87 results about "Turboprop" patented technology

The invention discloses a movable frame device for a high-power turboprop engine test bed and the engine test bed. The movable frame device comprises a movable frame framework, a front mounting seat and a counterweight system. The movable frame framework is arranged below the static frame device and used for being connected with the static frame device through a plurality of spring pieces arranged on the periphery of the movable frame framework, and then the rack system used for measuring the thrust of the turboprop engine is formed. The front mounting seat is used for being connected with an engine mounting rack for mounting a turboprop engine in a fast mounting manner so as to mount and support the turboprop engine to be measured. And the counterweight system is used for balancing the total weight of the turboprop engine and the engine mounting frame, so that the gravity center of the rack system is positioned in the center defined by the plurality of spring pieces. According to the movable frame device, the air inlet flow field of the propeller is not influenced by the size of the frame of the movable frame device, so that the engine performance is not influenced, the test result is high in accuracy, the test is safe and reliable, and the installation position of the propeller does not interfere with the structure of the movable frame device.

The invention relates to the technical field of aircraft power systems, and discloses a hybrid power management system suitable for a vertical take-off and landing aircraft and a control method. A lithium battery pack; a super capacitor; the power distribution module is electrically connected with the main turboprop engine, the lithium battery pack and the super capacitor and used for controlling energy flow among the main turboprop engine, the lithium battery pack and the super capacitor according to instructions. The control method comprises the steps that a staged power distribution strategy is executed according to judged flight stages such as vertical take-off and landing, horizontal cruise or emergency, instantaneous power is compensated through a super capacitor during vertical take-off and landing, and combined propulsion or energy recovery is decided according to key parameters such as the lift-drag ratio during horizontal cruise. According to the scheme, the vertical take-off and landing response and flight stability are improved, the endurance mileage is increased, and the intelligence and safety of the system are improved through self-adaptive control and redundancy design based on the ground digital twin model.

The invention discloses a lubricating oil system of an aviation turboprop engine. The lubricating oil system comprises a lubricating oil supercharging device, an oil-gas separation device, a lubricating oil return device, a lubricating oil filtering device, a lubricating oil monitoring device and a connecting guide pipe communicating all the devices. The lubricating oil supercharging device comprises a main lubricating oil pump, an auxiliary lubricating oil pump and a torque measuring pump; the oil-gas separation device comprises a ventilation oil tank, an air separator and an oil mist separator; the lubricating oil return device comprises a lubricating oil tank, a lubricating oil radiator, a middle and rear bearing oil well pump, a transmission box oil well pump and a direct-current electric lubricating oil pump; the lubricating oil filtering device comprises a lubricating oil filter and a middle-rear bearing oil return filter; the lubricating oil monitoring device comprises a temperature metal chip annunciator and a magnetic chip detection annunciator. Appropriate lubricating oil pressure and oil quantity are provided, lubrication and heat dissipation of friction surfaces of parts in an engine are improved, and the use reliability of the engine is improved.

The invention relates to the technical field of aero-engine vibration control, and discloses an on-line automatic balance system for a turboprop engine propeller rotor, which comprises a signal acquisition and analysis module, an automatic balance state monitoring and control module and an automatic balance actuator, the signal acquisition and analysis module is used for monitoring and acquiring an original vibration signal of the rotor system in real time to obtain vibration monitoring information, and transmitting the vibration monitoring information to the automatic balance state monitoring and control module; the automatic balance state monitoring and control module is used for receiving the vibration monitoring information and outputting a driving control instruction to the automatic balance actuator according to the vibration monitoring information; the driving control instruction comprises the moving direction and the stepping amount of the balancing weight; the automatic balance actuator comprises two balance weight discs, and the balance weight discs are used for moving according to the driving control instruction so as to carry out vibration suppression. According to the invention, the supporting stability and the assembling economy are improved.

The invention relates to the field of unmanned aerial vehicles, and discloses a tilting type composite thrust system of a small fixed-wing aircraft, comprising: a physical bearing and power unit module comprising wings and six groups of tilting thrust units on the wings, each group of unit being composed of a turboprop engine and a variable pitch propeller and generating thrust; the thrust vector adjusting module comprises six sets of electric tilting joints and drives the tilting thrust unit to rotate so as to adjust the thrust direction. The state sensing and feedback module is provided with an angle sensor for monitoring the tilting angle in real time; the cooperation and synchronization control module comprises a synchronization controller which is connected with the angle sensor and the electric tilting joints and controls the six sets of electric tilting joints to be converted in batches according to the monitored angle. According to the invention, the integrated electric tilting joint is arranged, and the joint contains a main tilting shaft for wide-angle mode switching and an auxiliary fine tuning shaft for small-range active compensation, so that two adjusting mechanisms with different functions are integrated in a single structure.

A propeller blade or airfoil for a turboprop engine made from composite material including a matrix-densified fibrous reinforcement, the propeller blade or airfoil including, in the direction of its span, a root and an aerodynamic profile. The fibrous reinforcement includes a fibrous preform having three-dimensional weaving with a root preform portion and an aerodynamic profile preform portion. The fibrous preform includes a separation delimiting a recess that forms a cavity extending both into the root and into the aerodynamic profile. A spar is present in the cavity, the spar including an aerodynamic profile shaping portion positioned in a first portion of the cavity and a root shaping portion positioned in a second portion of the cavity. The root has a rotationally symmetric shape.

Disclosed is an interconnection system (1) for an aircraft turbine engine, comprising a plate (2) that comprises rigid conduits (C) secured to the plate (2) and connecting at least two elements to one another and to at least one inlet and to at least one outlet, in order to allow oil to circulate.

The invention relates to the technical field of aero-engine overall performance and control, in particular to an acceleration rule design method based on a turboprop engine acceleration performance calculation model. According to the method, in the process of solving a performance calculation model of the turboprop engine, fuel flow iterative solving is carried out for limiting parameter errors, a pitch angle optimization method is matched, and implementation steps and a design process of turboprop engine acceleration law design are achieved. The calculation amount generated by frequently calling a turboprop engine performance calculation model is reduced, and the design efficiency of the turboprop engine acceleration law is remarkably improved. Meanwhile, by optimizing the pitch angle, it is ensured that the turboprop engine keeps safe and stable operation in the acceleration process, and the maximum performance potential of the turboprop engine is fully exerted.

The application belongs to the field of aviation medium-sized aircraft design, and relates to a large-angle glide performance calculation method of a fixed-wing turboprop aircraft. The method comprises the following steps: determining the combination characteristics of the power device of the turboprop aircraft according to the characteristics of the turboprop engine and the propeller; increasing the resistance and reducing the lift by using the double-open state of the spoiler, so as to realize large-angle glide; and obtaining the required large-angle glide performance of the aircraft by modeling and adjusting calculation.

An aircraft engine, has: a core having a compressor, a combustor, and a turbine disposed in serial flow communication, the turbine in driving engagement with a propeller; a nacelle extending around the core, the nacelle having a nacelle intake upstream of the core and a nacelle exhaust downstream of the core, the nacelle intake configured for receiving air from an environment outside the nacelle, the nacelle intake fluidly connected to the compressor for feeding air thereto, the nacelle exhaust configured to discharge combustion gases from the turbine into the environment outside the nacelle; and a blocking member movable between an open position in which the core is fluidly connected to the environment via both the nacelle intake and the nacelle exhaust, and a closed position in which the blocking member hinders fluid communication between the core and the environment via one of the nacelle intake and the nacelle exhaust.

This invention discloses a propeller flow rectification device and design method for testing high-power turboprop engines. The propeller flow rectification device includes a base, a mounting frame mounted on the base, a rectification ring for guiding and rectifying the propeller flow, and an adjustable tie rod for aligning the rectification ring. The rectification ring includes a rectification panel, a structural reinforcement on the outer wall of the rectification panel, a tie rod connecting seat on the structural reinforcement, and an outer panel on the structural reinforcement. The two ends of the adjustable tie rod are connected to the tie rod connecting seat and the mounting frame, respectively. Its ingenious structure ensures the overall flow field stability within the square test chamber. The alignment of the rectification ring can be achieved by adjusting the support length of the adjustable tie rod, eliminating the need to reinstall the engine and reducing the workload of alignment. By replacing the rectification ring, it can accommodate blades of different lengths, shortening the test cycle and improving test efficiency.

The invention discloses a medium-power turboprop engine transmission gear machining method which comprises gear rough machining and subsequent gear finish machining, and deformation restraining treatment is connected between the gear finish machining and the gear rough machining. The deformation inhibition treatment comprises subsection carburization, isothermal quenching and subzero treatment; the subsection carburizing is divided into a strong carburizing stage, a diffusion stage and a soaking stage; the isothermal quenching adopts a nitrate bath; the temperature of the subzero treatment is-70 DEG C to-90 DEG C. The problem that in the prior art, after gear machining is completed, operation of a turboprop engine is easily affected after storage and transportation links are conducted and then the gear is applied to the turboprop engine is solved.

The invention discloses a mounting rack for a plummer block test of a turboprop engine. The mounting rack comprises an underframe assembly, and a fixed main mounting section bracket, a moving main mounting section bracket and two groups of auxiliary mounting section brackets which are vertically supported on the underframe assembly. The fixed main installation joint support and the moving main installation joint support are used for being connected with two main installation joints on the two sides of the front end of an engine respectively, and the moving main installation joint support is arranged in a moving mode in the radial direction of the engine and used for compensating radial thermal expansion of the front end of the engine. The two sets of auxiliary installation joint supports are used for being connected with two auxiliary installation joints at the middle rear end of the engine correspondingly, and the two sets of auxiliary installation joint supports are arranged in a small-amplitude moving mode in the axial front-back direction, the radial inner-outer direction and the vertical up-down direction of the engine so as to be used for compensating axial and radial thermal expansion of the engine. The mounting rack can compensate the axial and radial thermal expansion amount of the engine, so that the additional load caused by the fact that four supporting points on the engine are not located on the same plane is eliminated, and the resonant frequency generated by the rack and the engine is shielded.

The invention discloses a waste liquid collecting system for a turboprop aircraft, and belongs to the field of aviation waste liquid collection, the waste liquid collecting system comprises an oil collecting tank, an oil drainage device, an exhaust device, a liquid level display device and a liquid drainage pipeline system, and the pipeline system is connected with the oil collecting tank and an engine; the oil drainage device is arranged at the bottom of the oil collecting tank and communicated with the outside of the fuselage, the liquid level display device penetrates through the aircraft skin to form an external visible window, and the exhaust device is communicated with the oil collecting tank. The method has the effects of improving the waste liquid collection efficiency and reducing the cost and the operation risk.

This invention discloses a heat-insulating and pressure-resistant engine vibration damper mounting structure, including a rear vibration damper. The rear vibration damper includes a rubber assembly one, a rubber assembly two, a main support assembly, and a mounting bracket assembly. One end of the mounting bracket assembly is inserted into the main support assembly. A main bolt passes through the rubber assembly two, the main support assembly, and the rubber assembly one in sequence and is then locked with a nut. The main bolt passing through the main support assembly is connected to one end of the mounting bracket assembly via a ball bearing one. The mounting bracket assembly includes a mounting bracket frame, and an engine connecting arm is provided on the turboprop engine. The heat-insulating connection method involves placing a high-strength heat insulation sheet between the engine connecting arm and the mounting bracket frame. During connection, a screw passes through the engine connecting arm, the high-strength heat insulation sheet, and the mounting bracket frame and is then locked with a locking nut, thereby connecting the turboprop engine to the rear vibration damper. After connection, the engine connecting arm and the mounting bracket frame are separated by the high-strength heat insulation sheet.

The application discloses a mounting rack for turbo-prop engine shaft table test, which comprises a chassis assembly, a fixed main mounting section support, a stringing main mounting section support and two groups of auxiliary mounting section supports vertically arranged on the chassis assembly. The fixed main mounting section support and the stringing main mounting section support are used for being connected with two main mounting sections on the front end of the engine respectively, and the stringing main mounting section support is arranged along the radial direction of the engine to compensate the thermal expansion in the radial direction of the front end of the engine. The two groups of auxiliary mounting section supports are used for being connected with two auxiliary mounting sections on the middle and rear end of the engine respectively, and are arranged along the axial direction, the radial direction and the vertical direction of the engine to compensate the thermal expansion in the axial direction and the radial direction of the engine. The mounting rack can compensate the thermal expansion in the axial direction and the radial direction of the engine, so as to eliminate the additional load caused by the four supporting points of the engine not being on the same plane and shield the resonance frequency generated by the test bench and the engine.

The invention discloses an oil supply system of an aviation turboprop engine and a control method.The oil supply system of the aviation turboprop engine comprises an oil tank part, an oil valve assembly, a heating assembly, a pressure relief assembly and a cooling assembly, the oil tank part is used for providing fuel oil for the engine, and the oil tank part comprises a fuel oil pump filter assembly; the invention discloses an oil supply control method of an aviation turboprop engine. The oil supply control method comprises the following steps: S1, fuel oil storage and conveying detection; s2, automatically releasing high pressure; s3, high-temperature fuel oil is cooled and conveyed; and S4, guaranteeing low-temperature fuel oil heating. Through the heating assembly, when the reciprocating sliding part reciprocates, heat flow in an engine cabin is introduced into the circulating part, fuel oil in the oil tank part is heated through heat circulating flow, the problems of viscosity and poor fluidity of the fuel oil caused by low temperature are effectively solved, pipeline blockage caused by viscosity rise of the fuel oil at the low temperature is avoided, and the service life of the fuel oil is prolonged. Smooth fuel oil conveying is guaranteed.

This invention discloses a high-speed laminar flow airfoil, belonging to the field of aircraft aerodynamic design. The airfoil has a maximum relative thickness of 14%C, a maximum relative thickness location of 41.66%C, a maximum relative camber of 2.14%C, and a maximum relative camber location of 79.06%C, where C is the chord length. The airfoil has a flat upper surface and a concave structure at the trailing edge of the lower surface. Its operating Mach number range is 0.2–0.7, and its operating Reynolds number range is 6 × 10⁻⁶. 6 ~20×10 6 Under design conditions, the laminar flow coverage of the upper surface of the airfoil reaches 50% of the chord length, at Mach 0.7 and Reynolds number 20 × 10⁻⁶. 6 Under these conditions, the maximum lift-to-drag ratio reaches 150, at Mach 0.2 and Reynolds number 6 × 10⁻⁶. 6 Under these conditions, the maximum lift coefficient reaches 1.774, and the stall angle of attack reaches 17°. This invention overcomes the technical challenges of low critical Mach number in high-lift airfoils and easy laminar transition in traditional laminar airfoils at high Mach numbers, achieving a balance between high-speed cruise and low-speed takeoff and landing performance, and is suitable for high-speed aircraft such as turboprop regional aircraft.

The invention relates to a twin-propeller turbine engine (101) for an aircraft, the turbine engine (101) having a main axis (X) and comprising:-a first propeller (S1),-a second propeller (S2),-a gas generator, the invention relates to a gas generator comprising a power transmission mechanism, a first shaft (104a) for rotating a first propeller (S1) about a main axis (X), and a second shaft (104b) for rotating a second propeller (S2) about the main axis (X), characterized in that the power transmission mechanism is a mechanical differential (160).

The invention relates to the technical field of fault diagnosis, and discloses a turboshaft turboprop engine main bearing fault sample generation method and device, and the method comprises the steps: obtaining a standard fault sample of a main bearing; generating a standard grey-scale map of the standard fault sample; training a fault sample generation model based on the standard grey-scale map and the standard fault label to obtain a target generator; obtaining a target fault label; and adopting a target generator to generate a plurality of target fault samples corresponding to the target fault tag. According to the invention, the standard fault sample carrying the label of the main bearing is obtained, the standard fault sample is converted into the grey-scale map, and the model is trained based on the grey-scale map and the fault label, so that the model learns the grey-scale map distribution characteristics of different fault categories. The target generator is adopted to generate the fault sample of the specified fault category, the physical authenticity of the fault is reserved through a semi-physical generation method, and effective expansion of the training data set is realized, so that the accuracy and generalization ability of main bearing fault diagnosis are improved.

De-icing method in a de-icing system (1) of a turboprop (100) comprising a fixed part (2) including an inverter (4) and an anti-icing control unit (7), a rotating part (3) including a rotary control unit (5) and a heating mat (9), and a rotary transformer (8) between the fixed and rotating parts, the method comprising upstream data transfers from the rotary control unit to the anti-icing control unit and downstream data transfers from the anti-icing control unit to the rotary control unit, wherein the method includes encoding steps: downstream data into frequency modulations of a supply voltage provided by the inverter around a predetermined nominal frequency; and upstream data into predetermined series of changes in an electrical power consumed by the rotary control unit.Figure for the abbreviation: [Fig.1].

An aircraft turbine propeller comprising a propeller fairing and a blade, the propeller fairing and the blade comprising a plurality of heating components (220) and a system (222) for supplying power to the plurality of heating components, the invention relates to a power supply system (290) mounted in a propeller fairing and configured to autonomously generate electrical energy by rotation of a turbine main shaft (216) driving the propeller in rotation, said power supply system comprising a motor (290) having: a rotating stator (230B) acting as an armature and fixed synchronously in rotation to the main shaft of the turbine; and a rotor (230A) which functions as an excitation inductor and is fixed to a high-speed auxiliary shaft (232) in a synchronous rotation manner, the rotation speed of the high-speed auxiliary shaft being greater than the rotation speed of the main shaft of the turbine.

The application discloses a kind of high-power turbofan engine lubricating oil interruption test system and method, including lubricating oil tank, reversing valve, U-shaped pipe section, suction pump and control assembly, lubricating oil tank is used to connect oil return stage, reversing valve is connected with U-shaped pipe section and lubricating oil tank respectively, for connecting outside atmosphere, to carry out oil circuit and air circuit switching, U-shaped pipe section is used to gather lubricating oil, suction pump is connected with U-shaped pipe section, for connecting oil supply stage, to respond time in the switching of reversing valve oil circuit to air circuit, suction lubricating oil in U-shaped pipe section, control assembly is electrically connected with reversing valve and suction pump respectively, for controlling reversing valve and suction pump work;Through lubricating oil tank, reversing valve, U-shaped pipe section, suction pump and control assembly cooperate and cooperate lubricating oil interruption test, compared with prior art, by passive "oil discharge" becomes active "oil suction", realizes complete oil, the accuracy of test result is high, repeatability is good, practicality is strong, is suitable for being widely promoted and application.

Lubricating oil is driven by the pressure of a lubricating oil pump to reach the position of a two-stage oil flinger through an in-shaft oil way channel, enters an oil flinging way of the two-stage oil flinger under the driving of the centrifugal force of the two-stage oil flinger rotating at a high speed, and is sprayed out through oil flinging holes in the oil flinger, and the lubricating oil is sprayed out through the oil flinging holes in the oil flinger. And the oil enters an oil receiving channel between the two-stage planetary gear and the sliding bearing, so that efficient and stable lubrication of the sliding bearing of the two-stage planetary gear reducer of the turboprop engine under high speed and heavy load is completed. According to the invention, the oil flinger lubricating mechanism is embedded into the in-shaft oil path system of the two-stage planetary gear, so that fixed-point lubrication of the sliding bearing is realized; the centrifugal acting driving force of the rotor structure on the lubricating oil is increased along with increasing of the rotating speed, the problem that the lubricating system affects the performance of the lubricating oil pump due to the fact that the requirements of a traditional lubricating system for lubricating and cooling are increased under the high-speed and heavy-load state of the turboprop engine is solved, and the lubricating efficiency and stability of the lubricating system are guaranteed.

The application discloses a multifunctional vibration isolator suitable for high-temperature environment application of a turboprop engine, comprising a mounting seat used for being connected with a mounting joint of the engine and provided with a U-shaped double lug at an upper end, a cavity framework nested on the U-shaped double lug, metal support plates connected at two sides of the cavity framework, a limiting shaft sleeve arranged between the cavity framework and the metal support plates, a metal wire mesh block filled in a filling space, and a shear bearing shaft with one end passing through a through hole in the middle of the metal support plate and abutting against a framework counterbore in the middle of the cavity framework and the other end fixed on a mounting frame of the engine through a tension bolt; wherein the pre-pressing bolt penetrates through the limiting shaft sleeve. The vibration isolator for the turboprop engine is simple in structure, closely assembled, and made of full metal material, and can adapt to severe temperature and corrosion environment, and is especially suitable for vibration isolation and mounting in a high-temperature environment above 200 DEG C at a mounting joint of the turboprop engine. Therefore, the application is suitable for wide promotion and application.

A system (34) for controlling the pitch setting of a propeller blade (10) of a turbine engine of an aircraft is disclosed, characterized in that it comprises: - a blade (10) comprising a blade (12) connected to a root (14); - a cup (58) comprising an annular wall (58a) extending around a pitch axis (A) of the blade (10) and a lower axial end closed by a bottom wall (58b); - a fixing ring (52) extending around the root (14) and located inside the cup (58); the system further comprises a safety element (110, 110') ensuring the retention of the root (14) with the cup (58), said at least one safety element (110, 110') having a substantially elongated shape and passing through an aligned hole (112, 114, 126) in the bottom wall (58b) of the cup (58) and in the free end (28) of the root (14).

The invention relates to the technical field of engine blade testing, in particular to a testing device for whole machine stress measurement of a turbine turboprop engine and a using method. A turbine turboprop engine whole machine stress measurement test device comprises a measurement assembly which comprises a strain gauge and a strain signal line group, the strain gauge is suitable for measuring the strain change of a turbine blade, the turbine blade is arranged in an engine exhaust casing, one end of the strain signal line group is connected with the strain gauge, and the other end of the strain signal line group extends into an exhaust protective cover; the cooling bin is arranged in the exhaust nozzle and provided with a containing space, the exhaust nozzle is connected with an engine exhaust casing, and the cooling bin comprises a first cooling layer; and the slip ring electricity leading device is arranged in the accommodating space. The invention provides a turbine turboprop engine whole machine stress measurement test device and a use method, and aims to solve the problems that a strain gauge is installed on a blade, a strain signal line group is led out from an air jet port of an engine, a slip ring electricity leading device is installed, the strain gauge is damaged due to high-temperature and high-pressure airflow impact, and the test is interrupted.

Interconnection system (1) for an aircraft turbomachine, comprising a plate (2) including rigid conduits (C) integral with the plate (2) providing the connection of at least two elements to each other and to at least one inlet and at least one outlet, for oil circulation. Figure for the abstract: [Fig 2]

The invention provides a turboprop aircraft engine air inlet channel design method with a foreign matter discharging function. The turboprop aircraft engine air inlet channel design method comprises the steps that firstly, the performance requirement of an air inlet channel is determined; 2, designing detailed parameters of the air inlet channel; 3, the theoretical appearance of the air inlet channel is determined; and 4, verifying the theoretical appearance of the air inlet channel. Compared with a traditional turboprop aircraft air inlet channel, the air inlet channel designed through the method is high in total pressure recovery coefficient and uniform in flow field, can meet the air inlet flow and pressure requirements of an engine, has the foreign matter discharging function and prevents foreign matter in air from entering the engine. On the premise of ensuring the static strength design, the weight of parts is further reduced, and the fuel consumption rate of an aircraft is reduced.

The invention discloses a turboprop engine rotor structure which comprises a shell used for wrapping a turboprop engine, a rotor assembly is arranged in the shell and is of a double-rotor structure, a rotating disc and blades are arranged on the rotor assembly, a limiting connecting assembly is arranged between the blades and the rotating disc, and the limiting connecting assembly is mainly composed of a mortise and tenon joint piece and a fixing piece. A plurality of mortises are formed in the rotating disc at equal intervals, extending roots are arranged at the bottom ends of the blades, tenon teeth are arranged at the bottom ends of the extending roots, and the blades are connected through clamping of the tenon teeth and the mortises. Multi-dimensional constraint is formed through a triple connecting structure of tenon tooth-mortise mechanical interlocking, fixed bolt pre-tightening and movable inserting rod limiting, the overall connecting rigidity and the pre-tightening force maintaining capacity of the blades and the rotating disc are remarkably improved, relative sliding and gap generation of the matching faces are effectively restrained, and the service life of the blades and the rotating disc is prolonged. And the risk of fatigue cracks caused by connection failure is reduced.