129 results about "Turboprop" patented technology

According to the turboshaft turboprop engine overall performance scheme design method meeting the multiple working conditions, an engine nonlinear equation set containing a plurality of design points is constructed through an aerodynamic thermal process equation and engine complete machine matching conditions, and cycle parameters of aerodynamic design points and working points and rotating speeds under other working conditions are iterated; the invention relates to a process and a method for enabling an engine to simultaneously meet the requirements of component matching, performance requirements and load limitation at a plurality of design points (20-30 points), namely a multi-design-point model modeling and performance matching method meeting multi-working-condition-point calculation. In the process of designing engine cycle parameters and control laws under a new working condition combination by adopting a multi-design-point model, whether physical rotating speeds, conversion rotating speeds and turbine front temperatures of residual working condition points exceed limits or not under the existing cycle parameters and control laws is checked in sequence; and in the process of replacing the original working condition in the multi-design-point model with the working condition if the working condition exceeds the limit, a three-item replacement principle method is adopted for the working condition points of the engine.

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 relates to the technical field of aero-engine speed reducer and transmission system tests, and discloses a transmission shaft loading device which comprises a force transmission assembly connected with the front end of a transmission shaft, and the front end of the transmission shaft is suitable for serving as a power input end; the radial loading assembly applies a radial load Fr in the z-axis direction to the transmission shaft through the force transmission assembly. The axial loading assembly applies an axial load Fx in the x-axis direction to the transmission shaft through the force transmission assembly; the radial loading assembly and the axial loading assembly apply a bending moment load My around the y-axis direction to the transmission shaft through the force transmission assembly. According to the technical scheme, the radial load, the axial load and the bending moment load borne by the transmission shaft can be effectively simulated. According to the transmission shaft loading device, the load states of a turboprop engine speed reducer and a tilting main speed reducer of a tilting rotorcraft transmission system in a horizontal working condition in actual work can be truly simulated, and when the speed reducer is subjected to a simulation operation test, the real performance of the speed reducer is obtained.

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 discloses a high-power turboprop engine preassembling device and a test run method, which are used for preassembling a high-power turboprop engine, and the high-power turboprop engine comprises an engine main body, a propeller and a test structure, the preassembling device comprises a preassembling frame used for supporting, a height adjusting structure arranged on the preassembling frame and an installing frame used for being detachably connected with the height adjusting structure, the preassembling frame is used for being installed in a test preparation room so as to preassemble the installing frame and the high-power turboprop engine, and the height adjusting structure is used for adjusting the vertical height of the installing frame when the installing frame is preassembled. The mounting rack is used for mounting the engine main body in a cantilever manner and mounting the propeller and the test structure on the engine main body after being connected with the height adjusting structure, and then hoisting the engine main body into a test room after being separated from the height adjusting structure; the preassembling requirement of the high-power turboprop engine is met, rapid reloading and preassembling of the high-power turboprop engine are achieved, and the utilization rate of a test bed is increased.

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.

A hybrid electric turboprop includes a propeller, a thermal engine, an electric motor, a gearbox, and an air inlet. The propeller is configured to rotate about a propeller axis and includes a plurality of blades extending radially outwardly from the propeller axis such that, as the propeller is rotated about the propeller axis, the plurality of blades define a substantially circular blade path area centered about the propeller axis and having a radius. The gearbox couples both the thermal engine and the electric motor to the propeller to enable each of the thermal engine and the electric motor to drive rotation of the propeller about the propeller axis. The air inlet defines an opening. A radial distance is defined from the propeller axis to a location within the opening. The radial distance is equal to or less than 60% of the radius.

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.

This application discloses a turboprop engine dual-rod timing coordinated control and disturbance-free switching system and method. The system includes: a power rod and a status rod; two absolute encoders for converting the angle input of the dual rods into encoder signals and transmitting them to a programmable logic controller (PLC); an FPGA module for receiving and converting control signals from the PLC into RVDT signals, which are then sent to the EEC and PEC, respectively; a host computer for presetting loading test lines, monitoring test data in real time, displaying test status and alarm information, and communicating with the PLC to write loading test line data to the PLC; a sensor group for collecting relevant engine parameters in real time; and a PLC for communicating with the absolute encoder, FPGA module, host computer, and sensor group, respectively. This application overcomes the risks of interference during mode switching, timing coordination issues, sudden changes in electrical signals, and response delays of abnormal protection mechanisms.

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.

The invention belongs to the field of flight performance design, and relates to a method for calculating the air-slip performance of a fixed-wing turboprop aircraft. The method comprises the steps that according to the characteristics of a turbine engine and a propeller, the combined characteristics of a propeller aircraft power device are determined, the influence of feathering of the propeller on the aerodynamic characteristics of an aircraft is considered, and the needed aircraft air-slip performance is obtained through modeling and adjustment calculation according to aerodynamic characteristic parameters and weight characteristic parameters of the aircraft.

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.

A turboprop engine blade or propeller airfoil made of composite material includes a fiber reinforcement densified by a matrix, the blade or propeller airfoil including along a longitudinal direction a root and an aerodynamic profile. The fiber reinforcement includes a fiber preform having a three-dimensional weave with a root preform part located in the root and an aerodynamic profile part located in the aerodynamic profile, the root preform and aerodynamic profile parts being connected to one another by three-dimensional weaving. The root preform part includes two woven skins wound around an insertion element.

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.

The invention discloses a turboprop engine test bench which comprises a static frame, a movable frame, a spring piece and a protection device, the static frame is used for being hung and fixed to a roof of a test plant, the movable frame is arranged below the static frame and used for installing a turboprop engine, and the spring piece is arranged between the static frame and the movable frame; the protection device comprises a first protection seat installed on the static frame and a second protection seat installed on the movable frame, a bearing groove is formed in the side face of the first protection seat, and the second protection seat comprises a supporting block which extends into the bearing groove from the side face of the first protection seat and has a preset gap relative to the bottom wall of the bearing groove. The supporting block is used for falling onto the bottom wall of the bearing groove when the spring piece loses efficacy and is supported and fixed through the bearing groove. According to the turboprop engine test bed, the distance between the movable frame and the static frame can be effectively reduced, the length of the spring piece is shortened, the bending moment and deformation borne by the spring piece in the test process are reduced, and the machining precision requirement of the spring piece is lowered.

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.

A system (34) for controlling the pitch of a propeller blade (10) of a turboprop engine of an aircraft is disclosed, characterized in that it comprises a cup (58) comprising an annular wall (58a) extending around an axis (A) serving as an axis of pitch setting of the blade, and a bottom wall (58b) configured to cooperate in a shape-matched manner with a free end (28) of said root (14), so as to prevent the cup (58) from rotating with respect to the root (14) around the axis (A), and a fixing ring (52, 52') configured to be mounted inside the cup (58) and to cooperate with the root (14) and with the annular wall (58a) of the cup (58), respectively, so as to ensure the axial retention of the root (14) in the cup (58).

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 present disclosure relates to a hybrid turboprop engine (100) for an aircraft, comprising: a gas generator (12) carried by a generator shaft (14); a free turbine (11) carried by a turbine shaft (13) and rotationally driven by the gas flow generated by the gas generator (12), the turbine shaft (13) being engaged with the main rotor (60) by a transmission unit (50), the transmission unit (50) comprising a first free wheel (51) oriented such that the main rotor (60) cannot drive the free turbine (11); and a reversible electric machine (30) engageable with the main rotor (60) via the transmission unit (50) to drive the main rotor (60) during an electric or hybrid mode of operation, the hybrid turboprop engine (100) comprising: a single oil pump (40) engageable with the transmission unit (50) to be selectively driven by the turbine shaft (13) or the electric machine (30) according to the mode of operation.