2015 Vol. 30, No. 4

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Field synergy analysis of highly-intensified piston oscillating cooling
ZHU Hai-rong, ZHANG Wei-zheng, YUAN Yan-peng
2015, 30(4): 769-774. doi: 10.13224/j.cnki.jasp.2015.04.001
Abstract PDF(842KB) icon1627 icon902
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The oscillating cooling process of oil in the oil cooling gallery of highly-intensified piston was simulated numerically and the effective speed was defined. Through the post processing of data, the distributions of field parameters of oil cooling gallery wall in time and space were obtained, and the connection between wall surface heat transfer coefficient and the field synergy performance was analyzed. The results show that: the field synergy performance is good enough in the space range of 60 degree to 120 degree and 240 degree to 300 degree. Distribution rules of effective speed match well with the field synergy angle cosine value; uniform distributions of surface heat transfer coefficients of oil cooling gallery wall are in accord with the distributions of synergy angle cosine value and effective speed, illustrating that field synergy mechanism can explain the convective heat transfer rules of oil oscillating cooling.
Flow characteristics in a rotor-stator cavity with cooling air inlet at low radius
CAI Xu, LUO Xiang, ZHAO Xi, XU Guo-qiang, CAI Jun
2015, 30(4): 775-783. doi: 10.13224/j.cnki.jasp.2015.04.002
Abstract PDF(1083KB) icon1434 icon899
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Experiments were conducted on a typical rotor-stator system where air entered through an annular slot at low radius and flowed out of the cavity axially through a rim seal between the rotor and the stator. For the seal in this rotor-stator system, the stationary shroud overlapped the rotating one. Pressure distributions at the stator surface and flow resistance coefficients of the rotor-stator cavity with a maximum gap of 67mm were measured under different dimensionless mass flow rates from 1.32×104 to 4.87×104 with a large range of rotational Reynolds numbers from 0.418×106 to 2.484×106. The results show that pressure on the stator surface decreases with the increase of rotational Reynolds number when the dimensionless mass flow rate is below 1.3×104; when the dimensionless mass flow rate is above 3.034×104, the trend reverses. This is the so-called "pressure inversion effect". However, dimensionless pressure does not show the same changes when rotational dynamic pressure is chosen as the denominator. The resistance coefficient of the rotor-stator cavity is determined by the dimensionless mass flow rate and rotational Reynolds number; for practical application, the resistance coefficient can also be estimated by the turbulent flow parameter in the range of turbulent parameter from 0.1 to 1.6.
Infrared radiation characteristics of axisymmetrical ejector convergent nozzle in turbofan engine
SHI Xiao-juan, JI Hong-hu
2015, 30(4): 784-792. doi: 10.13224/j.cnki.jasp.2015.04.003
Abstract PDF(1027KB) icon1735 icon911
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The infrared radiation (IR) characteristics of axisymmetrical ejector convergent nozzle in turbofan engine were investigated with the method of numerical simulation. The flow fields of exhaust systems were calculated with commercial software, while the IR characteristics were calculated with self-developed software NUAA-IR. The IR characteristics of 3-5μm waveband of axisymmetrical convergent nozzle and ejector convergent nozzle and the IR contributions of different solid walls in the nozzles on different detection directions were calculated and analyzed. The results show that the infrared suppression of ejector convergent nozzle mainly lies in the mixing of plume and environment atmosphere, helping to minimize the length of the plume, and then reducing the IR of gas. The shielding and cooling effect of ejector convergent nozzle on the nozzle's solid walls is very small, and this effect is effective only to the medium and low temperature walls such as main duct wall and inner wall of bypass duct at the azimuth angles greater than 20 degree. The total integral IR intensity of ejector convergent nozzle is less than that of the convergent nozzle except in the azimuth angles of 0 to 15 degree where these two nozzles' total integral IR intensities are nearly equal. The maximum reduction amplitude is about 34% at the azimuth angles of 40 degree.
Dynamic response characteristic of concentric stage swirling structure
QIN Hao, DING Zhi-lei, LIN Yu-zhen, LI Ji-bao
2015, 30(4): 793-799. doi: 10.13224/j.cnki.jasp.2015.04.004
Abstract PDF(805KB) icon1183 icon928
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The natural flow characteristics and their response to outer excitations of a LESS (low emissions with stirred swirls) combustor head exit was studied by experiment. Experimental results shows that periodical flow structures are in the pilot swirling flow zone, while absent in the main outlet zone;the flow distortion in the pilot flow zone will happen when outer excitation frequency coupled with natural flow frequency, and the velocity oscillation amplitude of periodical flow will increase dramatically.
Experiment and prediction of separation performance of hollow fiber membrane
SHAO Lei, LIU Wei-hua, SUN Bing, ZHAO Hong-tao, FENG Shi-yu
2015, 30(4): 800-806. doi: 10.13224/j.cnki.jasp.2015.04.005
Abstract PDF(762KB) icon1345 icon896
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Based on a test apparatus of the on-board hollow fiber membrane separation performance, the separation performance was experimentally investigated with variation of the bleed air pressure, bleed air temperature, flight altitude and other factors; the experimental data were applied into the artificial neural network prediction technology as the training and validation samples, so as to predict the performance of this membrane. The results show that: (1)the established mathematical model can effectively predict the separation performance of hollow fiber membrane; (2)the volume fraction of nitrogen-enriched air is inversely proportional to non-dimensional flow rate; so when the volume fraction of nitrogen-enriched air increases, the flow rate and efficiency of nitrogen-enriched air decrease; (3)under a given volume fraction of nitrogen-enriched air, non-dimensional flow rate of nitrogen-enriched air increases with bleed air temperature and bleed air pressure increasing; the efficiency of nitrogen-enriched air increases with the bleed air pressure increasing, but decreases with bleed air temperature increasing; (4) when the flight altitude increases, the non-dimensional flow rate and efficiency of nitrogen-enriched air increase, and the influence of volume fraction of nitrogen-enriched air decreases with the growing flight altitude.
Experiment on effect of dome and dilution holes on outlet temperature distribution for triple swirler combustor
DING Guo-yu, HE Xiao-min, XUE Chong, HONG Liang
2015, 30(4): 807-813. doi: 10.13224/j.cnki.jasp.2015.04.006
Abstract PDF(870KB) icon1206 icon1000
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Experiments were conducted to study the outlet temperature distribution of combustor for four different triple swirler schemes and two different dilution holes arrangement schemes with different inlet velocities and fuel-air ratios at fixed atmospheric pressure. The results show that: the outlet temperature distribution factors of different schemes change little with the inlet velocity and fule-ratio ratio; when intermediate swirler and outer swirler are counter-rotating, the outlet temperature distribution becomes more uniform; for two kinds of swirl number combination of 0.7-1-1.5 and 1.5-1-0.8, the outlet temperature distribution factor of the former is more sensitive to the change of outer swirler direction than that of the latter; compared with scheme B, the dilution holes arrangement of scheme A is more suitable for the triple swirler combustor; the dome and dilution holes arrangement should be matched properly in order to optimize the outlet temperature distribution.
Experiment of combustion performance in LPP low emission combustor with single dome
YAN Ying-wen, DANG Long-fei, DENG Yuan-hao, XU Rong, XU Hua-sheng
2015, 30(4): 814-822. doi: 10.13224/j.cnki.jasp.2015.04.007
Abstract PDF(937KB) icon2093 icon944
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Combustion performance experiment was conducted for three-stage swirler of lean premixing and prevaporizing (LPP) low emission combustor with single dome. In order to investigate the effects of different fuel air ratios, inlet air mass flow rates, inlet air temperatures and positions of pilot atomizer on the combustion performance of combustor outlet, and the outlet temperature profile, combustion efficiency and pollutant emission rules of combustor were obtained. The experiment results show: (1) emission of NOx increases with the fuel air ratio increasing, and the combustion performance of dome A is better than that of dome B; (2) at the same fuel air ratio, if inlet air temperature increases or inlet air mass flow rate decreases, pollutant emission raises correspondingly; (3) position of pilot atomizer has effect on the combustion performance of LPP low emission combustor.
Film cooling performance with internal coolant channel crossflow
JIA Guang-sen, ZHANG Li, LU Cong-ming, LUO Jian-xia, HUANG Xiao-yang
2015, 30(4): 823-830. doi: 10.13224/j.cnki.jasp.2015.04.008
Abstract PDF(1170KB) icon1264 icon930
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In order to investigate the flow and heat transfer performance of film cooling with internal coolant channel crossflow, narrowband transient liquid crystal measurement technique was used to gain the contours of cooling effectiveness and heat transfer coefficient downstream the film cooling hole for blowing ratios of 0.5, 1 and 2 with internal coolant channel crossflow, and the detailed flow field characteristics inside film cooling hole and downstream region were obtained by numerical simulation. The results show that the internal coolant channel crossflow has a notable effect on distributions of cooling effectiveness and heat transfer coefficient downstream the film cooling hole. The crossflow enhances the ability of lateral spreading of film cooling hole injection, and the film cooling effect is improved at high blowing ratio. Furthermore, the asymmetric vortices appear downstream the film cooling hole, and the structures of vortices are more complicated.
Nozzle structure's influence on fuel flow oscillation in unsteady conditions
TANG Guan-qiong, QIN Hao, LI Lin, LIN Yu-zhen, LI Ji-bao
2015, 30(4): 831-837. doi: 10.13224/j.cnki.jasp.2015.04.009
Abstract PDF(833KB) icon1828 icon919
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Assuming that fuel flow is continuous under small oscillation conditions, a correlation model about fuel flow oscillation and the nozzle structure was derived with flow number as an intermediate variable. Under this model, conclusion are made as follows: fuel flow oscillation is proportional to the flow number, and inversely proportional to the 0.5th power of average fuel pressure drop when the external excitation is given. Methods of decreasing fuel flow oscillation by decreasing flow number are obtained by nozzle calibration: increasing the throttling stage, decreasing throttling area. This method is validated through oscillation experiment. A general design procedure is suggested for nozzle designers in face of the instability problem.
Secondary flow control in high-turning compressor cascade using vortex generator jet
LIU Hua-ping, CHEN Huan-long, LI De-xiong, GUO Yu-jie, CHENG Xiao-qi, CHEN Fu
2015, 30(4): 838-845. doi: 10.13224/j.cnki.jasp.2015.04.010
Abstract PDF(1272KB) icon1895 icon914
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End wall secondary flow control using vortex generator jet in compressor cascade with the turning angle of 60 degree was performed. The effects of variation of the direction and the total pressure of the jet on aerodynamic performance and flow of compressor cascade were presented. The results indicate that with the skew angle of 0 degree in vortex generator jet and jet mass flow rate less than 0.5% of inlet mass flow rate of compressor cascade, the loss of cascade could be reduced obviously. The vortex induced by the vortex generator jet could weaken or even prevent the development of passage vortex, make the fluid of the main stream involve into the end wall boundary layer in the down-washed region and guide the low energy fluid to the main stream in the up-washed region. Thus the accumulation of the low energy fluid in the corner region was reduced, resulting in the decrease of the separation on the suction surface. Using the same jet total pressure at inlet with that of the compressor cascade inlet, the reduction of 21.5% of total pressure loss can be obtained; this effect can become more obvious with the increase of the jet total pressure at inlet.
Engineering research on prediction and suppression of blade flutter in compressor fan
SUN Hai, LI Jian, YANG Lin, CHEN Bao-shi, LIU Yi-xiong
2015, 30(4): 846-853. doi: 10.13224/j.cnki.jasp.2015.04.011
Abstract PDF(954KB) icon1553 icon985
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The aeroelastic stability prediction and the flutter margin point were achieved by energy method in original scheme of prototype compressor fan blade. The geometric modeling of aeroelastic unstable blade was modified to improve the aeroelastic stability. By comparison of blade geometric modeling, aerodynamic performance, vibration characteristic and aeroelastic stability in original and modified schemes, the blade flutter mechanism of compressor fan was analyzed on aerodynamic aspect. The effective engineering flutter suppression means for the compressor fan blade are achieved, such as increasing the chord length of blade and decreasing the aspect ratio, increasing the blade thickness and stiffness, decreasing the angle of incidence and improving the flow.
Influence of tip clearance heights on the flow field and performances of a supersonic expander
HUANG Zhen-yu, ZHONG Jing-jun, YANG Ling, HAN Ji-ang
2015, 30(4): 854-864. doi: 10.13224/j.cnki.jasp.2015.04.012
Abstract PDF(1331KB) icon1535 icon836
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The three-dimensional Reynolds-averaged Navier-Stokes equations and the standard k-ε turbulent model were adopted to simulate numerically the flow field and performance of a supersonic expander with different tip clearance heights. The results show that tip clearance height significantly impact on the local flow characteristics in the three-dimensional flow passage and the overall performances of the supersonic expander. With the increasing of the tip clearance height, the highest relative Mach number of airflow decreases, and the scope of high speed reduces. The leakage vortex strengthens, scale largens, transverse and radial movement significantly, and the leakage loss increased, but the loss of shock waves and interaction of shock and boundary layer reduced. Expansion ratio of supersonic expander firstly increased and then decreased, and the isentropic adiabatic efficiency reduced gradually. The main source of loss in the three-dimensional flow passage of supersonic expander with clearance is the friction loss of low energy fluid near the suction surface and bottom wall and the leakage loss. The scope of clearance height should be selected between 0.9%h0 and 1.5%h0.
Numerical simulation of effects of contoured slot seal configuration on turbine performance
ZHANG Jing-hui, MA Hong-wei
2015, 30(4): 865-874. doi: 10.13224/j.cnki.jasp.2015.04.013
Abstract PDF(2159KB) icon1527 icon684
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Effects of rim seal flow through uniform slot seal and contoured slot seal configurations on the turbine performance were investigated numerically. The results show that the gas ingress and egress structures are affected by stator disc, rotor disc and the tangential velocity in mainstream, and the rotating speed is less than that of rotor disc with the same direction. As a result, the flow patterns of rotor are changed, which strengthened the pressure side horseshoe vortex, thereby resulting in large changes of flow field of rotor exit. The interaction of seal flow and upstream vane weak induces the entropy rise in the rotor passages, leading to decrease of the turbine efficiency. Compared with uniform slot seal configuration, the contoured slot seal configuration collapses the big ingress and egress structures into smaller ones, and then has less negative effects on the turbine performance. The turbine efficiency is improved by 0.9%. The results prove that the contoured slot seal configuration has better turbine performance as well as better seal effectiveness.
Relationship between optimum curved blade generate line and blade camber angle in linear compressor cascade
LING Jing, DU Xin, WANG Song-tao, WANG Zhong-qi
2015, 30(4): 875-882. doi: 10.13224/j.cnki.jasp.2015.04.014
Abstract PDF(1020KB) icon1734 icon943
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The relationship between the optimum curved blade generate line and blade camber angle in linear compressor cascade was researched by optimization method. Curved angle and curved height of curved blade generate line were optimized at eight different blade camber angles. The stack line was composed of two Bezier curves and a straight line, cascade loss decreases with the increasing curved height at the same curved angle, and the optimum curved height of the curved blade is 0.5. As for an optimum curved angle, at which curved blade total loss is minimal, curved blade loss decreases with increasing curved angle when curved angle is less than curved angle, and loss grows with increasing optimum curved angle when curved angle is larger than optimum curved angle. The benefits of the curved blade improve with increasing blade camber angle. Optimum curved angle increases along with the increasing blade camber angle. The relationship of optimum curved angle and blade camber angle presents a similar linear type.
Characteristics under variant working conditions of a supersonic low-reaction aspirated compressor
ZHANG Long-xin, WANG Song-tao, RUAN Guo-hui, LIU Xun, WANG Zhong-qi
2015, 30(4): 883-890. doi: 10.13224/j.cnki.jasp.2015.04.015
Abstract PDF(1194KB) icon1232 icon766
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In order to explore the aerodynamic performance at off-design points, taking an supersonic first stage of three-stage highly-loaded low-reaction aspirated compressor as research objection, the variation of parameters and inner flow filed for an supersonic first stage operating at various speeds, aspiration mass flow rates, and near stall point were investigated by numerical simulations. As a result, the aspiration schemes of the aspirated compressor designed with the low-reaction concept can still meet the design demands at various speeds; the reduction of aspiration mass flow rate will not only decrease the stage aerodynamic performance, but also affect the matching with the next stage; the working range of aspirated compressor can be further broadened by adding additional aspirated slot/holes before the endwall aspiration slot.
Optimized profile of high subsonic axial flow compressor
CHEN Zhong-liang, JIANG Bin, ZHENG Qun, ZHANG Xiao-long, TAN Chun-lai
2015, 30(4): 891-900. doi: 10.13224/j.cnki.jasp.2015.04.016
Abstract PDF(1048KB) icon1563 icon811
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An optimization design method of compressor profile was studied by combining computational fluid dynamics and numerical optimization algorithm. The profile of a high subsonic axial flow compressor was taken as the research object with inlet Mach number of 0.7. The optimization variables were selected by Latin hypercube sampling method and an objective function considering incidence performance was built. Boundary layer transition was also taken into account and the Gamma-Theta transition model was used in numerical simulation. An optimized profile of high subsonic axial flow compressor that can improve incidence performance and reduce total pressure loss was obtained. The results show that the optimized profile can significantly reduce total pressure loss at both incidences of -4 degree and +4 degree with inlet Mach number from 0.2 to 0.8. It can increase the incidence of lower pressure loss by more than 4 degree at the design point (inlet Mach number of 0.7). In addition, the optimized profile can reduce the optimum consistency by 20% and improve the flow characteristics of cascade at low Reynolds numbers.
Effect of inlet preswirl on the aerothermodynamic performance of splitter mixing exhaust system
WU Fei, XIE Yi, SHAO Wan-ren, LIU Kun, RUAN Deng-fang, ZHENG Biao
2015, 30(4): 901-909. doi: 10.13224/j.cnki.jasp.2015.04.017
Abstract PDF(1020KB) icon1544 icon815
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In order to clarify the effect of core flow inlet preswirl angle of mixing exhaust system on the aerothermodynamic performance of splitter mixing exhaust system, investigation on the flow field of splitter mixing exhaust system at different inlet preswirl angles was carried through the three-dimensional numerical simulation based on Navier-Stokes equations, and the results were compared with flow along the axis. According to the results, the thickness of shear layer increases with the inlet preswirl angle, and it is the same with the twist of shear layer. Besides, at the splitter mixing exhaust system exit, as the inlet preswirl angle increases, the thermal mixing efficiency increase slightly and thermal mixing efficiency of 30 degree inlet preswirl angle model is 1.36 times the model with flow along the axis. Moreover, total pressure recovery coefficient and thrust of splitter mixing exhaust system also decline. Compared with the model with flow along the axis, the total pressure recovery coefficient of 30 degree inlet preswirl angle model falls by 0.0034 and its relative thrust declines by 0.081.
Air bridge technology for engine power simulation test in wind tunnel
ZHANG Rong-ping, WANG Xun-nian, HUANG Yong
2015, 30(4): 910-915. doi: 10.13224/j.cnki.jasp.2015.04.018
Abstract PDF(815KB) icon1995 icon905
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The design technology and correction method of the air bridge for engine simulation test in wind tunnel were presented. The layout of the air bridge was determined according to freedom analysis. By finite element analysis, the key beams of air bridge were optimized. The rigidity of the air bridge and the balance was matched based on air bridge and balance assembly numerical simulation. These design and optimization minimized the applied force of the air bridge and improved the capacity for overcoming pressure and temperature effect. The rigidity effect, pressure effect, temperature effect and mass flow effect of the air bridge were corrected by serial tests. These corrections further minimized residual force of the air bridge. A high aspect ratio airplane full-span turbofan powered simulators (TPS) test was conducted in 8 m×6 m low speed wind tunnel .The test results were good in repeatability. The standard deviation of drag coefficient was 0.0003. The test results also agreed well with related papers. The air bridge technology is thus proved successful and can meet the requirement of TPS nacelle power simulation test.
Effect of lobe number on aerothermodynamic performance of lobed S-shaped two-dimensional nozzle
DU Li-wei, LIU You-hong, SHAO Wan-ren, XU Su, DENG Hong-wei
2015, 30(4): 916-926. doi: 10.13224/j.cnki.jasp.2015.04.019
Abstract PDF(1129KB) icon1593 icon841
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Based on a lobed S-shaped two-dimensional nozzle of a turbofan engine, the length, inner divergence angle, outer divergence angle and ratio of high to width of lobed mixer were kept unchanged only with the lobe number fixed as 12, 14, 16, 18, 20. A group of lobed S-shaped two-dimensional nozzle models with different lobe number were built. By the use of validated CFD simulation, the effect of lobe number on aerothermodynamic performance of lobed S-shaped two-dimensional nozzle was investigated. The results show that within the flow field between the lobe trail cross section and the first S bend, the degree of mixing is highly affected by the lobe number and the thermal mixing efficiency increases as a response to the increment of lobe number. Within the flow field between the first S bend and the exit of the lobed S-shaped two-dimensional nozzle, the lobed S-shaped two-dimensional nozzle model with 16 lobes always has the lowest value of total pressure recovery coefficient and neither thermal mixing efficiency nor total pressure recovery coefficient has distinct change between other models. At the exit of lobed S-shaped two-dimensional nozzle, the lobed S-shaped two-dimensional nozzle model with 16 lobes has the highest thermal mixing efficiency of 0.850, while total pressure recovery coefficient of the model has dropped by 0.289% when compared with the model of the biggest total pressure recovery coefficient. Besides, the gradual narrowing flow passage of the lobed S-shaped two-dimensional nozzle can contribute to better mixing and more rapid consumption of stream-wise vortex.
Prediction model of cross-flow instability transition in swept wing boundary layers
XU Jia-kuan, BAI Jun-qiang, QIAO Lei, HUANG Jiang-tao
2015, 30(4): 927-935. doi: 10.13224/j.cnki.jasp.2015.04.020
Abstract PDF(1097KB) icon1737 icon840
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Through the theoretical analysis and numerical solution of classic Falkner-Skan-Cooke three-dimensional boundary layer similar solution, combining with the thought of two-dimensional boundary layer transition criteria, using the C1 criteria calibrated by the experimental data to get the cross-flow instability transition displacement thickness Reynolds number, the cross-flow transition criteria was established for fixed swept angle of wing leading edge by solving the equations and data fitting. The model was applied to conduct the numerical simulation of cross-flow instability transition on the ONERA-M6 wing with 30 degree swept angle of leading edge and NLF (2)-0415 infinite swept wing with 45 degree swept angle of leading edge. The simulated results show that the improved transition model can predict the location of cross-flow instability transition of swept wing precisely and be in good agreement with the experimental data. Therefore, the results indicate that the cross-flow instability transition criterion built is practical and valuable.
Numerical simulation of implicit fully coupled SST and TNT turbulence models for high speed flows
XIA Chen-chao, CHEN Wei-fang, GUO Zhong-zhou, NIE Liang
2015, 30(4): 936-943. doi: 10.13224/j.cnki.jasp.2015.04.021
Abstract PDF(1066KB) icon2759 icon764
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SST(shear stress transport) and TNT (turbulent/non-turbulent) turbulence models were solved by implicit fully coupled with the transport and mean flow control equations. Techniques of local time step and implicit treatment of turbulence source term were used to accelerate and stabilize the calculation. Test cases of hypersonic compression corner flow, conical cylinder flare flow and supersonic asymmetric shock wave/boundary layer interaction were simulated by the AUSMPW+ (AUSM by pressure-based weight functions) scheme and LU-SGS (lower-upper symmetric Gauss-Seidel) implicit fully coupled method. The results show that:SST and TNT turbulence models used can predict the wall pressure and heat flux of the wall at the 15 degree compression corner flows well. Discrepancies between calculated and experimental results increase with the increase of compression corner. Compressibility correction has a great effect on the pressure and heat flux of compression corner flows and has little effect on supersonic asymmetric shock wave/boundary layer interaction. The implicit fully coupled method shows better convergence characteristic than explicit coupled method.
Modeling of hypersonic vehicles propulsion system
XIAO Di-bo, LU Yu-ping, YAO Ke-ming, LIU Yan-bin, CHEN Bo-yi
2015, 30(4): 944-951. doi: 10.13224/j.cnki.jasp.2015.04.022
Abstract PDF(1072KB) icon1697 icon927
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In order to quickly obtain the thrust and moment of thrust at the early design stage, and satisfy relevant modeling and analysis control, a model for the propulsion system of hypersonic vehicle was proposed. An inlet model was developed based on the wave interaction method. Moreover, the dual-mode combustor was modeled by an increasing area duct with heat addition and friction, whereas the internal nozzle was modeled by a variable area duct with friction. Accordingly, the thrust was estimated according to the momentum theory, and then analytical expression was obtained by a curve-fitting method. Compared with the result of CFD, the error of Mach number and temperature at the dual mode ramjet engine inlet was less than 5%, and the pressure error was less than 10%; the computed thrust increased as Mach number, stoichiometrically normalized fuel-to-air ratio and angle of attack increased, and decreased as height increased; the time cost of one state point was less than 0.5s averagely. Computation result demonstrates that the proposed modeling approach meets the high efficiency and accuracy required in the control-oriented modeling process, contributing to the analysis and design related to dynamics and control of the hypersonic vehicles.
A new blade-casing rubbing model and its verification
CHEN Guo, WANG Hai-fei, LIU Yong-quan, FENG Guo-quan, JIANG Guang-yi, LI Cheng-gang, WANG De-you
2015, 30(4): 952-965. doi: 10.13224/j.cnki.jasp.2015.04.023
Abstract PDF(1453KB) icon2154 icon862
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For the aero-engine blade-casing rubbing fault, a new blade-casing rubbing model was put forward. On the basis of tranditional elastic rubbing model, the new model considerd the effects of blade number and the rotor-stator clearance change on rubbing forces, and it could simulate rubbing faults with various rubbing positions including single-point, multi-points, part and whole-cycle on casing and rotor. The new rubbing model was applied to the rotor-support-casing coupling dynamic model, and the casing acceleration responses under rubbing faults were obtained by the numerical integration approach. The aero-engine rotor tester including casing was used to carry out the rubbing test whose rubbing positions were the single-point on casing and whole-cycle on rotor, and it is found that the rubbing characteristics of the casing vibration acceleration signals have obvious periodic impact characteristics. The impact frequency is the frequency of the blades passing the casing, which equals to the product of the rotational frequency and blade number. There is a blade passing frequency and its multiple-frequency components in the high-frequency band of spectrum, and the impact is modulated by the rotational frequency. There are obvious quefrency components of rotational frequency components and its multiple frequencies in the signal cepstrum.The simulation results agree well with the test values, and the new blade-casing rubbing model is fully verified. Finnaly, based on the new blade-casing rubbing model, the other rubbing modes between rotor and casing were simulated, and the casing vibration charateristics were obtained.
Damping mechanism of floating ring squeeze film damper
ZHOU Hai-lun, FENG Guo-quan, LUO Gui-huo, AI Yan-ting
2015, 30(4): 966-971. doi: 10.13224/j.cnki.jasp.2015.04.024
Abstract PDF(875KB) icon1612 icon845
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In order to research on damping mechanism of floating ring squeeze film damper, the model of floating ring squeeze film damper was established based on the Reynolds equation of squeeze film damper and relative movement of floating ring and journal. Then damping mechanism of floating ring squeeze film damper was researched by means of displacement response of rotor system and bidirectional incentives test system of floating ring squeeze film damper. It is shown that the greater mass of floating ring, the thinner oil clearance, the wider oil film width or the greater viscosity of lubrication oil, the better damping performance of floating ring squeeze film damper, vice versa.
Experiment on vibration control of rotor with ring dynamic vibration absorber
ZHANG Bing-kang, HE Li-dong, YANG Xiu-feng, ZHANG Zhen-kun
2015, 30(4): 972-978. doi: 10.13224/j.cnki.jasp.2015.04.025
Abstract PDF(890KB) icon1264 icon759
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The application of ring dynamic vibration absorber in vibration control of rotor was studied by experiment. A kind of separable ring dynamic vibration absorber placed on bearing was designed based on single disc rotor model and vibration characteristics. Vibration absorber can be installed on optional position between two bearings without changing any support structure and affecting dynamic performance of rotor. The rotor experiment rig of vibration reduction with vibration absorber was built to measure the vibration with critical speed of rotor. The results show that vibration absorber can decrease vibration greatly, and the vibration amplitude is dropped by 70% at critical speed of rotor. The frequency range of vibration reduction can be broadened and the natural frequency of vibration absorber can be adjusted accurately by increasing the mass of vibration absorber. The relationship between vibration reduction effect and installation position of vibration absorber was studied by experiment. The results show that position of large amplitude is better than other area. The vibration reduction effect with different numbers of vibration absorbers was compared under rotor imbalance vibration, and the result indicates that the effect of two vibration absorbers is better than that of one.
Frequency domain optimal tuning method of PI control parameters for aero-engine
HUANG Jin-quan, LIU Nan, TANG Yu-ting
2015, 30(4): 979-984. doi: 10.13224/j.cnki.jasp.2015.04.026
Abstract PDF(862KB) icon1292 icon735
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For the aero-engine high and low rotor speed control loops and pressure ratio control loop, the method of determining the stability region of the proportion-integration (PI) control parameters satisfying the gain margin and phase margin was researched. According to the system stability analytical model in frequency domain, the stability region graphic of system parameters in frequency domain with PI controller considered was established. The optimal objective function was defined in terms of the typical dynamic performance indices and robustness indices, and a set of optimal PI control parameters were found in the obtained stability region of control parameters. This method was applied to controller parameters design of aero-engine for the high and low rotor speed control loops and pressure ratio control loop, and PI controllers under different states were designed. Simulation results demonstrate that the tuning method satisfies performance indices and robustness of closed loop system.
LQ/H∞ controller design for aero-engine based on improved NSGAⅡ
WANG Yuan, LI Qiu-hong, HUANG Xiang-hua, WU Qian
2015, 30(4): 985-991. doi: 10.13224/j.cnki.jasp.2015.04.027
Abstract PDF(877KB) icon1244 icon704
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The non-dominated sorting genetic algorithm Ⅱ (NSGAⅡ) was improved, and the traditional genetic operation was replaced by differential evolution (DE) algorithm with the non-dominated order as evolution direction. The convergence speed of the algorithm was faster. In order to solve the conservativeness of linear quadratic regulator (LQR) control and H2/H∞ control which had caused wide public concern recently in the field of aero-engine multi-variable control, the LQ/H∞ controller of aero-engine was designed by combining three performance indexes: time domain index, quadratic index and H∞ index and improving the weight matrixes Q, R of NSGAⅡoptimization. Compared with other controller design methods, the LQ/H∞ controller based on multi-objective optimization has clearer control objective, stronger robustness and less conservativeness. The simulation results show that, compared with the H2/H∞ controller based on linear matrix inequation (LMI), the LQ/H∞ controller based on multi-objective optimization can increase performance four times in time-domain and anti-interference ability by 15%, and decrease the regulating time by 50%.
Influential factors of nozzle performance of hot water rocket motor
SUN Wei-wei, WEI Zhi-jun, WANG Ning-fei
2015, 30(4): 992-998. doi: 10.13224/j.cnki.jasp.2015.04.028
Abstract PDF(851KB) icon1178 icon858
Abstract:
In order to learn the influential factors of nozzle performance of hot water rocket motor deeper, numerical simulation of the flow field in the nozzle were performed at different parameter conditions. The impact rules of different initial conditions such as pressure, vapor volume fraction, supercooling and different nozzle structures such as expansion ratio, converging half-angle and expansion half-angle to the thrust were investigated. The calculating results show that: the greater the initial pressure, the greater the thrust; as the initial vapor volume fraction increases, thrust has a process of increase first and then decrease; thrust is larger when the initial temperature of hot water rocket motor is closer to the saturation temperature; thrust has a process of increase first and then decrease with the increasing of expansion ratio; the impact of converging half-angle to the thrust is minor; the greater the expansion half-angle, the smaller the thrust.
Experiment on thrust adjustment characteristics of fluidic nozzle throat
GUO Chang-chao, YU Xin-yu, LI Bo, XIE Kan, WEI Zhi-jun, WANG Ning-fei
2015, 30(4): 999-1007. doi: 10.13224/j.cnki.jasp.2015.04.029
Abstract PDF(1091KB) icon1208 icon769
Abstract:
A cold-flow test research of fluidic nozzle throat was performed using the gas and water as secondary flow medium to study the thrust adjustment characteristics of fluidic nozzle throat of solid rocket motor. The influence of different secondary flow mediums, injection ways and injection mass flow rates on thrust response time, choke performance, thrust angle and thrust efficiency were studied. The experimental results show that the thrust response time is shorter using liquid secondary flow; the choke performance and thrust angle are related to injection positions and injection angle of secondary flow, and increase with increasing mass flow rate ratio; under the same mass flow rate ratio, thrust performance by using gaseous secondary flow is better than that by liquid secondary flow; but liquid secondary flow needs smaller pressure ratio and larger mass flow rate ratio adjustment range when the same mass flow rate ratio is given.
Mixed lubrication analysis in finite tooth contact surface of filtering reducer
PU Wei, WANG Jia-xu, ZHOU Guang-wu, LI Jun-yang, QIN De-cheng
2015, 30(4): 1008-1015. doi: 10.13224/j.cnki.jasp.2015.04.030
Abstract PDF(1153KB) icon1617 icon653
Abstract:
The numerical analysis model for mixed lubrication of filtering reducer tooth was built considering the tooth contact geometry, normal load of meshing point, tooth parameters, real surface roughness, rheological behavior of lubricant. A completely numerical solution of mixed lubrication analysis in tooth contact surface was achieved by solving the Reynolds equations and elastic deformation equations using compound iteration method and fast Fourier transformation (FFT) method respectively. The pressure and average film thickness were obtained at engaging-in point, engaging-out point and pitch point of filtering reducer, and the effects of speed on the average film thickness and contact area ratio in tooth contact surface were also obtained. The results show that the pressure distribution along the tooth width direction is not equal at engaging-in point, engaging-out point and pitch point, and pressures on both sides are higher; the variation of pressure is smaller on both sides and the average film thickness is greater at engaging-in point. When the tooth contact surface is rough, with the decrease of speed of filtering reducer, the average film thickness at engaging-in point, engaging-out point and pitch point is decreased, and the contact area ratio increases, leading to poor lubrication. Besides, the average film thickness of meshing points of rough tooth contact surface is less than that of smooth tooth contact surface under the same conditions.
Design of disturbance rejection control system for 5-DOF magnetic bearing
LI Ke-xiang, DENG Zhi-quan, LIU Cheng-zi, HUA Chun
2015, 30(4): 1016-1024. doi: 10.13224/j.cnki.jasp.2015.04.031
Abstract PDF(1059KB) icon1477 icon721
Abstract:
The disturbance rejection control system for 5 degree-of-freedom (DOF) magnetic bearing was designed based on proportion-integration-differentiation (PID) control combined with disturbance rejection process. Through analysis of the rotor motion model, a more accurate disturbance observer model of magnetic bearing was revised and obtained. The stability analysis of the discrete disturbance observer was conducted and the parameter tuning method was elaborated, making the parameter tuning process of disturbance observer more normative and convenient. The experimental results show that: when the disturbance rejection control is employed in PID control, displacement fluctuations of rotor disturbed DOFs are reduced by 38.7% and 67% respectively, and the motion track fluctuation of the rotor axis center is reduced by 21% with rotating speed of 20000r/min, thus verifying the superiority of the disturbance rejection control compared with PID control.

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