2020 Vol. 35, No. 4

Cover
  Previous Issue | Next Issue 
Display Method:
Development of comprehensive surrogate fuel model for RP-3 aviation kerosene
YU Jin, YU Binbin, YU Jiajia
2020, 35(4): 673-681. doi: 10.13224/j.cnki.jasp.2020.04.001
Abstract PDF(4610KB) icon1192 icon976
Abstract:
A comprehensive surrogate of RP-3 fuel was developed by the methodology of directly matching the molecular structure and functional groups. n-dodecane, 2,5-dimethylhexane, 1,3,5-trimethylbenzene and decalin were selected as the surrogate components. The physical surrogate model can well predict the main physical properties of RP-3 aviation kerosene from subcritical to supercritical pressures. Besides, the performances of chemical surrogate fuel were also validated against experimental data. Present surrogate fuel model was in good agreement with the experimental ignition delay times both at low and high temperature regions. In addition, present surrogate fuel could also well reflect the ignition delay time at low pressures (0.1-0.01 MPa). Results showed that present surrogate fuels could emulate both the physical and kinetic properties well. The success of present surrogate fuel lays a foundation for understanding of the coupling mechanism of fuel regenerative cooling process and combustion process, and achieving co-simulation of aero-engine regenerative cooling system and propulsion power system in scramjet.
Flow and temperature characteristics of wall film in bearing chamber located in hot region of aero-engine
FANG Long, CHEN Guoding
2020, 35(4): 682-692. doi: 10.13224/j.cnki.jasp.2020.04.002
Abstract PDF(4106KB) icon671 icon839
Abstract:
The wall film in bearing chamber is divided into flow control units which are related to each other. Based on the motion state of droplet and wall film, the mass, momentum and heat transfer were analyzed in each flow control unit. And then the unsteady mass, momentum and heat equations of each flow control unit were established and solved by house coding with help of Matlab. A theoretical method to analyze flow and heat transfer characteristics of wall film in bearing chamber was established. Results showed that the film thickness was within the range of 0-1 mm, the film velocity was within range of 0-5 m/s, and the film temperature was within the range of 100-140 ℃. At the same condition, the highest film temperature increased by 13% compared with the lowest film temperature. The film velocity and temperature increased with the rotor speed, while the film thickness decreased with the rotor speed. The film thickness and velocity increased with inlet flow rate, while the film temperature decreased with inlet flow rate. The research work is not only a theoretical foundation but also a data support for improvements of the efficiency of aero engines.
Effect of showerhead injection of leading edge on cooling characteristics for turbine vane pressure side
YAO Chunyi, ZHU Huiren, LI Xinlei
2020, 35(4): 693-703. doi: 10.13224/j.cnki.jasp.2020.04.003
Abstract PDF(3437KB) icon791 icon782
Abstract:
In order to evaluate the effect of showerhead injection of leading edge on the cooling characteristics of turbine vane pressure side with multirow film holes,an experiment was carried out in a high subsonic wind tunnel to obtain the film cooling effectiveness and heat transfer coefficient on the vane surface with/without showerhead injection of leading edge. The inlet Reynolds number of the cascade (based on the true chord length of the vane) varied from 2.0×105 to 4.0×105,the exit isentropic Mach number was 0.95. Six rows of cylinder holes were arranged on the leading edge and the range of mass flow ratio was 2.46%-4.57%,the pressure side was assigned with 6 rows of cylinder holes and the mass flow ratio varied from 2.00% to 3.71%. The experimental results showed that without showerhead injection of leading edge,the film cooling effectiveness on the front half of the pressure side was less affected by mass flow ratio,however,that on the rear half increased with the increase of mass flow ratio. The showerhead injection of leading edge raised the film cooling effectiveness of pressure side with multi-row holes by 20%~70%,and made the distribution of film cooling effectiveness more uniform along the flow direction. No matter whether the showerhead injection of leading edge existed,the heat transfer coefficient ratio on the pressure side increased with the mass flow ratio increasing. Viewed from streamwise, the showerhead injection of leading edge increased the heat transfer coefficient ratio in the region of leading and trailing edges on the pressure side,however,exerted a little effect in the middle region of pressure side.
Analysis on interaction between centrifugal compressor and ,pulse detonation combustor
WANG Lingyi, ZHENG Longxi, JIA Shengxi
2020, 35(4): 704-710. doi: 10.13224/j.cnki.jasp.2020.04.004
Abstract PDF(1973KB) icon792 icon801
Abstract:
In order to realize the stable operation of centrifugal compressor and pulse detonation combustor (PDC) combined system, it is necessary to investigate the interaction between them. A pulse detonation turbine engine principle test system was established, in which reversal-pressure wave was continuously generated from the high-pressure combustor to the centrifugal compressor. A simulation model was constructed to analyze the reversal-pressure wave. The simulation model was verified by the pulse detonation turbine engine principle test system. Based on that, the operating characteristics of the centrifugal compressor were calculated and analyzed. Results showed the reversal-pressure wave caused an instantaneous gas back flow in the compressor, leading to pressure oscillations at the compressor exit, and the compressor outlet was under unstable condition for a long time consequently. When the compressor worked together with the PDC, the compressor operating line got closer to the surge boundary, with the compressor efficiency dropping below 0.39, while the efficiency was over 0.81 under the condition of the compressor working alone.
Cooling performance of electric actuating mechanism on ,aircrafts based on heat pipes
NIU Wenjing, LIAN Wenlei, LIN Lingjiao
2020, 35(4): 711-721. doi: 10.13224/j.cnki.jasp.2020.04.005
Abstract PDF(2202KB) icon711 icon747
Abstract:
Research on cooling technology was carried out considering that the electric actuating mechanism on aircrafts would produce much heat in working. The empirical formula was fitted based on the experiment data. A cooling method for an aircraft electric actuating mechanism was proposed based on the heat pipe-fuel cooling system, and a numerical model describing the flow and heat transfer of the cooling system was established and validated feasibly. The influence of some factors on the cooling performance of the cooling system was studied by numerical simulation. Results showed that the error between numerical simulation and experiment data was within 14.6%. The heat pipe-fuel cooling system could reduce the temperature rise of the electromechanical actuator by 12.61% compared with natural convection. For the driver, the fan cooling method was superior to the heat pipe-fuel cooling method. When the effective thermal conductivity of heat pipes reached 7 000 W/(m·K), the heat pipe-fuel cooling system could achieve a good cooling performance, and continuously increasing the effective thermal conductivity of heat pipes couldnt effectively improve the cooling performance of the cooling system. This research provides a reference for designing an efficient cooling system for electric actuating mechanisms.
Simulation and analysis of temperature of aviation engine fuel system based on Flowmaster software
KANG Sizhao, XI Xiuzhi, LI Bo
2020, 35(4): 722-731. doi: 10.13224/j.cnki.jasp.2020.04.006
Abstract PDF(1927KB) icon1072 icon978
Abstract:
In order to study temperature rise of the fuel system under constant and variable speeds, based on the Flowmaster software and the Daqing RP-3, the aeroengine fuel system simulation model was established, the temperature simulation of engine was carried out, and experiment and simulation results were compared to verify the accuracy. Results showed that: the performance curve of components led to the calculation error. The maximum temperature was in the front of the main combustion chamber, up to 145 ℃ in some working cases, which could affect the safety of aeroengine, so the temperature of fuel must be controlled. The measure of taking fuel returning to the aircraft was useful. The temperature could not immediately achieve steady when the mass flow was instantly changed. With the rise of main combustion chamber flow, the fuel temperature declined when the returning mass flow was 0 kg/s and the centrifugal pump efficiency under different working conditions was equal. The temperature also achieved steady when mass flow was at a steady state. The calculation method of fuel temperature and fuel pump is proposed for the one-dimensional simulation further research of aviation engine system level.
Contact stress calculation and static strength analysis of zig-zag shrouded blade
SHEN Xiuli, GAO Pengxin, DONG Shaojing
2020, 35(4): 732-743. doi: 10.13224/j.cnki.jasp.2020.04.007
Abstract PDF(3881KB) icon923 icon957
Abstract:
The “displacement extraction method” was proposed to calculate the contact stress of the zig-zag shroud contact surface according to the convergent displacement of contact surface. The contact stress can be obtained accurately with the surface mesh size of 0.202 mm, and the maximum value was 453.48 MPa. The stress level and torsion degree of the shrouded blade under different conditions were analyzed. The stress level increased with the increase of initial installation tightness, and the blade torsion angle increased with the increase of centrifugal load. According to the relationship between the load on the contact surface and the torsion angle, the compressional rigidity and torsional rigidity of blade were obtained. The compressional rigidity was much higher than torsional rigidity, so the stress of blade body restricted the stress of contact surface. It was proposed that the average contact stress of shroud should not be greater than 40% nominal yield limit and the equivalent stress of blade body should not be greater than 75% nominal yield limit as a criterion for the static load strength analysis, and the shrouded blade model of 0.127 mm initial installation tightness can meet this strength requirement.
Low pressure rotor dynamic behavior simulation of Trent 900 engine after fan blade out
XU Xue, LI Hongxin, FENG Guoquan
2020, 35(4): 744-756. doi: 10.13224/j.cnki.jasp.2020.04.008
Abstract PDF(5728KB) icon969 icon940
Abstract:
To enhance the understanding of the dynamic behavior of the low pressure rotor and structural fuse in three-shaft high bypass ratio turbofan engine during a fan blade off （FBO） event, the Trent 900 was selected as the research sample.The structural characteristics which may cause important influence was analyzed, and the research on the influence of the dynamic coupling between the rotors, the simulation of structural fuse, and fan blade simplification was conducted. The three-dimensional rotor/stator simplified model for the FBO simulation was built. The simulation of 360 degree rotating of the fan rotor was performed and compared with the video of Trent 900 engine FBO rig test. The results revealed that: (1) the simulation results matched the video of FBO rig test in the time sequence of macro phenomena. (2) The breakage of main structural fuse can reduce the output vibration load by about 45%. (3) Secondary structural fuse can form a new front support for the low pressure rotor and protect the fan shaft.
Interval analysis of rotor dynamic characteristics based on ,Chebyshev polynomials expansion
WANG Cun
2020, 35(4): 757-765. doi: 10.13224/j.cnki.jasp.2020.04.009
Abstract PDF(3401KB) icon853 icon768
Abstract:
Interval analysis method on critical speeds and steady response was established to acquire accurate results with the bounds of parameters based on Chebyshev expansion. Critical speeds were expressed as Chebyshev polynomials expansion to calculate the interval results when support stiffness, density or elastic modulus were interval parameters. The comparison between Chebyshev method, perturbation method and combination method illustrated that the maximum error was less than 0.07%. The Chebyshev method on steady response was proposed while modal superposition was applied to correct the bounds. The defect of the original method arising from the large uncertainty of critical speeds was fixed. The comparison with Monte Carlo method illustrated that the maximum error of response upper bound was less than 8%. The differential equations of the original system do not need to be modified in Chebyshev method. The results illustrate that combined with commercial finite element software, the method has good applicability and advantages of high accuracy.
Dynamic similarity design method for aero-engine, low-pressure rotor system
MIAO Hui, ZANG Chaoping
2020, 35(4): 766-776. doi: 10.13224/j.cnki.jasp.2020.04.010
Abstract PDF(1503KB) icon1020 icon857
Abstract:
Dynamic similarity design of an aero-engine rotor system was presented to study its dynamic characteristics using the design model instead of the real structural system. Firstly, the similar equivalent model of the rotor full scale prototype, based on the dynamic equivalence principle of the rotor system, was established by dynamic optimization. Then, the dynamic similarity criterions of the rotor system were derived by equation analysis combined with dimensional analysis, and the reduced-scale model of the similar equivalent model was constructed according to the similarity relations of the rotor design parameters. Finally, using the model updating method, the dynamic similarity model of high prediction accuracy was obtained by updating the similarity error of the three-dimensional design model. Taking the dynamic similarity design for the low-pressure rotor of the dual-rotor aero-engine as an example, the effectiveness of the dynamic similarity design method was validated by finite element simulation. Results showed that the dynamic similarity model obtained by this method could effectively predict the critical speeds and unbalance responses of the full scale prototype within the design operation speed. The similarity errors for the first three critical speeds were 3.09%, 1.75% and 0.31%, respectively, and the corresponding modal assurance criterion values were all above 0.93. This method has great value for practical application, and provides a reference for dynamic similarity design of the whole aero-engine.
Vibration reduction analysis and design of friction ring damper in blisk
WEN Wei, QI Wenkai
2020, 35(4): 777-782. doi: 10.13224/j.cnki.jasp.2020.04.011
Abstract PDF(1843KB) icon856 icon853
Abstract:
A kind of annular friction damper adapted to the blisk was designed to control the blade-disk coupling vibration. The first-order harmonic balance method and the macro-micro slip model were expanded to the cyclic symmetry model to simulate the vibration reduction of blisk with ring damper. Results of simulating realistic blisk showed that the annular friction damper had obvious effect on the 0-diameter 4th-order dangerous modal vibration of the blisk, and damping effect was related to the quality of the damper and the installation position. The amplitude was reduce by 75% at the selected location. The optimal damper design, including mass and installation design, was established when its structure was determined.
Aeroelastic analysis of curved panels subjected to impinging, oblique shock based on fluid-structure coupling algorithm
LI Yingkun, CHEN Xiong, XU Jinsheng
2020, 35(4): 783-792. doi: 10.13224/j.cnki.jasp.2020.04.012
Abstract PDF(5154KB) icon920 icon874
Abstract:
A developed fluid-structure coupling solver was used to study the aeroelastic response behaviors of curved panels subjected to an impinging oblique shock. The geometric nonlinear large deformation equations of curved panel were solved by finite difference method,while the fluid governing equations were solved by means of finite-volume scheme. The conventional serial staggered algorithm was adopted for the fluid-structure interaction. Numerical results demonstrated that as the dynamic pressure was smaller than the critical flutter dynamic pressure, the curved panel exhibited a static equilibrium state,and the asymmetry of the panel deformation was more pronounced with the increasing dynamic pressure. When the dynamic pressure was greater than the critical flutter dynamic pressure, the panel vibration displacement increased firstly and then decreased,and finally reached a stable flutter state. This limit cycle flutter was asymmetric about the initial point. Meanwhile,as the dynamic pressure increased,the maximum positive and negative displacements of panel flutter,and flutter amplitude increased gradually,but the frequency of flutter reduced gradually. The panel flutter behaviors did not monotonously increase or decrease with altering curvature of panel. The smaller initial curvature can reduce the critical dynamic pressure of panel flutter. However,as the panel curvature further increased,the quasi-periodic irregular oscillation was excited due to the enhancement of the aerodynamic nonlinearity,and the critical dynamic pressure rose rapidly.
Fluid-thermal coupling on drag and heat reduction induced by spike and jet in hypersonic flow field
LI Yingkun, ZHU Liang, CHEN Xiong
2020, 35(4): 793-804. doi: 10.13224/j.cnki.jasp.2020.04.013
Abstract PDF(6035KB) icon691 icon442
Abstract:
To alleviate the serious temperature and pressure environment of the forebody of hypersonic vehicle, the combined drag and heat reduction scheme of spike and opposing jet was studied in depth. The Reynolds averaged Navier-Stokes equations were solved based on the finite volume method, and the shear stress transport k -ω turbulence model was used. The conjugate heat transfer method was used to solve the heat conduction equation of solid domain. Results showed that the application of the opposing jet significantly improved the performance of drag and heat reduction. The drag coefficient of blunt forebody decreased significantly with the increase of spike length. With the length-to-diameter ratio of the spike increasing from 0.5 to 2.0, the drag coefficient decreased to about 21%. The peak value of heat flux was almost unaffected. Increasing the total opposing jet pressure ratio can significantly reduce pressure on the blunt forebody surface. Considering the additional drag induced by the opposing jet, as a result, the drag reduction effect became worse. With the total opposing jet pressure ratio increasing from 0.4 to 0.8, the peak value of heat flux on the blunt forebody surface decreased by 62.5%. The result of conjugate heat transfer analysis shows the temperature of blunt forebody increases significantly with time, and the peak value of heat flux decreases slowly with time.
Numerical study of bypass dual throat nozzlewith parallelogram cross-section
WANG Yangsheng, XU Jinglei, HUANG Shuai
2020, 35(4): 805-814. doi: 10.13224/j.cnki.jasp.2020.04.014
Abstract PDF(3987KB) icon484 icon473
Abstract:
A parallelogram section bypass dual throat nozzle(BDTN) was proposed by combining the heterotypic outlet design with the fluidic thrust-vectoring nozzle, and then a comparative study of the flow filed structure and aerodynamic performance with the basic rectangle BDTN was carried out. The results of three-dimensional numerical simulation showed that the parallelogram configuration had the same aerodynamic parameter variation law as the rectangular configuration. Compared with the basic rectangle BDTN, the flow field of the parallelogram configuration changed due to the inclination of sidewalls, resulting in the decrease of the pitch vector angle and thrust coefficient in vectoring mode, while it had little effect on the thrust coefficient and flow coefficient in non-vector mode. The sidewall inclination angle was an important factor for the change of nozzle performance. Under the same nozzle pressure ratio, the lager inclination angle indicated the smaller vector angle. When inclination angles were not less than 60°, stable vector angles above 10° and maximum vector angles above 15° could be obtained. The parallelogram outlet had a better blending effect on the exhaust jet, which could greatly accelerate the attenuation of the centerline velocity of jet flow and contribute to improve the infrared stealth characteristics.
Flight performance of compound helicopter with twin propellers
HE Xiaoping, HAN Dong, YANG Kelong
2020, 35(4): 815-822. doi: 10.13224/j.cnki.jasp.2020.04.015
Abstract PDF(2539KB) icon749 icon478
Abstract:
A compound helicopter trim model with aerodynamic models of propeller and wing based on flight performance model of helicopter was constructed to investigate the effect of propeller and wing on the performance of compound helicopter with twin propellers. Taking the X-3 helicopter as an example, by adding different configurations of propeller and wing, the changes of each control strategy and attitude angle of the helicopter were given, and the effect of the variable proportions of propellers thrust and wings lift on the helicopter power required and helicopter lift-to-drag coefficient was analysed. Result showed that wing was useful to raise obviously peak lift-to-drag ratio, and also reduce collective and lateral cyclic pitch. The propeller could only improve the peak lift-to-drag ratio and reduce the helicopter power required until the flight speed is up to certain speed. The final compound helicopter configurated with propeller and wing could reliably achieve the less helicopter power required just by reasonably distributing thrust and lift, thereby improving the flight performance of the compound helicopter.
TLD simulation analysis with multiple faults for FADEC system ,under full repair policy
CAI Jing, HU We, CHEN Xi
2020, 35(4): 823-831. doi: 10.13224/j.cnki.jasp.2020.04.016
Abstract PDF(2508KB) icon689 icon417
Abstract:
The time-limited dispatch (TLD) analysis of the full authority digital control (FADEC) system is an important part of type certification of aero-engine. In order to make up for the existing TLD simulation analysis without taking into account the effect of the repair policy on TLD, and the short time as a variable, a TLD simulation analysis with multiple faults for FADEC system under full repair policy was proposed. The existing maintenance policies for FADEC system were compared, and according to the principle of opportunity maintenance or group maintenance, the full repair policy was presented. On the basis of the full repair policy, the analysis of simulation modeling was carried out, and the simulation process and method were designed and studied. The applicability and rationality of the proposed simulation analysis were verified through specific FADEC system. Results showed that the proposed simulation analysis was not affected by the number of parts in FADEC system, and in relation to single fault Markov model, the combined faults accounting for 8.84% of the total faults were analyzed, helping to improve the analysis accuracy.
Development of two-stroke kerosene engine controller ,based on MotoTron platform
KE Chang, HUANG Ying, ZHAO Zhenfeng
2020, 35(4): 832-843. doi: 10.13224/j.cnki.jasp.2020.04.017
Abstract PDF(4961KB) icon673 icon389
Abstract:
In order to realize the rapid development and strategy verification of the two-stroke kerosene aviation engine controller, in combination with the rich software and hardware resources of the MotoTron rapid control prototype development platform, based on the system design requirements of air-assisted direct injection and dual-spark plug ignition, an engine controller for controlling the four-cylinder engine using an eight-cylinder program was developed. The design method of control software based on the rapid control prototype software architecture was analyzed, and the development software, MotoHawk was used to configure the underlying program and build the engine control system. Then, in combination with Matlab/Simulink, the upper control strategy was established to completely develop the whole control system. The cold start bench tests were carried out, during which the speed of the engine was stable, the adjustment processes of the ignition advanced angle, the injection advance angle and the injection pulse width were in accordance with the control strategy. And the air-fuel ratio control was stable, with the error within 10%. The smooth transitions of the engine to normal conditions verified the cold start control strategy. And the fundamental functions of the controller had been realized, showing that the controller design could meet the requirements.
Neural network control of aircraft engine thrust ,degradation mitigation
YAN Zhaohong, QIU Xiaojie, HUANG Jinquan
2020, 35(4): 844-854. doi: 10.13224/j.cnki.jasp.2020.04.018
Abstract PDF(6209KB) icon627 icon417
Abstract:
A thrust degradation mitigation neural network control method of aircraft engine was proposed based on variable incremental linear programming (LP) optimization, due to performance deterioration of the gas path components. This method alleviated the engine thrust degeneration through control of high-pressure rotor speed and the engine pressure ratio by the inner loop, and correction of the engine command signal by the outer loop. The inner loop nonlinear autoregressive moving average (NARMA-L2) speed controller was obtained by neural network, and the outer command correction loop used the variable incremental LP optimization method to adjust the engine command signal. Simulations on a low-bypass-ratio-turbofan engine were performed. Results showed that under the 4 sets of simulation conditions, the designed control method can mitigate the thrust at least 46.5%, ensuring that the engine with performance deterioration was not overrun. The effectiveness of the method has been verified.
Working characteristic of silo ejector
QUAN Hui, XIE Jian, ZHANG Li
2020, 35(4): 855-866. doi: 10.13224/j.cnki.jasp.2020.04.019
Abstract PDF(1767KB) icon617 icon500
Abstract:
Considering the problem of silo design,the silo ejector model and ejector function were proposed. The properties of the static pressure matching function were studied. The critical conditions and ranges of the engine total pressure,the cross-sectional area of the silo and the outlet pressure of the mixing chamber were analyzed. This proved impossible to achieve the first limit of ejector,the zero points of the cross-sectional area of secondary flow and the constant velocity point; when the third limit of the ejector was not reached,the characteristic curve was subjected to the static pressure matching function; when the state of the ejector ranged from the third limit to the minimum limit,the characteristic curve was subjected to the optimal static pressure function; when the ejector worked in the over-congestion state,the characteristic curve was subjected to the minimum static pressure function. The results provide an important guidance for the design of silos and ejectors.
Flow field characteristics of angle-cut nozzle of solid rocket motor based on gas-solid two phase flow model
JIN Helong, JIANG Shuyuan, WANG Hao
2020, 35(4): 867-877. doi: 10.13224/j.cnki.jasp.2020.04.020
Abstract PDF(4377KB) icon825 icon475
Abstract:
In order to study the two phase flow characteristics of angle-cut nozzle of solid rocket motor (SRM), a gas-particle two-fluid model was adopted and combined with multi-region hybrid grid technology, and the interaction between gas phase and particles in the scarfed nozzle of solid rocket motor was researched. The effect of particle diameter and mass fraction on the gas-solid two phase flow characteristics of angle-cut nozzle was also investigated. Results showed that, the existence of solid particle phase had an important effect on the flow field structure of gas phase for angle-cut nozzle of SRM, bringing about a region with low velocity and high temperature of gas flow near the nozzle axis. At the same time, there was an area near the nozzle wall without particles. With the increase of particle diameter, the region gradually expanded without particles. The larger the particle diameter, the smaller the velocity; and the detention time was longer in the nozzle. The change of particle diameter and mass fraction also affected the flow field structure of rocket nozzle, with the increase of particle diameter, the gas phase Mach number first decreased and then increased along the nozzle axis, and the gas temperature first increased and then decreased, the trend of engine thrust was the same as that of Mach number, but both of which did not reach the extreme point simultaneously. The larger the particle mass fraction, the smaller the gas phase Mach number along the nozzle axis and the value of engine thrust; and the loss of two phase flow in the nozzle was larger.
Theoretical study on amplitude and frequency of compressor instability
SU Sanmai, CHEN Yaxin, SUN Zhanheng
2020, 35(4): 878-887. doi: 10.13224/j.cnki.jasp.2020.04.021
Abstract PDF(1743KB) icon720 icon492
Abstract:
In order to accurately describe the amplitude and frequency characteristics of compressor stall and surge during active stability control and health monitoring based on the compressor system model and the configuration of the stall cell, the stall behavior was analyzed by the modal wave theory from the perspective of dynamics. The surging phenomenon was analyzed by the strong nonlinear theory, based on the one-dimensional oscillation characteristics of surge. The theoretical description of the amplitude and frequency of the compressor instability was derived and given in the process of analysis. Simulation results show that, the amplitude and frequency of compressor stall and surge are the functions of the compressor characteristics, flow parameters and structural parameters. In the process of stall, the lower modes order means the easier occurrence and stabilization of amplitude. The higher modes order means the bigger rotation frequency of disturbance. In the process of surge, the smaller coefficient of compressor outlet throttling means the bigger value of the surge amplitude and oscillation frequency.
Numerical calculation and experiment on leakage characteristics of floating ring seal
HU Tingxun, ZHOU Kun, WANG Xiaoyan
2020, 35(4): 888-896. doi: 10.13224/j.cnki.jasp.2020.04.022
Abstract PDF(3600KB) icon724 icon455
Abstract:
To study the leakage characteristics of floating ring seal under different temperatures, pressures and rotational speeds, a numerical calculation model with CFD was built. The leakage characteristics of floating ring seal used in a specific aero engine were analyzed. For the change of leakage caused by rotational speed, a way was proposed to calculate the centrifugal expansion of rotor at different rotational speeds by finite element method, and to change the flow field size according to the centrifugal expansion. The change of state parameters of fluid caused by temperature and differential pressure was calculated by state equation, and the related parameters in constitutive equation were redefined. To verify the numerical calculation model, experimental study of leakage characteristics of floating ring seal under simulating real conditions was developed. Results showed that with the increase of differential pressures, computational results kept approaching the experimental results; when the differential pressures were greater than or equal to 0.1 MPa, the average deviations between computational results and experimental results were between 6.08%-9.43%.

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

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