11 results about "Aero engine" patented technology

The invention discloses an aero-engine turbine blade reliability evaluation method based on fracture mechanics. The method comprises the following steps: S1, establishing a simplified model of a bladesection by taking a turbine blade as a research object, calculating a stress intensity factor of the turbine blade containing an initial I-type crack at a blade root by applying a finite element method, substituting the stress intensity factor into a shape factor calculation formula, and determining a shape factor parameter; S2, establishing a turbine blade reliability model based on the generalized stress intensity interference model; and S3, establishing a probability density function of load and fracture toughness, establishing a probability life model and a reliability model for the bladein combination with a Paris formula, and solving the model to obtain the change condition of the reliability of the turbine blade along with the life. According to the method, initial defects existing in the blade are considered, on the basis of the initial defects, the crack propagation life of the I-shaped cracks on the surface of the blade under the cyclic load is established, and life prediction and reliability evaluation of the blade can be more accurate.

The invention discloses a design method for GH4169 alloy ring preforming blank with a tapered section, which solves the problems that the blank material with a rectangular section has low utilization ratio, the subsequent machining time is long, and the streamline structure of a ring is incomplete in the past, and on the like. The design method adopts the basic ideas that proper rolling proportion is selected according to the principle of the unchanging volume and the structural characteristics of the ring with the tapered section during the plastic deformation of metal, so that the structural sizes of parts of the preforming blank are obtained. The design method is simple, can effectively improve the utilization ratio of material, reduces the working hours for subsequent machining, can keep the complete inner streamline of the ring after machining is finished, and is beneficial to improving the performance of an aircraft engine and prolonging the service life.

The invention discloses a reverse flow combustor and a double-wall elbow structure thereof and relates to the field of aero-engines. The double-wall elbow structure comprises an inner elbow wall and an outer elbow wall; in the flowing direction of fuel gas, the inner elbow wall is divided into an upper inner elbow wall section and a lower inner elbow wall section, cooling structures of the two regions are different, the upper inner elbow wall section is provided with multiple rows of diverging holes, besides the fact that certain included angles are formed between the diverging holes and the wall of the upper inner elbow wall section, certain deflection included angles are formed between the diverging holes and the flowing direction of the fuel gas, and compound angles are formed; the lower inner elbow wall section is provided with a plurality of rows of diverging holes, certain included angles are formed between the diverging holes and the wall of the lower inner elbow wall section, and deflection included angles are not formed; and the inner elbow wall is sleeved with the outer elbow wall, a middle cavity is formed between the inner elbow wall and the outer elbow wall, and impactholes are formed in a shell of the outer elbow wall. By means of the elbow structure, the peak wall temperature at the start section of the inner elbow wall and the wall temperature gradient of an elbow can be effectively reduced, and therefore the strength of the elbow is improved, and the service life of the elbow is prolonged.

The invention discloses a test method for testing a P-S-N curve of high-cycle fatigue of an aero-engine material by a small sample amount. The method comprises the steps that same probability distribution points of fatigue lives of same samples under different stress levels in respective parent distribution meet an expression of a formula (4) defined in the specification, and in the formula (4), nji is logarithmic life of the sample i under the jth stress level, mu j is a logarithmic life mean under the jth stress level, sigma j is a logarithmic life standard deviation under the jth stress level, and meanings are same when a subscript j is changed to k; the life sample points under the different stress levels have a one-to-one corresponding relationship; according to the corresponding relationship in the formula (4), the fatigue life on a stress level is equivalent to another stress level, so that sample information aggregation is realized; equivalent large samples are obtained througha small sample test; and according to equivalent large sample data, life distribution parameter statistics is performed. According to the method provided by the invention, based on a sample aggregation principle, fatigue life distribution can be obtained through a small amount of test data, and the relatively accurate P-S-N curve is obtained.

The invention belongs to the technical field of aero-engine design, and particularly relates to an interpolation method suitable for an aero-engine real-time model. The method comprises the followingsteps: firstly, recording an interpolation algorithm of a current rotating speed position, a current pressure ratio position and a current adjustable blade angle position and the slope; performing first interpolation according to the current rotating speed; obtaining four discrete points, carrying out second interpolation according to the current adjustable blade angle; obtaining two discrete points, and finally, carrying out third interpolation according to the current voltage ratio to obtain the flow and the efficiency of the current working point; recording the interpolation position and the slope of the previous period in the difference calculation process of the next working point, judging whether the position value and the slope value need to be updated or not. Repeated circulation caused by searching is reduced, the interpolation efficiency in a real-time model is improved, and the real-time model running speed is increased.

The invention discloses an aero-engine high-pressure fuel and grease heat radiator. The heat radiator comprises a double-core-body fuel and grease heat radiator assembly (1), a fuel safety bypass valve (2) and an oil safety bypass valve (3). The high-temperature low-pressure oil flows along a shell pass of the double-core-body fuel and grease heat radiator assembly (1), and a low-temperature high-pressure fuel flows along a tube pass of the double-core-body fuel and grease heat radiator assembly (1). Heat of the oil is taken away by the fuel through heat exchange, so that the high-temperatureoil is cooled. The fuel safety bypass valve (2) and the oil safety bypass valve (3) are only started in the case that an engine fuel system becomes abnormal or an engine oil system becomes abnormal orflowing resistance of the fuel or flowing resistance of the oil of a heat radiator core body is increased during working in alpine region, the fluid does not flow through the heat radiator core body,so that the use safety of an engine is ensured. According to the heat radiator, the structure of the double-core-body fuel and grease heat radiator assembly is mainly designed, so that the effects that the heat exchange efficiency is highest, the structure is the most compact, and the weight is lightest within a certain space and certain size range are achieved, and the use requirements of a newly-developed-model engine can be met.

The invention relates to a manufacturing method for a high-speed permanent magnetic rotor. By performing the processes of assembling, polishing, dynamic balance and the like, and setting reasonable parameters, it can be ensured that magnetic poles of the motor rotor are firmly bonded on a rotor iron core at a high speed of 40,000 rpm. The permanent magnetic rotor manufactured with the method can work under a condition that the rotational speed is 50,000 rpm for a long term, is successfully applied to power motors of digital electronic control systems of aircraftengines, and meets operating requirements of high rotational speeds of aviation motors.

Provided in the invention is a steady-state-temperature-field-included simulation test bench of an aero-engine combustor. The simulation test bench comprises a box body, a combustion chamber simulation device, a test device, an air supply system, a ventilation system, a cooling system and a control system. The ventilation system includes an air suction type exhaust hood installed inside the lowerbox body and a wall-mounted industrial ventilation fan installed at the top of the box body. The air supply system and the cooling system are installed inside the lower box body. The control system includes a controller installed inside the lower box body and a human-machine interaction device installed on the outer side of the box body; and the controller is connected with the wall-mounted industrial ventilation fan, the human-machine interaction device, the combustion chamber simulation device, the test device, the air supply system, and the cooling system respectively. The simulation test bench can be used for simulating the actual working condition of the aero-engine combustion chamber; and the enabling performance of the thermal barrier coating material is studied preliminarily beforerun trial examination of the ground track. The steady-state-temperature-field-included simulation test bench has advantages of low cost, short cycle and low risk and the like.

The invention relates to an aircraft engine test site digital bus control device. The aircraft engine test site digital bus control device comprises a control system, a communication interface, an output interface and a power supply interface, and the control system comprises a main control circuit board, a central processing unit, an output module, a power supply module, a communication module and a relay module; the power supply interface is connected with the main control circuit board; the communication interface is connected with an engine test bed; the central processing unit is connected with the relay module which is connected with the output interface, and the output interface is connected with engine test bed site test equipment through a cable. The aircraft engine test site digital bus control device can monitor the status of an equipment site in real time and has good performance of safety, stability and anti-interference, intelligence of the control device is achieved better, and the requirement for aircraft engine test equipment site control is met.

The invention discloses an expansion ring and sealing brush combined internal combustion wave rotor and belongs to the technical field of sealing of aero-engine combustors. The expansion ring and sealing brush combined internal combustion wave rotor is characterized by mainly comprising a wave rotor body, end caps, an internal expansion ring, an external expansion ring, sealing brushes and the like, wherein the internal expansion ring and the external expansion ring prevent high-pressure gases in the wave rotor body from outleakage; the sealing brushes can prevent the high-pressure gases from leakage to adjacent wave rotor channels; and the internal expansion ring, the external expansion ring and the sealing brushes are all made of an elastic wear-resistant material, thermal expansion of a system can be accommodated to a certain degree, and the wave rotor cannot get stuck in the rotating process. The combined sealing method can simultaneously be used for sealing other multi-directional leakage systems in the aviation field.