17results about How to "Overcome or mitigate technical deficiencies" patented technology

A kind of turbine component transition state test import flow detection method, design utilizes the steady total pressure of turbine import before transition state starts, dynamic total pressure, the steady static pressure of turbine import before transition state starts, dynamic static pressure, and the dynamic total pressure of turbine import at time under transition state, dynamic static pressure, the corrected total pressure of turbine import at time under transition state, corrected static pressure are calculated to obtain, and then the Mach number of turbine import at time under transition state, flow coefficient are calculated, combined with the steady flow of turbine import before transition state starts, the steady flow of turbine import after transition state terminates, the conversion flow coefficient of turbine import before transition state starts, after transition state terminates is calculated, the flow coefficient of turbine import at time is calculated by steady dynamic mapping interpolation, finally the flow of turbine import at time is calculated.

The application relates to a honeycomb gasket sealing structure between a rotor wheel disc and a stator inner ring in an aero-engine, which comprises a stator inner ring, the inner side of which is connected with a honeycomb, and the stator inner ring can be assembled or integrally formed; a rotor wheel disc is arranged in the stator inner ring, the outer edge of the rotor wheel disc is provided with an annular tenon groove, and the rotor wheel disc is provided with an installation gap which is communicated with the annular tenon groove; a plurality of sector blocks are provided with tenon heads on the inner sides and are provided with gaskets on the outer edges; the tenon heads on the sector blocks are sequentially clamped into the annular tenon groove through the installation gap on the rotor wheel disc, are spliced into a whole ring structure, and the gaskets on the outer edges and the honeycomb on the inner side of the stator inner ring form a sealing structure; the sector blocks and the rotor wheel disc are circumferentially positioned through pins, wherein the heads of the pins are immersed into the side wall of the rotor wheel disc and are flush with the side wall of the rotor wheel disc, the top ends of the pins are hollow structures, the pins are outwardly expanded into horn shapes, and the top ends of the pins are flush with the side wall of the rotor wheel disc.

The application belongs to the technical field of folding drive design of folding wing of an airplane, and particularly relates to an integrated rotary drive type folding mechanism for folding wing of an airplane, which comprises a speed reducer, a driving cylinder, a fixing cylinder, a driving sleeve and a transmission cylinder, integrates folding, unfolding drive, locking and unlocking functions for the folding wing of the airplane, can complete the folding, unfolding, locking and unlocking process of the folding wing without the coordination of sensors, signal transmission and other devices, has good system stability, compact overall structure, small volume and is convenient to arrange at the folding position of the folding wing.

This application belongs to the field of airborne environmental control system design technology, specifically involving a parameter identification method for airborne environmental control systems based on Matlab and Amesim, including: Step 1, establishing a simulation model of the controlled object of the airborne environmental control system based on Amesim software; Step 2, setting input excitations, collecting the output response of the simulation model, and constructing parameter identification data and parameter verification data; Step 3, designing the mathematical transfer function of the controlled object of the airborne environmental control system; Step 4, based on the parameter identification data, using the Matlab parameter identification toolbox, performing parameter identification on the mathematical transfer function; Step 5, inputting the input excitations of the parameter verification data into the mathematical transfer function to obtain the output response of the mathematical transfer function; Step 6, comparing the output response of the parameter verification data with the output response of the mathematical transfer function to determine whether the parameter identification of the mathematical transfer function is effective.

A kind of afterburner in aero-engine, comprising: outer casing;Converging ring, be arranged in outer casing;Inner cone, be arranged in converging ring, hollow structure, rear end has multiple circumferentially distributed flame holes;Multiple support plates, are supported between converging ring, inner cone along the circumference;Precombustion chamber, be arranged in inner cone, for sandwich structure, inner wall is corrugated structure and has multiple air bleed holes, front end has oil injection port, rear end is shaped with multiple flame pipes;Each flame pipe is connected to inner cone, and the outlet of each flame pipe is communicated with a flame hole corresponding;Ignition electrode, pass through outer casing, converging ring, inner cone and a support plate are arranged, and ignition end is inserted into precombustion chamber;Air bleed pipe, pass through outer casing, converging ring, inner cone and a support plate are arranged, and outlet end is communicated to the sandwich of precombustion chamber;Oil injection nozzle, pass through outer casing, converging ring, inner cone and a support plate are arranged, and oil injection end is inserted into oil injection port.

The application relates to a surface heat flow measuring system for a silicon carbide composite material component of an aircraft engine tail nozzle, which comprises a countersunk nail connected to the silicon carbide composite material component, the head surface of the countersunk nail is flush with the surface of the silicon carbide composite material component, and the countersunk nail has a stepped through hole; the small hole of the stepped through hole extends to the head of the countersunk nail, and the large hole of the stepped through hole extends to the top end of the countersunk nail; a coaxial thermocouple is arranged at the head of the small hole of the stepped through hole; a threaded sleeve is sleeved on the lead wire of the coaxial thermocouple, screwed in the large hole of the stepped through hole, and abuts against the head of the coaxial thermocouple, so that the head of the coaxial thermocouple is flush with the surface of the head of the countersunk nail; a high-speed collector is connected to the coaxial thermocouple, collects the voltage signal of the measuring temperature of the coaxial thermocouple, and obtains the measuring temperature of the coaxial thermocouple; and a host computer is connected to the high-speed collector, takes the measuring temperature of the coaxial thermocouple as a boundary condition, and obtains the surface heat flow of the silicon carbide composite material component.

The application belongs to the technical field of bearing impeller type radial oil collecting device design in an aero-engine, and particularly relates to a bearing impeller type radial oil collecting device blade height design method, which comprises the following steps: making a sliding oil injection line L in the direction of an oil nozzle nozzle; taking the intersection of the sliding oil injection line and the outer wall of the oil collecting ring as a sliding oil movement limit point A; taking a sliding oil movement starting point B on the sliding oil injection line L from the sliding oil movement limit point A in a reverse direction, and the length of the sliding oil movement limit point A and the sliding oil movement starting point B is I, wherein I=V*T; T=60 / (M*N); wherein V is the sliding oil injection speed, m / s; T is the time difference of adjacent blades rotating to the same position, s; M is the number of blades; N is the aero-engine rotating speed, r / min; drawing a circle with the midpoint O of the oil collecting ring as the center and passing through the sliding oil movement starting point B, taking the radius R2 of the circle as the outer radius of the blade; calculating the height h of the blade, h=R2-R1; wherein R1 is the inner radius of the blade.

The invention belongs to the technical field of aero-engine combustion chamber design, and particularly relates to an aero-engine combustion chamber and an applicable fuel spray rod thereof, an integrated multi-point injection oil nozzle design is adopted, multi-degree-of-freedom control can be carried out on the flow of a plurality of fuel nozzles of a single oil nozzle, and a bluff body flame stabilizing combustion organization mode is combined, so that the combustion efficiency of the aero-engine combustion chamber is improved. The turbine given design requirements such as the combustion chamber outlet temperature profile shape and the hot spot temperature are met through the partition oil control function of the oil nozzle.

This application belongs to the field of high-temperature environmental testing heating design technology, specifically relating to a high-temperature environmental testing heating system and method. The high-temperature environmental testing heating system includes: a quartz lamp heater and a graphite heater; the quartz lamp heater includes a quartz lamp, a connecting plate, and connecting rods; multiple quartz lamps are arranged horizontally; the connecting plate is horizontally positioned above each quartz lamp; multiple sets of connecting rods connect to each quartz lamp and the bottom of the connecting plate at both ends; the graphite heater is rectangular, horizontally positioned below the quartz lamp heater, and has strip-shaped holes facing each quartz lamp for passage, thereby enabling the high-temperature environmental testing heating system to have: a high-temperature heating state where each quartz lamp is above the graphite heater; and a low-temperature heating state where each quartz lamp passes through the strip-shaped holes and reaches the bottom of the graphite heater.

This application belongs to the field of radar signal sorting technology, specifically relating to a method for determining the polarization ellipse parameters of a radar signal, including: calculating the polarization response under multiple polar angles, where i = 1, 2, ..., N are sampling points of the polar angles; if θ0 = 0 and axial ratio θ0 = 0, then determine the polarization ellipse θ0 = 0, where θ0 = 0; otherwise: calculate the potential estimate of the polarization ellipse tilt angle, where the polarization response ρ(θ0) corresponding to the potential estimate of the polarization ellipse tilt angle is calculated. t Furthermore, a radar signal sorting method is provided, comprising: determining the radar signal polarization ellipse parameters based on the aforementioned radar signal polarization ellipse parameter determination method; and using radar signal polarization as a feature dimension for sorting radar signals.

The application belongs to the technical field of high-pressure turbine working blade film hole measurement of an aero-engine, and particularly relates to a standard piece for calibrating a high-pressure turbine working blade film hole measurement device and a method, wherein the standard piece comprises: a first base plate having a standard blade basin curved surface, the standard blade basin curved surface having 1st-5th column standard film holes; a second base plate vertically arranged with the first base plate, having a standard blade back curved surface, the standard blade back curved surface having 6th-10th column standard film holes; and a third base plate vertically arranged with the first base plate and the second base plate, having a standard leading edge curved surface, the standard leading edge curved surface having 11th-15th column standard film holes; each column of standard film holes has six standard film holes, which are perpendicular to the plane of the base plate; the horizontal spacing of the 1st-5th column standard film holes, the 6th-10th column standard film holes and the 11th-15th column standard film holes is 4.5 mm, and the longitudinal spacing is 8 mm.

A method for simulating statically indeterminate boundary of static test of a component with connecting joint parts, comprising: performing finite element analysis of actual state of the component; performing finite element analysis of each connecting joint hinge constraint state on the boundary of the component; if the deviation of each connecting joint node force and element strain on the main force transmission path in the actual state of the component and the connecting joint hinge constraint state on the component exceeds the set deviation value, then: on the basis of the connecting joint hinge constraint state on the boundary of the component, releasing displacement constraint of part of the connecting joints in part direction according to relative deformation of each joint in the theoretical state of the component, and applying active load on part of the connecting joints, establishing an updated finite element model according to the new boundary condition, constructing an optimization model to correct the active load, until the deviation of each connecting joint node force and element strain on the main force transmission path is within the set deviation value range in the actual state of the component, and then obtaining the simulation of statically indeterminate boundary of static test of the component.

The application relates to a modeling method of an aircraft landing gear dynamics model, comprising the following steps: simulating the mechanical characteristics of a landing gear strut by using a rigid body unit; simulating the buffering characteristics of a landing gear tire by using a nonlinear spring damping unit, and constraining the movement of the landing gear tire by using a rigid body unit; simulating the buffering characteristics of a landing gear bumper by using a nonlinear spring damping unit, and constraining the movement of the landing gear bumper by using a rigid body unit; and constraining the movement of a torque arm by using an RBE2. In addition, the application relates to an aircraft landing simulation method, which is implemented based on the above-mentioned modeling method of the aircraft landing gear dynamics model.

The application belongs to the technical field of aero-engine combustion chamber design, and particularly relates to a fuel cooling structure of an aero-engine combustion chamber, which comprises a combustion chamber outer wall, a combustion chamber inner wall, a flame tube, a fuel passage, an internal partition plate, and a cavity partition plate.

A small-thrust engine fan stator assembly connection structure and method, wherein the assembly connection structure comprises: an outer casing made of composite material, the front end and the rear end having annular connection edges, both of which are covered with front and rear metal edges; a containing ring made of composite material, wound around the outer circumference of the front end of the outer casing; an inner casing front section made of aluminum alloy, arranged in the outer casing front section, the front end and the rear end having annular connection edges, wherein the annular connection edge of the front end is connected with the rearward bent inward protruding part of the front metal edge of the outer casing front end annular connection edge by rivets; an inner casing rear section made of aluminum alloy, arranged in the outer casing rear section, the front end having an annular connection edge; a stator vane ring made of aluminum alloy, arranged in the outer casing, between the inner casing front section and the inner casing rear section, the outer ring front end and the rear end having annular connection edges, connected with the annular connection edges of the rear end of the inner casing rear section and the front end of the inner casing rear section by rivets, and matched with the outer casing with a small gap.

The application belongs to the technical field of inter-stage combustion chamber design of an aero-engine, and particularly relates to an aero-engine hydrogen-fueled inter-stage combustion chamber, which comprises a combustion chamber outer wall in the shape of a ring, a combustion chamber inner wall in the shape of a ring arranged in the combustion chamber outer wall, a plurality of guide vanes supported between the combustion chamber outer wall and the combustion chamber inner wall in a circumferential direction, each guide vane being hollow, the leeward side wall of each guide vane being provided with a row of leeward side injection holes distributed along the height of the guide vane, the windward side wall of each guide vane being provided with a row of windward side injection holes distributed along the height of the guide vane, and each guide vane being provided with a row of pilot injection holes distributed along the height of the guide vane and a pilot slot extending along the height of the guide vane, the pilot slot being located behind the windward side injection holes and being connected to the pilot injection holes, a plurality of hydrogen gas injection rods connected to a hydrogen gas source and penetrating through the combustion chamber outer wall and being inserted into each guide vane, and each hydrogen gas injection rod being provided with three rows of hydrogen gas nozzles on the side wall of the hydrogen gas injection rod and being inserted into the leeward side injection holes, the windward side injection holes and the pilot injection holes of the corresponding guide vane.

The application specifically relates to an assembly structure of a mid-pressure compressor rotor-stator casing of an aero-engine and a method thereof, wherein the assembly structure of the mid-pressure compressor rotor-stator casing of the aero-engine comprises: two-stage rotor outer casings which are integral ring structures, the rear end of the front-stage rotor outer casing is provided with a rear end stop groove, and the front end of the rear-stage rotor outer casing is provided with a front end stop groove; a plurality of stator sector blocks which comprise arc-shaped upper edge plates, arc-shaped lower edge plates and a plurality of stator blades; each stator sector block is arranged between the two rotor outer casings and is spliced to form an integral ring structure, wherein the two ends of each arc-shaped upper edge plate are positioned between the two-stage rotor outer casings through a stop opening, and an intermediate stop groove is arranged on one arc-shaped upper edge plate; a stop strip is arranged in the rear end stop groove, the front end stop groove and the intermediate stop groove; and a stator outer casing which is a horizontally split casing is arranged between the two-stage rotor outer casings and is used for pressing the arc-shaped upper edge plates and the stop strips thereon against the two-stage rotor outer casings.