2016 Vol. 31, No. 12

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Influence of curvature on flow stability in circular pipe
LI Jia, GAO Jun
2016, 31(12): 2817-2823. doi: 10.13224/j.cnki.jasp.2016.12.001
Abstract PDF(1035KB) icon1321 icon773
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In order to research the influence of curvature on flow stability in circular pipe, two circular pipe models of axial flow of spanwise curvature and circumferential flow of axial curvature were studied. For two models, the stabilities and disturbance evolutions of difference curvatures were calculated by the linear stability theory. And the results were compared with channel flow. Then the influences of curvature on basic flow, eigenvalue and eigenfunction were obtained. Result showed that, along with the greater curvature, the basic flow would deflect inside, the axial wavenumber of disturbance wave became greater, the growth rate changed smaller, the unstable region of neutral curve diminished and symmetry of eigenfunction got worse. So curvature has a stabilizing influence on disturbances in boundary layers and the stability will be better when the curvature increases gradually in certain rang.
Effects of mutational load on performanceof gas turbine generator set
ZHANG Ya-dong, JIANG Li-yun, HAN Xiao-guang, XU Xing, ZHANG Ke
2016, 31(12): 2824-2832. doi: 10.13224/j.cnki.jasp.2016.12.002
Abstract PDF(1022KB) icon1658 icon782
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According to the fast response and real-time simulation requirements of mutational load, a split shaft gas turbine power generation model was built using volumetric method to better accommodate rapid load changes. The control rules of step changes in load to be adopted and the cascade proportion integration differentiation (PID) control policy were presented. The key influential factors to dynamic performance of step changes in load were discussed. Simulation results show different limitations of parameters bring different effects on the gas turbine dynamic performance. Selecting limitations of parameters reasonably can optimize gas turbine dynamic performance. Results of model were presented and compared with the test date, indicating the model and engine agreed well for each other. The results show the model has good real time, convergent and dynamic response property. Theoretical analysis and test result demonstrate the fuel limitation curve should be operated in 1s after rapid load changes to achieve better dynamic performance. The minimum fuel limitation is necessary for load rejection, and reasonable selection of the limit value is the key to ensure stable combustion and control speed overshoot.
Aerodynamic characteristics and lift enhancement mechanism of mini-TED in transonic flow
JI Le-qiang, DING Dong-jian, LI Jian-bo, CUI Zhao
2016, 31(12): 2833-2841. doi: 10.13224/j.cnki.jasp.2016.12.003
Abstract PDF(1227KB) icon1313 icon650
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By means of solving two dimensional and compressible Navier-Stokes equations, the transonic flow-field characteristics of NACA0012 airfoil equipped with mini trailing edge lift enhancement device (mini-TED) was investigated, and the influences of geometrical parameters on the vortices and aerodynamic characteristics of airfoil were analyzed and compared with the Gurney flap(GF). The two orthorhombic geometry detailed parameters were defined as chord-wise length and effective height to describe the geometric features. Only the chord-wise length changed at the same angle of attack, although the trailing edge vortices varied greatly, the aerodynamic characteristics of airfoil had little change. On the other hand, the trailing edge vortices and the aerodynamic characteristics of airfoil varied significantly with the effective height changed, and the aerodynamic characteristics were very close to the GF in the same condition of effective height. Result shows that the effective height plays a dominant role in affecting the aerodynamic characteristics of airfoil. Rather than the microstructure and shape, the size of the vortices is the main factor influencing the flow field of airfoil as this is mainly decided by the effective height. With the mini-TED, compared with the original airfoil, the shock wave on the upper surface of the airfoil moves backward and the strength of the lower shock wave is weakened, so the pressure coefficient of airfoil surface changes. With the increasing effective height, the induced vortices scope becomes larger, enhancing the drainage effect and increasing the lift coefficient, drag coefficient and nose-down pitching moment coefficient. The lift-drag ratio is improved at the same angle of attack as well.
Impact of variable area bypass injector on variable cycle engine performance
ZHOU Hong, WANG Zhan-xue, LIU Zeng-wen, ZHANG Ming-yang
2016, 31(12): 2842-2850. doi: 10.13224/j.cnki.jasp.2016.12.004
Abstract PDF(1044KB) icon1755 icon749
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The geometry regulating laws of forward and rear bypass injectors of variable cycle engine were analyzed. A mathematic model capable of calculating the variable cycle engine steady and transition performance was established. It considered the local total pressure loss caused by sudden area change at forward bypass injector and mode selector valve. The impacts of geometry regulating laws of forward and rear bypass injectors on variable cycle engine performance and stability were simulated with single bypass mode, double bypass mode and mode transitioning, and the simulation results were compared with NASA test data. The results indicate the opening of forward bypass injector's inner bypass area can decrease recirculation margin and increase fan and core drive fan stage surge margin in single bypass mode. During the late period of transforming from single into double bypass mode, the reduction of fan pressure ratio will diminish by closing rear bypass injector's outer bypass area. It's suggested that the forward bypass injector's inner bypass area should be changed, and the total area should be kept unchanged, while the rear bypass injector's outer bypass area should be changed, and the inner bypass area should be kept unchanged.
Calculation and dynamic characteristics analysis for the driving moment of three bearing swivel nozzle
LIU Shuai, WANG Zhan-xue, ZHOU Li, LIU Zeng-wen
2016, 31(12): 2851-2858. doi: 10.13224/j.cnki.jasp.2016.12.005
Abstract PDF(1136KB) icon1323 icon726
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The dirving moment calculation model of each nozzle segment of the three bearing swivel nozzle(3BSN) was built based on the deflection law model and the given deflection law of the 3BSN. The changing law of the moment produced by each nozzle segment's gravity and gas aerodynamic force to the instantaneous rotation axis with the varying nozzle deflection angle was obtained by respective force analysis for each nozzle segment. The effect of the nozzle geometrical parameters on the moment to the nozzle instantaneous rotation axis was obtained by changing the realizable maximum nozzle deflection angle and the length of the second and third nozzle segments. The maximum moment of the second and third nozzle segments' aerodynamic force to the instantaneous rotation axis increased with the realizable maximum deflection angle of the nozzle and occurred at the location in approximately 70% of the maximum deflection angle. The moment of the second and third nozzle segments' aerodynamic force to the instantaneous rotation axis increaseed with the nozzle segment's length increases.
Evaluation on turbulence models in simulating complex flows around high-lift airfoil
TANG Yu-meng, LIU Yang-wei, LU Li-peng
2016, 31(12): 2859-2869. doi: 10.13224/j.cnki.jasp.2016.12.006
Abstract PDF(1555KB) icon1521 icon751
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The flows around the three-element high-lift airfoil under the approaching conditions using seven eddy-viscosity turbulence models commonly applied to engineering in Fluent software were simulated numerically. Through a systematic comparison of the simulation results with the experiment results and the corresponding delayed detached eddy simulation (DDES) results, including the pressure coefficient, Mach number, vorticity and turbulent kinetic energy of airfoil, conclusions were drawn about their predicting abilities for this airfoil. Results show that, the standard k-ω model has the best predicting abilities for the averaged flow field, providing a better simulation for the pressure coefficient around the airfoil, the separating point and separating zone magnitude on the flap; the shear stress transport (SST) k-ω model also has the same abilities. The one-equation SA(Spalart-Allmaras) model and four-equations v2-f model shares similar abilities; but the simulation results given by the k-ε series models seem worse compared to all. As for the predicting abilities for the turbulent kinetic energy, conclusions similar to the averaged flow field are drawn. However, the "largest" turbulent kinetic energy distribution around the separating zone on the flap isn't captured by all the turbulence models.
TRPIV experimental investigation of drag-reductionmechanism in turbulent boundary layerover superhydrophobic surfaces
WANG Er-dan, TIAN Hai-ping, ZHANG Jing-xian, YAO Zhao-hui, JIANG Nan
2016, 31(12): 2870-2877. doi: 10.13224/j.cnki.jasp.2016.12.007
Abstract PDF(1226KB) icon1595 icon653
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Drag-reduction in turbulent boundary layer (TBL) over superhydrophobic surfaces was investigated by the time-resolved particle image velocimetry (TRPIV). Time series of velocity vector fields in TBL over superhydrophobic surfaces and hydrophilic surfaces were measured in a cyclical water channel. The distributions of mean velocity profile, Reynolds shear stress and turbulence intensity along wall-normal direction were acquired at the same free-stream velocity (0.17m/s). The center of the head of hairpin vortex in wall turbulence was detected with the multi-scale spatial locally-averaged vorticity. Conditional sampling and phase average methods were utilized to extract the spatial topologies of spanwise vortices at different normal positions. Streamwise fluctuation velocity and the pitching angle of spanwise vortex in both cases were compared. The features of streamline topologies around spanwise vortex were analyzed from the saddle-focus dynamical system. Results revealed that a drag reduction of 10.1% is acquired if Reynolds numbers are about 13500. For superhydrophobic surfaces, the average velocity seems to be increased, Reynolds shear stress is decreased, streamwise turbulence intensity is weakened, the pitching angle of spanwise vortex is smaller, and the development of coherent structure near wall region is suppressed.
Reduced mechanism of surrogate fuel for RP-3 aviation kerosene and verification
YAN Ying-wen, DAI Chao, LI Jing-hua
2016, 31(12): 2878-2887. doi: 10.13224/j.cnki.jasp.2016.12.008
Abstract PDF(1023KB) icon1466 icon709
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N-decane was chosen as the surrogate fuel of RP-3 aviation kerosene, and a reduced mechanism (36 species and 62 steps) was established. A Bunsen burner with premixed pre-evaporation combustion was designed to study the combustion flame of RP-3 aviation kerosene. And two verified kerosene simplified mechanisms (23 steps and 38 steps respectively) and the 62 steps mechanism were used by numerical simulation. The numerical results were compared with the experimental data. Results show that, the distributions of temperature and volume fraction of CO2 first increase and then decrease in the axial direction, and the temperature reaches the maximum at axial distance from nozzle of 0.020m.While the volume fraction of O2 first decreases and then increases, the volume fraction reaches the minimum at axial distance from nozzle of 0.025m. Compared with the 38 steps and 23 steps reduced mechanisms, the distributions of temperature and volume fraction of O2 are in good agreement with the experiment data when 62 steps reduced mechanism is applied. At the same time, the numerical results with 62 steps reduced mechanism can also well simulate the volume fraction of CO2, but the numerical results of 23 steps and 38 steps can only predict the change trend of the experimental results. Therefore, n-decane can be used as the surrogate fuel of RP-3 aviation kerosene and the 62 steps reduced mechanism can accurately predict the combustion characteristics of RP-3 aviation kerosene.
Flow field analysis for confined flames characteristics of counter dual-stage swirler
ZENG Qing-hua, KONG Wen-jun
2016, 31(12): 2888-2894. doi: 10.13224/j.cnki.jasp.2016.12.009
Abstract PDF(1145KB) icon1087 icon677
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Flow characteristics of the confined combustion flames of counter dual-stage swirler were analyzed. Research found that,the flow characteristics of the counter dual-stage swirler changed with the confinement ratio variation. When the confinement ratio was less than 3.8, the center backflow bubble on this flow field grew longer as the confinement ratio increased, with the growth law the same as that of the center backflow bubble on the single-stage swirler flow field. However, as the confinement ratio continued to increase, the growth law of the center backflow bubble was different between the counter dual-stage swirler and the single-stage swirler. The center backflow bubble on the single-stage swirler flow field continued to grow longer, while the center backflow bubble on the counter dual-stage swirler flow field was split into two different backflow bubbles. This result has deepened the understanding of the effect of confinement on flow and combustion, and provided a guidance for the optimal design of the counter dual-stage swirler.
Effect of wedge film cooling on the infrared signature of two-dimensional wedge nozzle
ZHOU Bing, JI Hong-hu
2016, 31(12): 2895-2903. doi: 10.13224/j.cnki.jasp.2016.12.010
Abstract PDF(1214KB) icon1052 icon658
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The experimental model of two-dimensional wedge nozzle with cooling structure was designed, and a series of experimental investigations were conducted to measure the effect of wedge film cooling on infrared signature of the nozzle. The wedge wall temperature, the jet flow temperature on the nozzle exit and the infrared radiation intensity of the nozzle were measured at blowing ratios of 0 to 1. Results show that,with the increase of blowing ratio, the wedge wall temperature gradually decreases and the film cooling effectiveness increases up to 0.68; the jet flow temperature on center area of the nozzle exit decreases to 14%, the infrared radiation intensity of the nozzle gradually decreases to 52.8% at azimuth angle of 0° and 13% at azimuth angle of 90°.
Experimental and numerical study of Y-jet nozzle of LN2-spray temperature-adjusting system
WANG Ze, XING Yu-ming, LIU Xin, ZHAO Liang
2016, 31(12): 2904-2912. doi: 10.13224/j.cnki.jasp.2016.12.011
Abstract PDF(1445KB) icon1222 icon640
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Based on Euler-Lagrange method, the characteristic of LN2 (liquid nitrogen) atomized by the Y-jet nozzle was studied by discrete phase model and air blast atomizer model of software FLUENT. Because of the evaporation effect of LN2 droplets, the spray angle was about 10°, significant smaller than the spray angle of water, and the spray distance was shortened. The effect of operating pressure on the Sauter mean diameter (SMD) and the distribution of volume fraction and quantity percentage was analyzed. The results indicate that the SMD along the spray direction is essentially unchanged. The SMD is controlled by gas velocity and gas-liquid ratio, and gas velocity is the major factor in the low pressure conditions while gas-liquid ratio is the major factor in the high pressure conditions. The distribution of LN2 droplets diameter is more concentrated than in the cases using water, due to the evaporation of droplets. The effect of temperature control of LN2-spray temperature-adjusting system using Y-jet nozzles with different flow rate demands was also investigated. The temperature of outlet was lower than 98K with temperature difference of 12K between inlet and outlet, and the temperature difference of outlet was within ±1K, which could satisfy the temperature control requirements.
V2F simulation on secondary flow and heat transfer of turbine cascade endwall
TIAN Xing-jiang, CHANG Hai-ping, ZHANG Jing-yang, CHENG Feng-na
2016, 31(12): 2913-2920. doi: 10.13224/j.cnki.jasp.2016.12.012
Abstract PDF(1261KB) icon1662 icon630
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Numerical simulation using V2F model based on a linear eddy viscosity assumption, was carried out on subsonic flow and heat transfer of turbine cascade endwall with low aspect ratio airfoil. The results show that complex vortex structures in turbine cascade are mainly composed of horseshoe vortex, passage vortex, pressure side corner vortex and suction side corner vortex. The horseshoe vortex and passage vortex are main sources of secondary loss, and their intensity and location have direct effect on heat transfer of endwall, showing obvious partition property; prediction on endwall limiting streamlines shows that single horseshoe vortex separation line with V2F model is consistent with the experiment data, and superior to double horseshoe vortex separation line predicted by SST(shear stress transport) k-ω model. With introduction of a new turbulent scale, which coincides well with experiment data on endwall static pressure loss coefficient distribution, locations of horseshoe vortex and passage vortex and total pressure loss coefficient distribution of cascade outlet and endwall Standon number distribution, V2F model could predict better on both cascade aerodynamic loss and endwall heat transfer than SST k-ω model.
Preliminary study on combustion performance of dual-mode scramjet engine
TIAN Ye, XIAO Bao-guo, ZHANG Shun-ping, XING Jian-wen
2016, 31(12): 2921-2927. doi: 10.13224/j.cnki.jasp.2016.12.013
Abstract PDF(937KB) icon1245 icon565
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The effect of equivalence ratio and fuel distribution on combustion performance of the dual-mode scramjet engine was investigated by numerical simulation and experiments. Results showed that, under the inflow condition with Mach number of 2.0, total temperature of 1100K, and total pressure of 1.0MPa. When the total equivalence ratio was 0.6, the combustion mode was dual-mode subsonic combustion, the thermal throat was at the exit of the cavity. When the total equivalence ratio was larger than 0.6, the combustion mode was subsonic combustion, the thermal throat was at the same location in different cases. The flow structure of this dual-mode scramjet was stable under subsonic combustion mode. The fast thrust analysis method is created by choosing pressure reference points is developed, the error of this method is rather small.
Finite element simulation for effect of loading on HCF life scatter
YANG Xiao-guang, MIAO Guo-lei, HAN Shi-wei, SHI Duo-qi
2016, 31(12): 2928-2932. doi: 10.13224/j.cnki.jasp.2016.12.014
Abstract PDF(910KB) icon1072 icon568
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The finite element simulation method of high cycle fatigue(HCF) life scatter based on crystal plasticity theory was introduced. According to the features of microstructure of typical titanium alloy TC4, the grains model was established by Voronoi method. Combining the crystal plasticity constitutive equations of titanium alloy, the simulation of the effect of stress level on the fatigue life scatter was performed. The parameter of fatigue life dispersion was given. Results show that,the fatigue life is longer and the life is more dispersed as the loading decreases.
Optimization design for laminate scheme of fiber reinforced composite shaft
SHA Yun-dong, JIA Que-yue, LUO Li, HAO Yan-ping, LI Shou-qiu, ZHAO Feng-tong
2016, 31(12): 2933-2940. doi: 10.13224/j.cnki.jasp.2016.12.015
Abstract PDF(984KB) icon1055 icon576
Abstract:
Based on micro mechanics finite element method, the improved micro mechanics model of the representative volume element (RVE) was used to predict effective mechanical properties of continuous fiber reinforced metal matrix composite. Then square diagonally arranged RVE model was used to calculate its mechanical performance, which was better than that of quadrilateral RVE model. Taking continuous fiber reinforced metal matrix composite shaft as research objects, a micro and macro mechanics model of continuous fiber reinforced metal matrix composite shaft was built. The carrying capacity of this shaft, the rule of shaft carrying capacity, critical buckling load, and critical speed under the influence on different arrays between fiber and matrix were calculated. On this basis, in order to achieve optimization design about laminate scheme of continuous fiber reinforced metal matrix composite shaft, total thickness was taken as objective function with design condition of an aero-engine, and laminate scheme from micro RVE to macro shaft was obtained by random design method.Result shows that, use square diagonally arranged RVE model to calculate the mechanical performance is better than that of quadrilateral RVE model. Fiber and matrix is square diagonally arranged shaft carrying capacity, the critical buckling load and critical speed can be improved. The scheme is the same as GE(General Electric) Company SiC/Ti low pressure turbine shaft layer solution.
Method of loading calibration for actual measurement ofaero-engine thrust based on installation structure
HUANG Yi-feng, LIU Ke-ge, YAN Chu-liang, HAN Xiao-jin
2016, 31(12): 2941-2948. doi: 10.13224/j.cnki.jasp.2016.12.016
Abstract PDF(963KB) icon1225 icon531
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For a certain aero-engine, strain gauges layout and bridges design on the main load-bearing components of areo-engine installation structure were carried out. To establish the relationship equations between strain output and load input, load calibration tests were conducted on the installation structure components and the whole structure respectively. And the correlation and significance for calibration equation of the thrust was analyzed in this approach. The results of correlation coefficient and F-measure reached more than 0.999 and 22000 respectively. Comparing the thrust calculations and load values, the errors under one-direction and two-direction were less than 2%. Results show that calibration equation has high accuracy. With introduction of the flight data into calibration equations, the engine thrust-time history curve was obtained. By contrast, the developmental trends of engine thrust-time history curves and engine fuel flow-time history curves coincided well. The results verify effectiveness and feasibility of the engine thrust test method further, and provide an enforceable technical approach for engine thrust measurement in actual flight.
Influence of tenon structure contact surface geometry on contact stress
ZHANG Ye, SHEN Xiu-li
2016, 31(12): 2949-2956. doi: 10.13224/j.cnki.jasp.2016.12.017
Abstract PDF(900KB) icon987 icon569
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Targeting the contact boundary calculation accuracy problems, grid partition strategy suitable for the tenon structure contact area was proposed in combination with the photo-elastic experiment. On this basis, further analysis of the influence of tooth structure and tooth profile angle on the contact stress was taken, and applied to verify the trapezoidal tooth structure. Result shows that, the trapezoidal tooth of reasonable structure can effectively reduce the contact stress, with less influence on other dangerous parts, so the reasonably designed trapezoidal tooth has a certain advantage compared with the circular arc tooth. After the optimization which takes the trapezoidal tooth contact angle as design variable, the maximum equivalent stress of contact boundary drops 25.54%.
Compliance verification method for pressurized engine static parts of aircraft engine airworthiness
WANG Tao, HU Dian-yin, WANG Rong-qiao
2016, 31(12): 2957-2964. doi: 10.13224/j.cnki.jasp.2016.12.018
Abstract PDF(861KB) icon1423 icon568
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The newly added airworthiness terms of CCAR33.64 (pressurized engine static parts) in "aircraft engine airworthiness regulations"(CCAR33-R2) was studied and analyzed. Experiment design method and numerical simulation method were utilized in the process to do research on pressurized engine static parts with compliance verification method and verification flow. Compliance verification process for CCAR33.64 was founded and the example of a certain type of low pressure turbine back rear casing for aircraft engine was performed. The maximum working pressure and 1.1 times maximum working pressure were taken in numerical simulation to analysis. Maximum equivalent stress were 453MPa and 534MPa. Maximum deformation were 0.366mm and 0.432mm. No permanent deformation or rupture of casing proves its compliance which provides reference for establishing airworthiness guide for pressurized engine static parts of aircraft engines.
Primary resonance response analysis on folding rudder with gaps
LIU Fang, DING Qian
2016, 31(12): 2965-2971. doi: 10.13224/j.cnki.jasp.2016.12.019
Abstract PDF(1074KB) icon1120 icon557
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Primary resonant response of a folding rudder with gaps was investigated. The vibration equations of a folding rudder with gaps both in up-down and pitch directions were established. Harmonic external excitation forces with same frequency were considered in these two directions of motion. The primary resonant periodic solution of the system was analyzed by the average method. Combining with numerical solution, the amplitude-frequency response curves were presented, and the influence of gap, external excitation force, preload on primary resonance of rudder were discussed. Influences of gap, external excitation force and preload on the primary resonance of folding rudder were investigated. The results show that, as the gap increases, the overall main resonance response curve shifts leftwards; non-linear characteristic is more prominent, amplitude jump and hysteresis amplitude appear. With the increase of external excitation force, the proportion of response in the gap range is reducing in line with the nonlinear effects, the vibration characteristic of the system is nearly linear; with increase of fixed preload, the resonance shape changes greatly.
Coupled heat transfer characterisitcs of stepped cooling channel of liquid oxygen/methane rocket engine
SUN Bing, SONG Jia-wen
2016, 31(12): 2972-2978. doi: 10.13224/j.cnki.jasp.2016.12.020
Abstract PDF(1013KB) icon992 icon581
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To study the flow and heat transfer characteristics of transcritical methane in regenerative cooling channel of liquid oxygen/methane rocket engine, and acquire the influence of large-scale sudden expansion and sudden contraction of cooling channel on cooling effect, the three-dimensional coupled heat transfer of thrust chamber was simulated with the whole field direct coupling approach, and the flow of hot gas with non-equilibrium chemical reactions was considered in simulation. Through computation, the three-dimensional temperature field and flow field of thrust chamber were obtained. The computational results show that the maximum value of hot gas side wall temperature appears upstream the throat, because powerful secondary flow exists near the throat section. Due to strong vortices in sudden expansion and sudden contraction sections, the turbulence intensity of coolant is enhanced, the heat transfer coefficient of coolant side wall is significantly increased, the local minimum value of hot gas side wall temperature and large local loss are generated. Most of heat from combustor is absorbed by coolant on the bottom surface and side surface inside the cooling channel, because thermal resistance of copper liner is much smaller than that of nickel jacket. The axial variation of temperature and heat flux on bottom surface is more violent than that on top surface.
Design and implementation of PXI based test measurement and control system for attitude and orbit control liquid rocket engine
WEI Qiang, GUO Hong-jie, LIU Lie, LIANG Guo-zhu
2016, 31(12): 2979-2987. doi: 10.13224/j.cnki.jasp.2016.12.021
Abstract PDF(1210KB) icon1118 icon647
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According to high accuracy, reliability and flexibility of attitude and orbit control liquid rocket engine ground test, a PXI (PCI extensions for instrumentation) controller based test measurement and control system was designed. The control system could operate 40 valves in manual, sequential and automatic modes. The measurement system could acquire 120 signals synchronously with fault diagnosis. The software was developed with LabVIEW in an ideology of modularization, providing favorable expansibility. A process-oriented panel was designed, and embedded control was integrated into the system to improve the reliability. Besides, the safety of fire test was guaranteed with emergency shutdown. To improve the accuracy, calibration was introduced together with a new impulse damping circuit. The test measurement and control system had been successfully implemented in many ground tests of attitude and orbit control liquid rocket engine. The design methods were proved in efficient enhancement of the expansibility and reliability. The accuracy of measurement and control was 0.5% and 0.1ms, respectively, fully fulfilling the ground test requirement of various attitude and orbit control liquid rocket engines.
Flow field in and near the open-end swirl injector
CHEN Chen, YANG Yang, GAO Hong-li, YAN Zhi-hui, YANG Shun-hua
2016, 31(12): 2988-2995. doi: 10.13224/j.cnki.jasp.2016.12.022
Abstract PDF(1070KB) icon1082 icon553
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A numerical simulation using the VOF (volume of fluid) interface tracking method and RNG(renormalization group)turbulence model was carried out in order to deeply understand the atomization mechanism of the open-end swirl injector. The feasibility of the numerical calculation was verified by using PDPA(phase Doppler particle analyzer) and camera to measure the injector exit velocity and the spray angle. The central air core and the recirculation regions in the injector were captured. And the interaction between the tangential inlets and swirling chamber was revealed. What's more, the evolving characteristic of the liquid film thickness and the development law of the internal and external velocity field were emphatically analyzed. With the increasing of the pressure drop, the breaking length decreases while the opening length and the spray angel increase. The unique mechanism of the open-end swirl injector is revealed so it can be helpful to predict of the atomization performance and the structure optimization.
Thrust vectoring features about fluidic throat for solid rocket motor
LI Bo, YU Xin-yu, XIE Kan, GUO Chang-chao, WEI Zhi-jun, WANG Ning-fei
2016, 31(12): 2996-3003. doi: 10.13224/j.cnki.jasp.2016.12.023
Abstract PDF(1238KB) icon1277 icon563
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A cold-flow experiment and numerical simulation research were performed on the fluidic throat using water as working fluid. The general rules of the fluidic throat for solid rocket motor were studied, including the injection styles choke performance, the thrust angle and the thrust performance of the fluidic throat under different conditions of secondary flow injection styles and secondary flow ratio. Numerical simulations were in good agreement with experimental results. Results show that, the choke performance is related to the point, angle and flow ratio of secondary injection. Meanwhile, the choke performance will be better with the increase of the flow ratio of the secondary flow and main flow. Secondary injection of the nozzle throat can effectively regulate the effective throat area which contributes to the thrust magnitude. When the flow ratio is 0.4, the maximum effective throat area ratio is 0.8. Secondary injection at the divergent section can regulate the thrust direction effectively. When the flow ratio is 0.4, the maximum thrust angle is 20°. The phase difference of the throat injection and divergent section injection is used for reducing the interference from throat and divergent segment.
Experiment of breakup characteristics on conical sheets of pintle injectors
FANG Xin-xin, SHEN Chi-bing, ZHANG Xin-qiao
2016, 31(12): 3004-3009. doi: 10.13224/j.cnki.jasp.2016.12.024
Abstract PDF(976KB) icon1117 icon558
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The fragmenting images of surface wave of a pintle injector in different injection pressure drops and structural parameters were obtained using high-speed photography, and the breakup length and breakup time of the conical sheets were measured. Moreover, the variations of the breakup length and breakup time in different operating conditions of the pintle injector were analyzed. According to the experimental results, the mass flow rate of the pintle injector can achieve linear variation in the condition of constant injection pressure drop. When designing a pintle injector, the slit width should be chosen as small as possible within the reasonable range of momentum ratio of propellants. The conical sheets breakup more quickly in the low working condition. In addition, both of the breakup length and breakup time of the conical sheets reduce along with increase of injection pressure drop.
Analysis on shockwave and energy loss in transonic centrifugal impeller with splitter blades
LIU Zheng-xian, CAO Jiang-tao, WANG Mu-su, ZHAO Yi-jia
2016, 31(12): 3010-3016. doi: 10.13224/j.cnki.jasp.2016.12.025
Abstract PDF(938KB) icon1113 icon537
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Based on typical transonic centrifugal impeller with splitter blades, the numerical simulation results were analyzed and judged by the basic formula of shockwave. The existence of shockwaves in the flow passage of impeller was proved theoretically. Result shows that, the high-speed air flow easily formed shockwaves at the leading edges of the splitter blade and the main blade, and also at the flow channel between the suction surface of main blades and the pressure surface of splitter blades separately. The distribution characteristics of shockwave lines affected by the tip clearance and the working condition were compared, furthermore the energy loss of shockwave was analyzed. Result shows that, the position of shockwave lines move upstream with the decreasing flow rate at a constant tip clearance, the tip clearance makes the shockwave lines move to the downstream flow passage, and the strongest shockwave is generated at the leading edge of the splitter blades, leading to obvious energy loss.
A method of optimal design of highly-loaded airfoil
ZHAO Qing-wei, PAN Ruo-chi
2016, 31(12): 3017-3025. doi: 10.13224/j.cnki.jasp.2016.12.026
Abstract PDF(1230KB) icon1156 icon592
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3-section of cubic Bezier spline was used to fit the tangent of camber line and 2-section of cubic Bezier spline was used to fit the thickness distribution. The airfoil was built on streamline-surface. S1 flow field was evaluated by MISES software and optimization algorithm was performed by genetic algorithm(GA). By introducing a parallel mode to the traditional GA, computational efficiency was improved considerably. Automatic design and optimization of highly-loaded airfoil was achieved and the design criteria of highly-loaded airfoil for design of supersonic airfoil were investigated. Results show that compared with the original airfoil, incident angle range of airfoil I is broadened about 2 degree under the premise that loss coefficient of total pressure reduces a little and the airflow angle of outlet is unchanged; if loss coefficient of total pressure of airfoil II obviously reduces, incident angle range is broadened about 0.6 degree and deviation angle reduces about 3.9 degree.
Assembled method for estimation of turbofan engine high pressure turbine inlet temperature
ZHAO Yong, LI Ben-wei, SONG Han-qiang, SUN Tao
2016, 31(12): 3026-3033. doi: 10.13224/j.cnki.jasp.2016.12.027
Abstract PDF(919KB) icon1216 icon551
Abstract:
Considering that direct and reliable measurement of high pressure turbine inlet temperature on board is not feasible at present, a new assembled method for estimation of turbofan engine high pressure turbine inlet temperature was proposed. Based on the energy conservation principles of a turbofan engine, thermodynamic equations between high pressure turbine inlet temperature and other gas path parameters were presented, thus six estimation models of high pressure turbine inlet temperature were derived. Unmeasured parameters of these models were defined as temperature model coefficients as a whole, which were fitted with measured engine power setting parameters using an engine performance model. Validation results show that assembled estimation models have good precision for both health and deteriorative engines, and the mean squared error of its best performance model is not above 1%. Moreover, by comparison with linear and neural network fitting methods, the assembled estimation method has a clear advantage in precision and stability.
Method of generating airfoil manufacturing coordinates based on cold airfoil
CAO Chuan-jun, LI Bin, LIU Xian-zhi, FENG Jin-zhang
2016, 31(12): 3034-3041. doi: 10.13224/j.cnki.jasp.2016.12.028
Abstract PDF(1179KB) icon950 icon488
Abstract:
Method of generating airfoil manufacturing coordinates based on cold airfoil was proposed. It comprised of ordering, interpolating, adding location bias, removing endwall fillet, and fitting endwall airfoil. This method was used to generate multistage axial compressor blade manufacturing coordinates. The difference between manufacturing coordinates and the cold airfoil below 0.01mm except near endwall section was found. Both the waviness of blade surface and curvature of leading or trailing edges were acceptable. Numerical results show the difference between manufacturing coordinate airfoil and cold airfoil has less impact on aerodynamic performance. Mass flow versus total pressure ratio and isentropic efficiency characteristics change little. Besides, flow velocity and flow angle of blade row inlet and outlet, and the detailed flow field have almost no changes. This method can be applied for generating airfoil manufacturing coordinates.
Partition method for improvement of piecewise linear model with full envelope covered
YANG Shu-bo, WANG Xi, LONG Yi-fu, LI Zhi-peng, HU Zhong-zhi, YIN Kai, ZHANG Rong
2016, 31(12): 3042-3053. doi: 10.13224/j.cnki.jasp.2016.12.029
Abstract PDF(1454KB) icon951 icon498
Abstract:
Since ambient conditions vary in a wide range within the full flight envelope, the existing piecewise linear model (PLM), which is based on sea-level static condition with use of corrected parameters for other points in the flight envelope, cannot meet the accuracy for replacing the nonlinear model. To obtain more accurate linear models, a method of partitioning the flight envelope over a grid of Mach number and altitude boxes was suggested. Then, a set of linear models for a given operating condition was selected by picking the nearest (Mach number, altitude) box in the flight envelope. Through the selected set of linear models, interpolating for power level based on a weighted sum of corrected rotor speeds can obtain a linear model with acceptable accuracy. Simulation results of different points within the full flight envelope showed that the maximum error between nonlinear model and the existing PLM was more than 50%, while the maximum error between nonlinear model and the improved PLM was within 8%. It is concluded that the improved PLM performs accurately, especially under the non-standard conditions. In addition, the improved PLM can satisfy the real-time requirement better than the existing PLM.
Experiment on electrostatic sensors airborne monitoring of a turbofan aero-engine
YIN Yi-bing, ZUO Hong-fu, FU Yu, LIU Chen-ning, MAO Hui-jie
2016, 31(12): 3054-3063. doi: 10.13224/j.cnki.jasp.2016.12.030
Abstract PDF(1157KB) icon1254 icon562
Abstract:
A kind of electrostatic sensor with specific size was designed and assembled on exhaust nozzle in wall of turbofan engine nearby the low pressure turbine, and the civil turbofan engine was served as electrostatic sensor carrier and monitoring object in exhaust electrostatic monitoring experiment. The whole performance test process was taken as test period, and exhaust electrostatic signal was fully acquired during the test period. Data processing and contrastive analysis of electrostatic signal data and aero-engine performance parameters were taken, meanwhile the activity level(AL) and negative event rate (NER) of abnormal signal in the test period were calculated, and reason of the abnormity was analyzed. This experiment verified the feasibility of airborne installing for electrostatic monitoring technology. Results show that,the electrostatic signal is about 5mV in normal condition, electrostatic level changes with the operating power and has an obvious positive correlation trend; electrostatic sensor has great effects on monitoring abnormal particle in exhaust, the AL and NER can reach 14pC and 8% in fuel-rich combustion condition.
Performance analysis of air foil thrust bearings withdifferent top foil taper heights
XU Fang-cheng, ZHANG Guang-hui, SUN Yi, LIU Zhan-sheng, ZHANG Wen
2016, 31(12): 3064-3072. doi: 10.13224/j.cnki.jasp.2016.12.031
Abstract PDF(1374KB) icon1011 icon506
Abstract:
The foil structural 2-D thin plate model of bump type air foil thrust bearing was created, finite difference method (FDM) and finite element method (FEM) were coupled together to solve the compressible air Reynolds equation and film thickness equation. The bearing performance such as pressure distribution, film thickness, top foil deformation and power loss were obtained in specific bearing load condition, and the effects of taper height on bearing performance were studied by comparing simulation results with 25, 70, 100, 200, 300μm taper heights. The results show that reducing taper height can make pressure distribution more uniform and reduce the concentrated force on top foil. However, with an optimum value, the bearing can maximize the minimum film thickness for the required same load capacity. Also the bearing can realize the lowest power loss, and the bearing efficiency is improved. These results can be treated as a guideline of air foil thrust bearing design.

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