2019 Vol. 34, No. 7

Catalogue Cover
  Previous Issue | Next Issue 
Display Method:
Accurate prediction method of flow coefficient for 2D curved surface variable inlet
2019, 34(7): 1409-1415. doi: 10.13224/j.cnki.jasp.2019.07.001
Abstract PDF(1643KB) icon1114 icon799
Abstract:
In order to meet the flow requirements of Mach number range of modal translation (incoming Mach number 2.2-3.2) for the 2D curved surface variable inlet, in view of three adjustment schemes of cowl translation, rotation and translation + rotation, based on theoretical analysis and the flow field of the reference inlet, an accurate prediction method of flow coefficient was proposed, and verified by numerical calculation, and the overall performance was obtained. The results showed that their flow coefficients were exactly equal to the predicted values and did not require repeated trial-calculation, which met the design expectations and could be extended to the axisymmetric inlet. In relation to the reference inlet, the flow coefficient and compression efficiency increased simultaneously as the cowl moved forward. The total pressure recovery coefficient of the exit was equal and the pressure ratio was increased by 14.6% at incoming Mach number 2.5. Under the condition of reducing the same flow coefficient, the backward movement of the cowl made the pressure ratio and compression efficiency lower, the total pressure recovery coefficient of the exit was basically equal and the pressure ratio was reduced by 12.9% at incoming Mach number 2.5, while the rotation of the cowl further reduced the pressure ratio by 9.1%, and the performance of cowl back movement scheme was better.
High speed wind tunnel test of civil aviation turbine engine nacelle
2019, 34(7): 1416-1424. doi: 10.13224/j.cnki.jasp.2019.07.002
Abstract PDF(3975KB) icon1272 icon753
Abstract:
Research of three different civil aviation turbine engine nacelle profiles by high speed wind tunnel test and numerical simulation method was finished, and the static pressure distribution and induct performance of nacelle were obtained. By comparing the result of wind tunnel test with the numerical simulation, it showed that nacelle upside cowl would have separation when the mass flow rates reduced or the attack angle increased; the induct of downside would have separation when the mass flow rates increased; the numerical simulation result of static pressure distribution coincided with the wind tunnel result when there was no separation, but when mass flow rate reduced, the CFD results were separated before the test.
Flow field structures and density fluctuation characteristics around a blunted double cone with cooling jet
2019, 34(7): 1425-1462. doi: 10.13224/j.cnki.jasp.2019.07.003
Abstract PDF(4775KB) icon916 icon679
Abstract:
Large eddy simulation was utilized for the simulation of supersonic cooling jet flow filed around a blunted double cone, and the flow structures and density fluctuation characteristics were analyzed. The large eddy simulation method was based on the implicit sub-grid model, and the high-order flux limited compact scheme and explicit Runger-Kutta method were utilized for the spatial discretization and time advancing. The present simulation clearly captured the flow field discontinuity. The processes of flow becoming unstable and transiting to turbulence were predicted subtly. Besides, the characteristics of the flow field density fluctuation were obtained directly. By comparison with the none-jet state, it revealed that gas film formed by the supersonic jet can effectively cool the optical window, and the mixing layer formed by the main flow and jet flow became unstable, and transited to turbulence quickly, causing large scale turbulence structures and intense density fluctuation. In addition, the spatial development characteristics of cooling jet flow field around the whole model are obtained.
Correcting method of fighter intake/exhaust test data
2019, 34(7): 1433-1439. doi: 10.13224/j.cnki.jasp.2019.07.004
Abstract PDF(1471KB) icon829 icon731
Abstract:
In order to improve the accuracy of the fighter intake/exhaust test results, a method based on the principle of thrust and drag bookkeeping, which can be used in correcting the effects of inlet internal flow and labyrinth seal pressure difference, was presented. Two measures were used to increase the liability of the correcting method: total pressure measurement points were distributed by the flow field, particularly intensive on nonevent area; the total pressure was calculated with weighted average method. In order to verify the correcting method, a fighter intake/exhaust test was conducted in the FL-14 wind tunnel. The test results indicate that, the impact of correcting method on testing accuracy is within the index of GJB; the correcting method can only be used within the medium angle of attacks; with the use of internal flow effects correction, the lift slope and the lift itself decreases, while the drag increases.
Aerodynamic performance of shrouded coaxial anti-rotating rotor propulsion unit with ground effect
2019, 34(7): 1440-1449. doi: 10.13224/j.cnki.jasp.2019.07.005
Abstract PDF(3888KB) icon1132 icon773
Abstract:
In order to study the aerodynamic characteristics of shrouded coaxial rotor propulsion unit hovering near ground, experimental measurements were conducted to investigate the effects of rotational speed, rotor spacing and shroud-diffuser-height from ground. CFD simulations were performed to analyze the velocity field and pressure field distribution around propulsion unit, and the flow-field characteristic of ground effect for shrouded coaxial rotor propulsion unit was studied. Results showed that, when the distance from shroud diffuser to ground was less than two rotor radii, the ground effect began to work; when the distance was less than one rotor radius, the ground effect was very significant, and figure of merit (FM) increased by about 30% at most. As the distance decreased, the shroud thrust decreased due to the declining inflow velocity leading to the decrease of the peak pressure at the shroud lip. The rotor spacing had little effect on ground effect. The thrust and counter-torque of shroud and rotors were approximately proportional to the square of rotor speed.
Experiment on flight performance of dragonfly during take-off
2019, 34(7): 1450-1458. doi: 10.13224/j.cnki.jasp.2019.07.006
Abstract PDF(4213KB) icon1078 icon778
Abstract:
In view of the experimental observation of the free takeoff process and the quasi free take-off process of dragonflies, two high-speed cameras were photographed with each other perpendicular to the optical axis. The motion parameters of the dragonfly body and wings during the two take-off processes were accurately captured by feature point matching and three-dimensional reconstruction, and the dynamic analysis and comparison were carried out. The experimental results showed that the dragonfly adopted synchronous vibration wing during the free take-off process and gradually transformed into asynchronous vibration wing (around 110 degrees), the maximum instantaneous vertical acceleration could reach 20m/s2; with use of asynchronous vibration wing (phase difference of 180 degrees) in the course of quasi free take-off, the phase difference gradually decreased, and the maximum instantaneous vertical acceleration was 12m/s2. In addition, synchronous flap, large angle of attack and large flapping angle were favorable for lift.
Unsteady flow field interaction of coaxial rotor
2019, 34(7): 1459-1470. doi: 10.13224/j.cnki.jasp.2019.07.007
Abstract PDF(4974KB) icon891 icon731
Abstract:
The aerodynamic interaction of coaxial rotor during the meeting of blades is most intense in the periodic interaction. In order to explore the characteristics of the flow field interference during the meeting of upper and lower blades of coaxial rotor, a numerical simulation method for aerodynamic interaction based on the unsteady Reynolds averaged Navier-Stokes (RANS) equations was established. Moving overset grid was applied to simulate the motion of the coaxial rotor. The motion of two airfoils with relative horizontal translations in the flow field was simulated to investigate the characteristics of flow field interference at a specific cross section of coaxial rotor. The variation of aerodynamic performance and the characteristics of flow field for the coaxial rotor and two airfoils system were analyzed, and the parameter effects of two airfoils system on the aerodynamic interaction were discussed. Results indicated that the lift increased first and then decreased dramatically during approaching of blades of the coaxial rotor and two airfoils system, and finally became stable. The lift fluctuation amplitude of upper and lower rotors was 30% and 22%, respectively. The aerodynamic interaction between the upper and lower airfoils decreased with the increase of the separation distance, and the lift fluctuation of the upper and lower airfoils changed differently to the variation of separation distance. The amplitude of the lift fluctuation of the two airfoils rose with the relative thickness of airfoil, and the secondary fluctuation occurred. The amplitude and range of the lift fluctuation of the airfoils with smaller relative inflow velocity were larger when there was extra inflow.
Field synergy and entropy generation of convective heat transfer in microchannels with complex structure
2019, 34(7): 1471-1482. doi: 10.13224/j.cnki.jasp.2019.07.008
Abstract PDF(8189KB) icon1161 icon720
Abstract:
The effects of fin width, cavity width and Reynolds number (Re) on the convective heat transfer characteristics in the microchannels with cavities and fins were studied by field synergy and entropy generation principle. The heat transfer enhancement mechanism of micro structures was analyzed, and the comprehensive performance of the microchannel was assessed. The results showed that the increase of fin width and cavity width was able to reduce the heat transfer synergy angle, improve the synergy relationship between flow field and temperature field obviously and enhance the convective heat transfer. However, the local vortexes caused the decrease of flow synergy angle and increased the pressure drop of the microchannels; the increase of fin width was helpful to improve the energy utilization efficiency and enhance heat transfer associated with the increase of the entropy generation rate of fluid friction; increasing the cavity width properly was able to reduce the entropy generation rate of heat transfer, but the extremely large cavity width led to the increase of the heat transfer irreversibility and flow friction; considering the pump power, relative fin width and relative cavity width, new correlation of the thermal resistance was proposed; the microchannel heat sink with relative fin width 0.2 and relative cavity width 2 yielded the lowest thermal resistance and the best overall performance.
Modeling and thermodynamic performance analysis of thermal control system based on terminal circuit with TEC in space station
2019, 34(7): 1483-1492. doi: 10.13224/j.cnki.jasp.2019.07.009
Abstract PDF(2463KB) icon956 icon668
Abstract:
A thermal control system was designed based on terminal single phase fluid circuit with thermoelectric cooler (TEC), which was capable of solving the problem of overranging temperature of science loads, when the intermediate circuit temperature fluctuated greatly. A thermal control module based on TEC was presented for improving the cooling ability of the fluid circuit, to recover its cooling capacity. Extra cooling power was provided to lower the working fluid temperature of inlet flows through the cold plate, providing a relative lower temperature region around science load, when the fluid circuit power was insufficient because of the high temperature of the intermediate circuit flow. Moreover, a mathematical model and a numerical simulation model were established separately for the proposed thermal control system. Furthermore, the thermal control performance of the TEC module was verified in the cases of four thermal disturbances, such as heat load disturbance of science load, temperature disturbance of the intermediate circuit, mass flow rate disturbance of the terminal circuit, and heat disturbance of parallel branches. Simulation work shows that: the designed thermal control system keeps the temperature fluctuation margin of science load below 1K for strict cooling requirement, in several conditions such as: science load heating power increase by 30%, intermediate circuit temperature increase by 5K, terminal circuit mass flow rate decrease to 0.0015kg/s.
Experiment on infrared radiation characteristics of S shaped 2D nozzle
2019, 34(7): 1493-1500. doi: 10.13224/j.cnki.jasp.2019.07.010
Abstract PDF(2317KB) icon1064 icon703
Abstract:
A experiment model of S shaped 2D nozzle was designed based on axisymmetric exhaust system and studied on the infrared characteristic experiment rig. Flow parameters and the far field infrared radiation characteristics of the experiment model were measured under experimental conditions, and compared with the baseline axisymmetric nozzle experiment data. Results showed that, infrared radiation power of S shaped 2D nozzle was mainly in the upper detection plane, and the maximum radiation was in 15 degrees, due to the largest radiation contribution of solid walls in this direction angle. Infrared radiation power in lower and side detection plane was much less than that in upper detection plane because of the shielding effect of S shaped 2D nozzle. Compared with the axisymmetric nozzle, S shaped 2D nozzle had significant effect of infrared suppression, the infrared signature of the S shaped 2D nozzle reduced by 81% in the rear direction, and the maximum radiation intensity reduced by 47%, while the gas radiation reduced by 50%.
Influencing factors of the lean blowout of alternative fuels
2019, 34(7): 1501-1509. doi: 10.13224/j.cnki.jasp.2019.07.011
Abstract PDF(2272KB) icon955 icon698
Abstract:
In order to research the effect of fuel and inlet air pressure on the lean blowout (LBO), the experiment using standard kerosene fuel (RP-3), diesel and high boiling Fischer-Tropsch oil (FT) in three different combustor inlet pressures was conducted. The results showed that the influence of inlet pressure (19.17%) on LBO was greater than the effect of fuel (6.26%). The main factors of LBO include inlet air pressure, flame volume, combustion temperature, the fuel atomization diameter and the LHV and density of fuel. The flame volume and fuel atomization diameter were mainly affected by the fuel property while the combustion temperature was mainly influenced by the inlet pressure. The change of LBO affected by inlet pressure was attenuated by the affected flame volume and combustion temperature. Fuels with low carbon number and high alkane content may increase the effect of flame volume on LBO, and result in a better LBO in low inlet pressure.
Control method of thermoacoustic instability with sound lining with heating bias flow
2019, 34(7): 1510-1518. doi: 10.13224/j.cnki.jasp.2019.07.012
Abstract PDF(4029KB) icon948 icon648
Abstract:
The passive control method of thermoacoustic instability was researched through experimental methods. A horizontally placed Rijke tube, a thermoacoustic instability experimental device, was set up; and an electrically heated hot gauze was used as a heat source in Rijke tube. It was found in the experiment that the temperature of the air before and behind the hot gauze had a crucial effect on the amplitude of the thermoacoustic oscillations. Moreover, sound linings with back cavity and perforated plate structures were used to control thermoacoustic instability. The bias air can be drained into the back chamber, and the flow and temperature of the bias air can be adjusted. It was also found that the biased air in the back cavity can enhance the inhibitory effect on the instability, and the effect became better with the increase of the bias flow rate; in addition, increasing the temperature of the biased air can not significantly increase the control effect of sound linings on thermoacoustic instability.
Theory and experiment on the leakage characteristics of brush seals based on fluid-structure interaction
2019, 34(7): 1519-1529. doi: 10.13224/j.cnki.jasp.2019.07.013
Abstract PDF(4002KB) icon908 icon491
Abstract:
The leakage flow characteristics and the influencing factors of the brush seals were studied both theoretically and experimentally. A fluid-structure interaction theoretical model was established to analyze the leakage flow characteristics of brush seals. Experiments were presented to test the leakage quantity of brush seals with different pressure ratios and rotational speeds and compared with the theory solution model to verify. Based on the accuracy of the theory computational model, the effect law of the brush seal leakage quantity in different structural parameters was analyzed, and the formula for predicting the leakage quantity of brush seals was constructed. The results showed that, the leakage quantity of the brush seals increased approximately linearly with the increase of the pressure ratio when the radial clearance was fixed. The rotational speed had little influence on the leakage quantity of the brush seals. Bristles diameter and the clearance of the bristles affected the leakage quantity through affecting the effective flow area of the gas in the brush area, the smaller effective flow area led to the smaller leakage quantity. With the increase of the bristles diameter and the axial rows of the bristles, the leakage quantity of the brush seals decreased gradually. The leakage quantity increased with the increase of the bristles clearances. As the backing plate fence height and the axial clearance between the last row bristles and the backing plate fence increased, the leakage quantity firstly increased rapidly, and finally became stable. The constructed theoretical formula of leakage quantity can accurately predict the leakage quantity of brush seals and provide a theoretical basis for the optimization design of brush seals structure.
Summary on the development of auto-balancing technology of aero-engine
2019, 34(7): 1530-1541. doi: 10.13224/j.cnki.jasp.2019.07.014
Abstract PDF(4380KB) icon1353 icon553
Abstract:
For the problem of unbalanced vibration of aero-engine during operation, the auto-balancing technology was used to provide online compensation quality and reduce the vibration effectively. At present, the auto-balancing technology on aero-engines had problems and challenges in three aspects: structure, control algorithm and engineering application. The electromagnetically driven auto-balancer based on the influence coefficient algorithm was applied in the US C-130H transport aircraft. The test results from the four aero-engines in the flight showed that with auto-balancer, the propeller was balanced to below 1.27mm/s, 94% reduction. Meanwhile 75% reduction in propeller-induced vibration was measured at the compressor.Research indicates that aero-engine auto-balancing technology can achieve continuous online dynamic balancing, which will greatly reduce the vibration fatigue damage of the aero-engine and its components，reduce the maintenance cost and prolong the maintenance cycle. Auto-balancing technology requires a gradual process for the aero-engine applications. In order to achieve this goal, it is necessary to carry out the research work of structure design, control strategy, system integration and so on. At the same time, it is necessary to take the auto-balancing structure into account at the very beginning of the aero-engines design.
Experiments on force coefficients identification of elastic ring squeeze film damper
2019, 34(7): 1542-1550. doi: 10.13224/j.cnki.jasp.2019.07.015
Abstract PDF(2562KB) icon923 icon514
Abstract:
The test rig for the force coefficients of elastic ring squeeze film damper (ERSFD) was set up. ERSFD was excited by two shakers located orthogonally providing harmonic force. The results were obtained under centered circular orbit and utilized for force coefficients identification according to the orbit kinetic of the damper. The results revealed that the damping and stiffness of the film and elastic ring were both significant, the stiffness and damping of the film dropped dramatically with the increase of the height of the pedestal while the force coefficients remained almost constant. The thickness of the ring had little influence on the force coefficients of the film. The damping of the film first turned lager with the increase of feeding pressure and remained constant under high pressure while the stiffness of the film first increased and then decreased a little bit with the increase of feeding pressure. Uncertainty analysis indicates that the uncertainty of the film force coefficients Cxx,Cyy,Kxx,Kyy is 12.2%,11.5%,18.2%,12.7%, respectively while the uncertainty of the force coefficients of the elastic ring Cxx,Cyy,Kxx,Kyy is 30.7%,33.1%,17.0%,12.8%.
Characteristic of magnetic field for LIPS-200 ion thruster based on equivalent magnetic circuit
2019, 34(7): 1551-1557. doi: 10.13224/j.cnki.jasp.2019.07.016
Abstract PDF(1703KB) icon692 icon451
Abstract:
The magnetic circuit structure and magnetic field configuration of LIPS-200 ion thruster capable of working steadily after performance enhancement were obtained to minimize the anode shock voltage and discharge loss by using orthogonal experimental method. Combined with the equivalent magnetic circuit method and the finite element discretization form, the magnetic field model of LIPS-200 ion thruster was established.Then the permanent magnet alternative proposals for electromagnet in specific spatial arrangements were studied.And the equivalence of alternative proposals and the effectiveness and feasibility of the simulation method were validated by comparing with the magnetic induction intensity and working performance.Results showed that the relative error of magnetic induction intensity in key points was less than 5% after magnetic field conversion, and the sensitive parameters changed in line with expectations, the design goal of magnetic circuit structure optimization and performance improvement for LIPS-200 ion thruster was well realized.
Numerical simulation of spoiler ring’s effect on combustion and flow of powdered rocket motor
2019, 34(7): 1558-1567. doi: 10.13224/j.cnki.jasp.2019.07.017
Abstract PDF(3191KB) icon787 icon494
Abstract:
In order to improve the combustion efficiency of powdered rocket motor, the effects of spoiler ring’s existence, flowing diameter and position on the combustion and flow characteristics in the chamber were studied by numerical simulation. Result showed that the spoiler ring can increase mixing and heat transfer intensity between particle and gas species, and improve the evaporation and combustion of Al particle, which resulting in the increase of combustion efficiency of the powdered rocket motor. When the ratio of spoiler ring flowing diameter to the combustion chamber diameter was between 0.538 and 0.846, the combustion efficiency increased with the reducing flowing diameter of spoiler ring. When the ratio of spoiler ring’s distance from the chamber’s head position to the chamber’s total length was between 0.3 and 0.8, the combustion efficiency increased as the spoiler ring moved toward the chamber’s head. The spoiler ring with small flowing diameter and distance from the head between 30% and 40% of the chamber length could be better for the spoiler ring design of powdered rocket motor.
Experiment of Hall thruster plume effects on different ground experimental conditions
SU Yang, CAI Guobiao, HE Bijiao, SHANG Shengfei, LIU Peng, LIANG Wei
2019, 34(7): 1568-1576. doi: 10.13224/j.cnki.jasp.2019.07.018
Abstract PDF(2663KB) icon736 icon518
Abstract:
Reducing the influence of back-sputtering effect can effectively improve the accuracy of the measurement of the Hall thruster plume effect. Quartz crystal microbalance (QCM) was used to measure the deposition and sputtering distribution of HET-40 thruster plume on two different experimental conditions: case 1, using liquid nitrogen heat sinks and case 2, without using liquid nitrogen heat sinks. Meanwhile, X-ray photoelectron spectroscopy (XPS) was used to analyze the composition of the QCM surface after two experiments. The results of the two experiments showed that the sputtering rate under the condition of case 1 was slightly higher than case 2. Especially within the range of 90° to 110° relative to the thruster axis, case 1 experiment result showed sputtering effect, while case 2 experiment showed deposition effect. Through analysis of the experimental results, it can be found that using liquid nitrogen heat sink to reduce the temperature of the inner wall surface of vacuum chamber can effectively adsorb the particles sputtered by the plume and reduce the concentration of back-sputtering particles, leading to the above phenomenon.
Research on micro centrifugal compressor heat transfer effect estimation model
2019, 34(7): 1577-1585. doi: 10.13224/j.cnki.jasp.2019.07.019
Abstract PDF(2280KB) icon654 icon512
Abstract:
For non-adiabatic boundary condition of micro-centrifugal-impeller,after heat transfer process was analyzed in the micro-centrifugal-impeller,the disk and blade were treated as annular and straight fins respectively,then a new math model was built to describe the phenomenon in it;the model was modified for actual situation based on differential equations of heat conduction and thermal equation of fins;after 3-D numerical simulation validation, the model showed that,at least,4% accuracy improvement was achieved compared with zero-dimension-network-model;compared with old design philosophy, which only assessed heat addition effect by numerical simulation,the new one,with the help of our model,can improve 30% efficiency and 11% total pressure radio, and cut the design cycle down to 14.3%.
Numerical simulation of tip leakage vortex trajectory and stall prediction for centrifugal impeller
2019, 34(7): 1586-1597. doi: 10.13224/j.cnki.jasp.2019.07.020
Abstract PDF(5466KB) icon772 icon522
Abstract:
The Krain impeller with total pressure ratio of 4.1 was taken as the research object. The effects of mass flow rate, rotational speed and tip clearance on the tip leakage vortex (TLV) trajectoryand the interface between the tip leakage flow and the main flow (ITLMF) were numerically investigated. Results showed that the decrease of mass flow rate, the increase of the rotational speed and the decrease of the tip clearance made the TLV trajectory move away from the suction surface and the ITLMF move upstream. By simplifying the main flow and tip leakage flow interaction as the free-stream and counter-flow wall jet interaction, a subsonic centrifugal impeller stall-onset prediction model was developed. In this model, the momentum balance analysis was applied to identify the position of main flow/tip leakage flow interface. The tip leakage flow velocity was modeled. Zhao’s model was used to predict the TLV trajectory. The relationship between the virtual origin of tip clearance jet and the tip clearance was determined. Results showed that the stall point predicted by current model agreed well with CFD, with the root mean square error was below 2.42%.
Load sharing performance of two branches split-torque transmission system with cylindrical gears
2019, 34(7): 1598-1606. doi: 10.13224/j.cnki.jasp.2019.07.021
Abstract PDF(3557KB) icon904 icon478
Abstract:
The principle of non-uniform load distribution in torque-split transmission system of cylindrical gears was analyzed. By establishing the dynamic model of this torque-split transmission system the effects of structure parameters between two sides on the load sharing performance were researched. The test of the split-torque transmission with elastic shaft was carried out, and correctness of the analysis method above was proved.Results showed that the load condition, torsional stiffness, support stiffness and mesh phase differences between two sides had an effect on the load sharing performance of the system, among which smaller torsional stiffness, larger support stiffness and suitable meshing phase difference were beneficial for load sharing.
Oil film dynamic characteristics during oil droplet’s normal impact onto metallic surface
2019, 34(7): 1607-1614. doi: 10.13224/j.cnki.jasp.2019.07.022
Abstract PDF(2229KB) icon797 icon454
Abstract:
To explore the spread behavior of the deposited oil film formed by droplet-housing impact in aero-engine bearing chamber, a three-dimensional numerical model was presented using volume of fluid (VOF) method to predict the oil/air two-phase flow during the normal impact of oil droplet onto a smooth stainless steel surface. The static contact angle between droplet and surface was measured by experiment. The dynamic spreading and recoiling process of deposited oil film during the normal impact was analyzed, and the effects of droplet diameter and impact velocity on dynamic characteristics such as spread diameter, nominal film thickness, spread velocity and impact force were discussed. The calculation results showed preliminarily that at maximum spread diameter the film appeared to be a flat disk with slightly concave center, and in the recoiling period, it appeared to be a disk with thin rim and thick center. The recoiling velocity of film and the impact force in the recoiling period reduced nearly to 0 and remained approximately constant. The oil film spread diameter, nominal film thickness, spread velocity and impact force increased with the increasing droplet diameter. The spreading diameter and impact force increased with the increasing impact velocity of droplet, whereas the nominal film thickness decreased. Finally, the reliability and correctness of the model presented was validated by comparing with relevant experimental data.
Model reference adaptive control with smooth switching scheme for piecewise linear systems and its application in turbofan engine control
SUN Penghui, WANG Xi, YANG Bei, ZENG Detang
2019, 34(7): 1615-1623. doi: 10.13224/j.cnki.jasp.2019.07.023
Abstract PDF(1350KB) icon764 icon491
Abstract:
A model reference adaptive control (MRAC) with smooth switching scheme was proposed for piecewise linear systems, and the method was utilized in turbofan engine control to avoid the discontinuity of control input. In this scheme, each sub-region of the operating envelope had its own MRAC controller, and smooth indicator function based smooth switching scheme was introduced to switch multiple controllers smoothly at the boundary of adjacent sub-regions. The Lyapunov stability analysis indicated that the proposed smooth switching scheme can guarantee the convergence of the closed-loop system during the controllers switching. The tracking error system was converted into a switched system to analyze the global stability of the closed-loop system. The advantage of the method was that the chattering of system output and instability caused by asynchronous switching can be eliminated. The simulation illustrates the effectiveness of the proposed control scheme in comparison with the existing MRAC controller with gain scheduling for turbofan engine.
Fuel metering unit performance degradation detection and remaining useful life estimation method based on RF-SVR
2019, 34(7): 1624-1632. doi: 10.13224/j.cnki.jasp.2019.07.024
Abstract PDF(3264KB) icon967 icon456
Abstract:
In order to realize the safety state monitoring and health management of the aero-engine fuel system, the fuel system performance degradation detection and remaining useful life estimation was researched.Taking the fuel system fuel metering device as an example, the main performance degradation mode was analyzed.The health indicators (HIs) selection scheme based on current-speed data was designed.Considering the uncertainty of environment and model parameters, the model uncertainty simulation was carried out.The component performance degradation was detected based on the Mahalanobis distance between the healthy data and the performance degradative data.A remaining useful life (RUL) estimation method based on random forest-support vector regression (RF-SVR) was proposed.The component RUL estimation was realized by SVR model optimized by RF feature selection.Finally, the method was validated based on the simulation data of a certain type of civil turbofan engine mechanical hydraulic model.The results show that the performance of the method has a false alarm rate and a false negative rate of less than 2%, and the RUL estimation error is less than 3%. This provides a reference for predictive maintenance of aero engine fuel systems.

Copyright © 《Journal of Aerospace Power》 All Rights Reserved.

Supported by: Beijing Renhe Information Technology Co., Ltd.