting a plasma spraying process, and carrying out vacuum heat treatment on the thermal barrier coating in advance. The high-entropy alloy bonding layer provided by the invention has better thermal stability, and can still keep strong bonding force between the ceramic layer and the bonding layer alloy after water quenching high-temperature thermal cycle service. And the transition group high-entropy alloy bonding layer is beneficial to formation of a thermal growth oxide layer, has an inhibition effect on diffusion of metal elements and coating degradation in the hydrogen-induced corrosion and high-temperature steam corrosion processes, and can prolong the high-temperature thermal shock service time of the hydrogen-doped gas turbine.

A brush seal for use with a ceramic matrix composite (CMC) component of a gas turbine engine may feed a cooling flow to the CMC component, which has a hot side configured for exposure to a hot gas path of the gas turbine engine and an opposing cold side having a cavity for feeding a cooling flow to at least one film cooling hole of the CMC component. The brush seal includes: an outer diameter (OD) backing plate; an inner diameter (ID) backing plate; a plurality of brush seal bristles sandwiched between the OD backing plate and the ID backing plate; at least one hole forming a passage extending through the OD backing plate, the brush seal bristles, and the ID backing plate; and a cover plate attached to an inner surface of the ID backing plate and configured to extend into the cavity of the CMC component.

A thermal management system for an aircraft engine includes a closed-loop fluid circuit including a coolant supply leg, a first electric power system leg, and a second electric power system leg. A coolant supply system is fluidly coupled to the coolant supply leg for providing a coolant fluid to the coolant supply leg, the first electric power system leg, and the second electric power system leg of the closed-loop fluid circuit. A flow control device fluidly couples the coolant supply system via the coolant supply leg, to the first electric power system leg, and to the second electric power system leg. A first heat source is in thermal communication with the coolant fluid via the first electric power system leg, and a second heat source is in thermal communication with the coolant fluid via the second electric power system leg.

The application aims to provide a single-motor driven multi-mode ornithopter simulating the flight posture of insects and birds, comprising a fuselage, a foldable wing mechanism, a flapping wing and ground walking mechanism, and a steering tail wing mechanism; the flapping wing and ground walking mechanism is a spatial five-link mechanism; in the flight state, the motor outputs power, the five-link mechanism drives the wing to complete the flapping action; in the ground walking state, the electromagnet connection hole is not electrified, the connecting rod driving wheel and the electromagnet right-angle connecting shaft are disconnected, the wing is in the folded state, and the rotation output of the motor is the rotation movement of the driving connecting rod driving wheel; the application integrates the connecting rod and the wheel, realizes the function of single-motor driving two modes, simplifies the mechanism complexity, reduces the manufacturing difficulty, lightens the ornithopter weight, and has good obstacle crossing ability and can cross narrow space.

A gas turbine engine is provided having a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (AHPCExit) in square inches and the turbine section having a drive turbine defining a drive turbine exit area (ADTExit) in square inches, the turbomachine further comprising a drive turbine shaft coupled to the drive turbine; wherein the gas turbine engine defines a maximum exhaust gas temperature (EGT) in degrees Celsius, a maximum drive turbine shaft torque (TOUT) in Newton meters, and a corrected specific power (CSP) in Newtons squared times degrees Celsius over meters squared, wherein the corrected specific power is determined as follows:(TOUTADTExit)2*EGTAHPCExit*1⁢0-1⁢1;wherein CSP is greater than 0.0001194×EGT2−0.103×EGT+22.14 and less than 0.0003294×EGT2−0.306×EGT+77.91; and wherein EGT is greater than 525 degrees Celsius and less than 1250 degrees Celsius.

A VTOL aircraft (300, 600) includes a lift propeller (312, 612) configured to provide lift during take-off, landing, and hover operations. The lift propellers (312, 612) are configured to stow close to the boom surface (323, 623) when not in use, for example, during a cruise flight, to minimize drag around the surface of the propellers (312, 612). The lift propellers (312, 612) are vertically displaceable relative to the booms (322, 622) when switched to a lift configuration to provide a greater spacing from the booms (322, 622) to improve lift efficiency and reduce unwanted vibrations. The lift propeller (312, 612) may have a fail-safe configuration that allows the propeller (312, 612) to passively move between a lift configuration and a stowed configuration, and to fully rotate even when fully retracted.

An airfoil assembly and engine component for a turbine engine having an engine core extending along an engine centerline, the airfoil assembly having an airfoil extending between a root and a tip in a radial direction, away from the engine centerline, to define a span length. The airfoil comprising a composite core and a set of skins defining at least a portion of an exterior surface of the airfoil. A shank extending in the radial direction from the root toward the engine centerline and having a flared cross-section.

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

A double-flow turbomachine has an exhaust casing through which a hot primary flow passes and surrounded by a cold secondary flow, and a cone carried by the exhaust casing. The exhaust casing has a hollow radial arm passing through the primary flow to convey air into the cone in order to form a cooling flow for components located in the cone. The cone has an outer wall and an inner wall which runs alongside the outer wall and is radially spaced therefrom to delimit a thermal insulation space through which the cooling flow passes from upstream to downstream before it is discharged through an outlet opening at the end of the cone. It has an inner cover surrounding a component to channel the cooling flow so that it runs along the component to cool it before passing through the thermal insulation space to be discharged there.

Carbide-based fuel assembly includes outer structural member of ceramic matrix composite material, the interior surface of which is lined in higher temperature regions with an insulation layer of porous refractory ceramic material. A continuous insulation layer extends the length of the fuel assembly or separate insulation layer sections have a thickness increasing step-wise along the length of the fuel assembly from upper (inlet) section towards bottom (outlet) section. Fuel element positioned inward of the insulation layer and between support meshes has a fuel composition including HALEU and the form of a plurality of individual elongated fuel bodies or one or more fuel monolith bodies containing coolant flow channels. Fuel assemblies are distributively arranged in a moderator block, with upper end of the outer structural member attached to an inlet for propellant and lower end of the outer structural member operatively interfaced with a nozzle forming a nuclear thermal propulsion reactor.

A fan section of an aeronautical propulsion system includes a fan rotor including seventeen blades to twenty blades and having a solidity strictly less than 1.0. The solidity is equal to a ratio between a chord at the blade tip and an inter-blade pitch at the blade tip. The fan section has a hub-tip ratio greater than or equal to 0.22 and less than or equal to 0.32, a pressure ratio greater than or equal to 1.05 and less than or equal to 1.5, and a peripheral speed at the blade tip greater than or equal to 260 m / s and less than or equal to 400 m / s. The pressure ratio and the peripheral speed are measured at cruising speed.