2016 Vol. 31, No. 8

  Previous Issue | Next Issue 
Display Method:
Design and performance test of a kind of three-blade propeller
XIANG Song, TONG Gang, WU Jiang, ZHAO Wei-ping, YANG Kang, WANG Ji
2016, 31(8): 1793-1798. doi: 10.13224/j.cnki.jasp.2016.08.001
Abstract PDF(993KB) icon1649 icon862
Abstract:
According to the flight velocity, the required thrust and rotation speed of propeller, the three-blade propeller of an aircraft was designed, and the chord length and pitch angle distribution along the radial direction of propeller were obtained. The three-blade composite propeller (diameter of 1.75m) was manufactured based on the designed results, and the thrust of this propeller under various rotation speeds was measured using a propeller test rig. In order to obtain the dynamic performance of propeller, the scaled propeller(diameter of 0.96m) was manufactured, while the aerodynamic performance data of thrust, torque, power and efficiency were measured at the three-dimension test section of Northwestern Polytechnical University NF-3 wind tunnel. Results show that the maximum efficiency of propeller is 85.63%, this three-blade propeller is suitable for the 70kW engine (for example: Rotax 912), so this designed method of propeller has good applicable value.
Numerical investigation of geometric parameters' influence to aerodynamic performance of two dimensional convergent and divergent nozzle
SHI Xiao-juan, JI Hong-hu, SI Ren
2016, 31(8): 1799-1810. doi: 10.13224/j.cnki.jasp.2016.08.002
Abstract PDF(1178KB) icon1519 icon768
Abstract:
The computational cases were designed using the orthogonal experiment design method. The effect of eight geometric parameters on the aerodynamic performance of two dimensional convergent and divergent (2D-CD) nozzle was investigated. The geometric parameters studied include length to diameter ratio, cross-sectional area change rate, long half shaft rate, short half shaft rate of the super elliptic transition section, and throat width to height ratio, throat radius ratio, convergence half angle, divergent half angle of primary nozzle. The sensitivity analysis of throat aspect ratio, throat radius ratio, convergent half angle, divergent half angle on aerodynamic performance was performed. The results show that throat radius ratio is the main parameter of 2D-CD nozzle's aerodynamic performance, therefore the throat radius ratio should be increased in the 2D-CD nozzle design, especially when the convergent half angle is larger. Width to height ratio of throat is not a major parameter of 2D-CD nozzle's aerodynamic performance. The aerodynamic performance of 2D-CD nozzle is independent of transition duct.
Numerical simulation of powered high-lift jet effects for large transport
GONG Zhi-bin, LI Jie, JIANG Sheng-ju, ZHANG Heng
2016, 31(8): 1811-1819. doi: 10.13224/j.cnki.jasp.2016.08.003
Abstract PDF(1185KB) icon1319 icon764
Abstract:
Taking C-17 transport engine installation as reference and considering external mixing of engine jet gas streams, a powered high-lift configuration with externally blowing flap (EBF) and the corresponding cruise model were constructed. Based on the structured multi-block grid techniques and Reynolds-averaged Navier-Stokes methods, numerical validations were carried out on a high-lift model and a fan-jet engine nacelle, and then the EBF effects were investigated. For powered high-lift configuration at low speed, most jet flow impinged directly on the lower flap surfaces and then deflected downward, while the rest was sucked and speeded up by the slots, and then blown upon the upper flap surfaces. The two parts of them converged at the flap trailing edge and extended downward. Spanwise lateral flow characteristics existed when the jet flow affected on the flap. Circulations increased significantly not just around the flaps but also on the slat and main wing. Both the available and maximum lift coefficients were larger than 4.0, breaking through the limit of the traditional mechanical high-lift systems. Longitude trim problems may be severer due to larger nose down pitching moment. For the cruise model at high speed, the engine jet flow will obviously affect the pressure distributions on lower wing surface, making the lift decrease and the drag increase significantly. Thus jet effects should be well considered during the clean wing design for powered high-lift configuration.
Aeroelastic analysis of rotor airfoil based on CFD/CSD tight coupling method
HUANG Dao-bo, LI Jian-dong, LIU Yong
2016, 31(8): 1820-1829. doi: 10.13224/j.cnki.jasp.2016.08.004
Abstract PDF(1288KB) icon1207 icon672
Abstract:
By coupling the calculation of the unsteady aeroforce and the vibration response through the development of CFD/CSD (computational fluid dynamics/computational structure dynamics) tight coupling method, the Fluent code was calculated, and the data of structural response and dynamic load were exchanged by UDF (user defined function)and reading and writing the I/O (input/output) files. The aerodynamic characteristics and vibration response of the airfoil at the condition of pitching motion, plunging (flapping) motion and periodic alternating incoming flow were studied. Numerical results indicated that the aerodynamic load was periodic, showing a smoothing hysteresis curve in rotating process,and the flapping responses also showed a cyclical phenomenon, the pitching response showed slight oscillation at slight stall condition. At the deep stall condition, the aerodynamic load or structural response showed strong nonlinear oscillation, high frequency components were more obvious.
Effect of forced-transition on performance of hypersonic inlet
YI Miao-rong, ZHAO Hui-yong, LE Jia-ling
2016, 31(8): 1830-1837. doi: 10.13224/j.cnki.jasp.2016.08.005
Abstract PDF(1408KB) icon1299 icon674
Abstract:
Diamond and ramp forced-transition trips with different heights were designed for a hypersonic inlet. On the condition of freestream Mach number 6 and the angle of attack 1 degree,the effects of forced-transition on wall heat flux distribution, pressure, total pressure and Mach number of isolator inlet were investigated by both wind-tunnel experiment and three-dimensional numerical simulation. The computational data agreed well with wind-tunnel experiment results in wall pressure, inlet total pressure and Mach number. And there was at most 35% difference in wall heat flux between simulation and experiment results. The results show that both the diamond and ramp forced-transition trips can promote forced-transition effectively. The transition location moves forward with the increasing height of the trips. For the started inlet, the forced-transition trips have little effect on the static pressure of the inlet bottom surface, but the total pressure and Mach number near the bottom surface of the isolator will decline; in such case, the mass-averaged total pressure and the mass flow rate in the isolator decline about 3.5%. The decline of the mass-averaged Mach number is only about 1%.
Design of an ornithopter with multisection flexible morphing wings
HUANG Ming-yang, XIAO Tian-hang, ANG Hai-song
2016, 31(8): 1838-1844. doi: 10.13224/j.cnki.jasp.2016.08.006
Abstract PDF(1053KB) icon1969 icon951
Abstract:
The multisection flexible morphing wings mimic the complex movement of seagull's flapping wings. By observing movement periods of seagull's flapping wings, a new flapping-wing model including fold, torsion, flexible morphing and slowly beating up and down was built. In order to provide the design of this ornithopter with a theoretical basis, the quasi-steady method was applied to calculate aerodynamic force. The three-dimensional model and motion simulation were built in CATIA and 3DMAX, and an ornithopter with multisection flexible morphing wing was built for flight test. Flying attitudes including level flight,climbing and yawing were studied, and test results including lift and thrust agree with the calculation. Compared with conventional ornithopters, the ornithopter with multisection flexible morphing wing can fly while flapping with a low frequency, and adjusting the shape of flapping wings.
Experiment for improving aerodynamic performances of a flying wing by millisecond pulsed plasma actuation
MA Jie, LIANG Hua, WU Yun, HAN Meng-hu, WEI Biao, ZHAO Guang-yin
2016, 31(8): 1845-1851. doi: 10.13224/j.cnki.jasp.2016.08.007
Abstract PDF(936KB) icon1118 icon658
Abstract:
The experiment for improving aerodynamic performances of a flying wing by millisecond pulsed dielectric barrier discharge (DBD) plasma actuation was performed in a wind tunnel at incoming speed of 30m/s. Plasma actuator was planted along the leading edge of the flying wings; the discharge energy of each pulse was at the order of 0.1mJ/cm when the peak-to-peak voltage was 9.5kV. The effect of actuation frequency and duty cycle on the aerodynamic performances, such as lift/drag coefficient, lift/drag ratio and pitch moment coefficient, was investigated through force measurement by a six-component balance. The results show that the aerodynamic performances of the flying wing at high angles of attack can be improved with plasma actuation. When the optimum non-dimensional actuation frequency F+≈1, the stall angle of attack increases from 14° to 17° and the maximum lift coefficient increases by 10%. The duty cycle plays an important role in flow control efficiency. The best duty cycle is 5% in the current investigation, which can reduce the power consumption tremendously. Meanwhile, the changes of pitch moment coefficient indicate that longitudinal static stability of flying wing is improved by plasma actuation.
Experiment and numerical computation of impingement cooling in narrow space with pin fins
RAO Yu, WAN Chao-yi, CHEN Peng
2016, 31(8): 1852-1859. doi: 10.13224/j.cnki.jasp.2016.08.008
Abstract PDF(1127KB) icon1212 icon733
Abstract:
Experimental and numerical computation were conducted on impingement cooling in a narrow space with full-height pin fins turbulated flow, and the performance comparisons were made with another impingement cooling on the flat target plate under maximum cross flow scheme. The jet Reynolds number ranges from 15000 to 30000. Transient liquid crystal thermography method was used in the experiments to obtain the detailed impingement heat transfer distribution on the endwall of the target plates, and numerical computation was done to examine the flow structure and heat transfer characteristics in the impingement cooling systems. Experimental study shows that the pin-fin can slightly improve the average heat transfer characteristics on the endwall of the target plate by about 7.0%, and increase the pressure loss by about 17.9%. It is also found that heat transfer uniformity is also improved on the endwall of the target plate with pin fin. On the other hand, numerical computations still show that the wall jets and the upwashing vortex in the space strongly interact with the pin fins, and the pin fins appreciably increase the heat transfer area, therefore the impingement cooling system with pin fins has significantly improved the total heat transfer performance by about 27.0% compared with the impingement with the flat plate.
Influence of indirect initiation on the thermal efficiency of pulsed detonation engine
QIU Hua, WANG Wei, XIONG Cha
2016, 31(8): 1860-1865. doi: 10.13224/j.cnki.jasp.2016.08.009
Abstract PDF(861KB) icon1079 icon638
Abstract:
To analyze the influence of indirect initiation, a method was first established to evaluate the average cycle thermal efficiency of the system whose working fluid experienced different combustion processes in combustion chamber. Then, the thermal efficiency and deflagration to detonation transition (DDT) characteristics of two detonation tube models with different obstacles were investigated by the 2-D numerical simulations. The ratio of combustion wave run-up distance to DDT distance (defined as ξ) was also studied. The results show that: (1) the thermal efficiency of the pulsed detonation engine (PDE) model can be both influenced by the indirect initiation and different obstacles; (2) when ξ approaches to 1, the average thermal efficiency of the system is only 60% of ideal detonation cycle; (3) the effects of indirect initiation are weakened with the increase of ξ; (4) when ξ approaches to 1.5, the average thermal efficiency will reach 93% of ideal detonation cycle.
Pressure loss characteristics in a rotating cavity with de-swirled nozzles
HU Yan-li, GUO Wen, WANG Lei, LUO Xiang
2016, 31(8): 1866-1873. doi: 10.13224/j.cnki.jasp.2016.08.010
Abstract PDF(1098KB) icon1331 icon692
Abstract:
De-swirled nozzles were fitted in a rotating cavity to reduce the inlet flow tangential velocity thus to reduce the pressure loss associated with radial inflow in the cavity. In this work, experiments were conducted to measure the pressure difference through the test rig and k-ω SST model was applied to simulate the flow structures. Experimental results revealed that the proportion of pressure loss through nozzles to the overall pressure loss through test rig exists maximum and non-maximum conditions. While the proportion of pressure loss through nozzles to the overall pressure loss reaches the top there are no large scale vortex. In the other conditions there are always two large scale vortexes in the cavity, which lead to the largest pressure loss in the rotating cavity.
Studies on combustion and emission characteristics of coal-based jet fuel in ground individual can combustor
YANG Wen, ZHANG Dong-mei, CAO Wen-jie, WU Xiu-zhang, WANG Shu-lei
2016, 31(8): 1874-1882. doi: 10.13224/j.cnki.jasp.2016.08.011
Abstract PDF(947KB) icon1140 icon687
Abstract:
By the improved individual can combustor, the differences of the combustion and emission characteristics between coal-based jet fuel and petroleum-based RP-3 jet fuel were evaluated. The results showed that the coal-based jet fuel had better lean ignition and lean blowout performances; in addition, the combustion and emission characteristics of these two jet fuels exhibited the same trends when the fuel flow or exit temperature changed. However, the combustion and emission characteristics of the coal-based jet fuel were slightly lower than those of the RP-3 at the same fuel flow, furthermore, at the same exit temperature, the wall temperature of the coal-based jet fuel was lower than that of the RP-3.
Effect of carbon deposit removing hole area on the outlet flow field of double-swirler
ZHENG Jian-wen, SONG Shuang-wen, PENG Jian-yong, LIU Tao
2016, 31(8): 1883-1888. doi: 10.13224/j.cnki.jasp.2016.08.012
Abstract PDF(1125KB) icon1071 icon681
Abstract:
Particle image velocimetry (PIV) was used to study the effect of atomizer shroud carbon deposit removing hole area on the outlet flow field of double-swirler. The result showed that, when the ratio of the carbon deposit removing hole area to the double-swirler area increased from 0 to 0.24, the recirculated air velocity decreased and less air was recirculated into the double-swirler. When the ratio of the carbon deposit removing hole area to the double-swirler area was more than 0.18, the carbon deposite removing air would destory the sysmmetrical double-vortex structure and disturb the outlet flow field, which with X/D=0-0.8 was critically influenced by the carbon deposit removing air.
Numerical investigation of convective heat transfer on a wedge-shaped concave surface subjected to staggered offset impinging jets
GUAN Tao, ZHANG Jing-zhou, SHAN Yong
2016, 31(8): 1889-1896. doi: 10.13224/j.cnki.jasp.2016.08.013
Abstract PDF(1199KB) icon1118 icon593
Abstract:
Three-dimensional numerical simulations were conducted to investigate the convective heat transfer on a wedge-shaped concave surface subjected to staggered offset impinging jets. Under the presented parameters, such as offset spacing ratio (L/d) of 0-2.5, jet to concave leading edge spacing ratio (H/d) of 6-12, and jet Reynolds number (Re) of 10000-28000, the results showed that staggered offset of the impinging jets introduced complicated vortex flow inside the concave cavity and enhanced the local heat transfer correspondingly to the impingement stagnation zone. By comparison with the jets arranged in a line, relative little staggered offset of the impinging jets enhanced the laterally-averaged Nusselt number in the vicinity of concave leading edge. With the increase of offset spacing ratio, the chordwise location corresponding to the maximum laterally-averaged Nusselt number moved downwards. To improve the laterally-averaged Nusselt number in the vicinity of concave leading edge while maintaining local heat transfer capacity at the leading edge, the offset spacing ratio of 1 approximately is more appropriate.
Reliability model of aero-engine structural system
HONG Jie, ZHANG Zi, XU Xiao-li, MA Yan-hong
2016, 31(8): 1897-1904. doi: 10.13224/j.cnki.jasp.2016.08.014
Abstract PDF(912KB) icon1557 icon639
Abstract:
In order to make the error of aero-engine reliability model smaller, error analysis was made for traditional aero-engine reliability model, the reason that reliability indexes over expand was pointed out, and the aero-engine reliability model based on failure modes in structural system considering aero-engine structural system and the failure mechanics characteristics was established. The reliability calculation method of each layer was given. Structural system reliability model of aero-engine based on failure modes, which considers relative independence, timeliness and time varying of failure rate, can avoid problems such as missing important failure modes and over expansion of reliability index, so it has great engineering application value.
Equivalent modeling and vibration isolation design ofwing-mounted engine structure
XU Fei, HE Er-ming, LI Jing-xu
2016, 31(8): 1905-1912. doi: 10.13224/j.cnki.jasp.2016.08.015
Abstract PDF(1110KB) icon1208 icon665
Abstract:
In order to research the vibration isolation characteristics and optimize isolator design of wing-mounted engine mounting structure, the theoretical analysis model and finite element model of engine mounting joint-pylon-wing structure were built. The modal verification and vibration isolation characteristic of the mounting structure was analyzed using finite element method. The dynamic response and vibration transmission main path under three typical load spectra was analyzed. The influence of isolator parameters and position on the engine mounting system isolation performance was also analyzed based on the vibration transmission path method. The results show that the peak response of vibration load under rotation frequency of low pressure rotor, rotation frequency of high pressure rotor can reduce to 22.03% and 14.65% by the mounting structure; the vibration transmission main path under rotation frequency of low pressure rotor is the engine-former engine mounting joint-pylon-wing while under rotation frequency of high pressure rotor is the engine-latter engine mounting joint-upper connection rod-wing. The vibration isolation rate of mounting system can reach 50.41% by reasonably setting the position of the vibration isolator, and even reach 70.67% when the frequency ratio of isolator is 5 and the damping ratio is 0.25. The installation position and the parameters of isolator must be appropriately selected to acquire optimal vibration isolation effect of the entire engine mounting system.
Thermal-structural coupling characteristics of axial piston pump slipper pair
TANG He-sheng, YIN Yao-bao, LI Jing
2016, 31(8): 1913-1920. doi: 10.13224/j.cnki.jasp.2016.08.016
Abstract PDF(1453KB) icon987 icon602
Abstract:
To improve the wear resistance of axial piston pump slipper pair, the transient thermal-structure coupling model was established based on the slipper/swash plate friction pair. Under the pressure shock condition, the slipper surface temperature, stress and deformation were analyzed. The results show that slipper temperature in a piston pump periodically changes with the piston chamber pressure and the range of slipper temperature from 45.5℃ to 49.8℃. The highest slipper temperature appears in the pump suction-discharge pressure transition zone. When the slipper runs in discharge pressure zone of pump, the maximum axial stress is 250MPa, which is concentrated on the edge region between slipper pocket and sealing belt. The layed axial stress of slipper is significant, causing the slipper deformation differentiation. The range of slipper deformation is from 12.5μm to 15μm, which appears at the edge of slipper. Shear force of abrasive particle is enhanced by input heat flux of slipper intensifying micro cutting and extrusion deformation. Furthermore, strip exfoliation and pits are presented by wear surface of slipper, which shows adhesive and abrasive wear characteristics.
Non-deformability evaluation of aero-engine's case
XU Feng, MA Yan-hong, ZHANG Li, HONG Jie
2016, 31(8): 1921-1928. doi: 10.13224/j.cnki.jasp.2016.08.017
Abstract PDF(1018KB) icon1213 icon661
Abstract:
Based on the structure and dynamic characteristics of aero-engine's case, an evaluation method for the whole engine case's non-deformability under safety design regulations was established. In order to evaluate the non-deformability from three aspects: structure characteristics, deformation characteristics under steady load, and deformation characteristics under impact load, three parameters including structure normalized equivalent stiffness, clearance matching coefficient, and impact amplification coefficient were raised respectively. Then a non-deformability assessment on typical high bypass ratio turbofan engine was conducted. The result shows that using the ratio of equivalent stiffness of each cross section to maximal equivalent stiffness as the structure normalized equivalent stiffness, using the ratio of minimal clearance to maximal clearance of each cross section under steady loads as the clearance matching coefficient, and using the ratio of peak force of impact load to bearing's or mount's reaction force as the impact enlargement coefficient, all the three parameters can reveal the dangerous position of the whole engine's case and the response characteristics under ultimate load. This result can provide a reference for structural design and optimization of other aero-engine's cases.
Component mode synthesis for dynamic reduction of 3-D finite element rotor system and its application
ZUO Yan-fei, WANG Jian-jun
2016, 31(8): 1929-1934. doi: 10.13224/j.cnki.jasp.2016.08.018
Abstract PDF(955KB) icon1087 icon637
Abstract:
Rotating substructure method based on component mode synthesis was proposed to reduce 3-D finite element rotor model. Planar (undamped) mode matrix was used to truncate the degrees of freedom of the substructure and the reduced gyroscopic matrix was got by multiplying a speed coefficient to the unit reduced ones. Compared with complex (damped) mode reduction method, it's no need to recalculate the truncation matrix, and it's more accurate than Guyan reduction method. 87% degrees of freedom of an aero-engine rotor model was reduced. The maximum critical speed error got from Campbell diagram is 0.04%. Unbalance response calculated by the reduced model is almost the same as the original's, and less than 20% RAM (random access memory) and computation time is used. All the results show that the proposed method is valid.
Experiment and numerical simulation of external performances and internal flow of a molten salt pump
SHAO Chun-lei, GU Bo-qin, ZHOU Jian-feng, CHENG Wen-jie
2016, 31(8): 1935-1942. doi: 10.13224/j.cnki.jasp.2016.08.019
Abstract PDF(1105KB) icon1199 icon584
Abstract:
In order to reveal the regularity of two-phase flow in the molten salt pump, an experimental method of a molten salt pump was researched, and the external performances were obtained. The internal flow field near the volute tongue was measured by using PIV (particle image velocimetry) technology. Unsteady numerical simulation of two-phase flow in the molten salt pump was carried out. The differences between experiment and numerical simulation results were elaborated. The distribution of volume fraction of crystal particles was obtained and its evolutional regularity was analyzed. The results show that the distribution of volume fraction of crystal particle is stable in the molten salt pump. In the impeller, particles are accumulated near the suction side of blades. The particle volume fraction is affected significantly by the particle diameter. The larger particle size means more obvious particle accumulation phenomenon. The molten salt pump outlet's static pressure fluctuates periodically, and the amplitude of pressure fluctuation is affected by the particle volume fraction distribution in the impeller passage.
Study on improvement of the compressor anti-swirl distortion capacity with bowed stator
TU Bao-feng, LIU Hua, HU Jun
2016, 31(8): 1943-1949. doi: 10.13224/j.cnki.jasp.2016.08.020
Abstract PDF(1358KB) icon1043 icon647
Abstract:
In order to reduce the effect of swirl distortion on performance and stability of axial compressor, the first-stage stator of two-stage low speed axial compressor was designed into bowed blade with three different positive bending angles. A full annulus three-dimensional steady simulation method was used to investigate the effect of twin swirl on performance and stability of compressor. The results show that stator with appropriate positive bending design can improve the anti-twin swirl capacity of compressor. By reducing the load of blade at end wall regions using bowed stator blades, the suction surface boundary layer separation at hub will be decreased, loss will be reduced, and both of efficiency and total pressure ratio will be increased as a result. The maximum efficiency can be increased by 1.5% and mass flow rate at stall boundary point can be decreased by 6.95% when bending angles of stator blades equal to 10 degree. The efficiency and total pressure ratio will be increased by 1.7% and 0.03% respectively when mass flow rate equals to 18.5kg/s. Load at mid-span region and friction loss will increase when bending angle is too large, making it harmful to improve the efficiency and total pressure ratio.
Effect of tip clearance on the operating mechanism of offset splitter blades in a centrifugal compressor
CHEN Jie, HE Min-xiang, HUANG Guo-ping
2016, 31(8): 1950-1956. doi: 10.13224/j.cnki.jasp.2016.08.021
Abstract PDF(1066KB) icon1128 icon651
Abstract:
The good circumferential position of offset splitter blades and its mechanism were studied in small centrifugal impellers, which operated with different tip clearances by the computational fluid dynamics. When the impeller worked with the small tip clearance, the main loss was attributable to the separation near the main suction surfaces, and then the offset splitter bias toward the main suction surface can suppress this loss and enhance the impeller performance. The tip leakages increased with the tip clearance, so the loss concentrated in the passage between the main pressure surface and splitter, therefore the splitters tended to the main pressure surface, and then the offset splitters can reduce the concentration of the tip leakage and the loss. The tip clearance and the circumferential position of the splitter affected together the tip leakage, the separation near the suction surface and their interrelation. The design of the splitter circumferential position can adjust the crosswise pressure gradient in the blade passage and the location where the tip leakage mixes in the main flow, and then improve the ratio of the losses in the passages at both sides of splitter, therefore the best circumferential position of the splitter changes with the tip clearance.
Experiment on surge phenomenon of small multi-stage axial compressor
HE Xiang, MA Hong-wei, YIN Yue-qian, SHAN Xiao-ming, YAO Zheng-rong
2016, 31(8): 1957-1963. doi: 10.13224/j.cnki.jasp.2016.08.022
Abstract PDF(1188KB) icon1217 icon612
Abstract:
An experimental investigation on the compressor surge phenomenon was carried out in a small multi-stage axial compressor under different rotor speeds. The dynamic pressure signals of 1st rotor tip were measured by the Kulite dynamic pressure transducers and the fast response pressure probes. Results indicated that, at mid and higher rotor speed, the modal wave appears and evolves into rotating stall in the pre-surge period and, subsequently, the compressor is rapidly brought to deep surge by the huge blockage of the stall cell; However, at lower rotor speed, the classical surge is directly triggered by modal wave with the existence of rotating stall in the surge cycle.
Aero-elastic stability analysis of an axial compressor with splitter rotor
HUANG Zhong-shan, WANG Yan-rong, FU Zhi-zhong
2016, 31(8): 1964-1972. doi: 10.13224/j.cnki.jasp.2016.08.023
Abstract PDF(1350KB) icon1108 icon580
Abstract:
The aero-elasticity of an axial compressor with splitter rotor was investigated by employing energy method and time-domain method. The aerodynamic modal damping ratio of blade and unsteady aerodynamic work distribution on the blade surface were obtained while the frequency domain analysis of blade vibration displacement were made; the conventionally designed rotors and the splitter rotor were also compared. The results show that the structural mistuning and aerodynamic mistuning had significant positive effects on the aero-elastic stability of the system. Aerodynamic mistuning improved the unsteady aerodynamic work distribution on the blade surface and the structural mistuning weakened the effect of interblade phase angles. For the first vibration mode, aerodynamic modal damping ratio (AMDR) of the splitter rotor has the value of 0.000791, and that of conventionally designed rotors have the value of 0.000217.Therefore,the structure of the splitter rotor not only has a better aerodynamic performance, but also enhanced the aero-elastic stability.
Review of research on solid fuel scramjet engine
LV Zhong, XIA Zhi-xun, LIU Bing, LIU Yuan-chun
2016, 31(8): 1973-1984. doi: 10.13224/j.cnki.jasp.2016.08.024
Abstract PDF(1112KB) icon1502 icon687
Abstract:
The configurations of scramjet using solid fuel as propellant were summarized, and the progress and status of different configurations of scramjet using solid fuel propellant were elaborated, respectively. According to the operating feature of solid-fuel scramjet, solid-fuel dual-combustor scramjet and solid-fuel rocket scramjet, the advantages and existing problems of different configurations were analyzed, and some suggestions for further research were proposed. The research results reveal that though solid fuel can ignite and burn in supersonic combustor of solid-fuel scramjet, the combustion efficiency of propellant is low and the scramjet cannot operate for a long time. The solid-fuel rocket scramjet is beneficial to ignition and stable combustion of solid fuel, and can also operate for a long time. The solid-fuel rocket scramjet has a bright research and application prospect.
Influence of combustor shape on self-ignition in solid fuel scramjet
CHI Hong-wei, WEI Zhi-jun, LI Biao, WANG Li-he, WANG Ning-fei
2016, 31(8): 1985-1994. doi: 10.13224/j.cnki.jasp.2016.08.025
Abstract PDF(1111KB) icon1463 icon596
Abstract:
The transient self-ignition mechanism of PMMA (polymethylmethacrylate) and the effect of combustor shape on self-ignition in solid fuel scramjet with cavity were simulated numerically. An unsteady two-dimensional axisymmetric RANS (Reynolds-averaged Navier-Stokes) equation, SST(shear stress transport) k-ω turbulent model, finite-rate/eddy dissipation reaction model were solved numerically. Main conclusions are as follows: the reactants reach the ignition temperature when the heat of reaction accumulates and increases continuously within the residual time provided by the cavity, so the self-ignition is permitted in solid fuel scramjet. The cavity length, cavity diameter, convergence angle and the cylindrical section diameter of the combustor are the main influential factors to the self-ignition. A step type cavity is suggested to enhance the self-ignition performance.
Impact effect of liquid rocket motor exhaust plume on launching platform
CAI Hong-hua, NIE Wan-sheng, HOU Zhi-yong
2016, 31(8): 1995-2001. doi: 10.13224/j.cnki.jasp.2016.08.026
Abstract PDF(1126KB) icon1017 icon542
Abstract:
To study the impact effect of liquid rocket motor exhaust plume on launching platform, a mathematical calculation model of exhaust plume impact on launching platform for liquid rocket motor was built. For the impact effect of LOX (liquid oxygen)/ kerosene exhaust plume on the launching platform, the influences of the propellant flux and afterburning on the impact effect from the nozzle exit to the platform in two different cases of 3m and 5m were studied, the differences and the reasons for the differences were also analyzed. Results show that, with the increase of the propellant flux, the distance of shock wave expansion and compression of the free jet region and the area of the wall-jet region increased. Afterburning can not only change the structure of the free jet region and the stagnation zone, but also increase the temperature and the area of the wall-jet region. Afterburning and propellant flux influence the impact effect by generating influence on the structure of the exhaust plume, and the impact effect is related to the distance from the nozzle exit to launching platform.
On-orbit refilling technologies and schemes of liquid propellant
WANG Lei, LI Yan-zhong, MA Yuan, XIE Fu-shou
2016, 31(8): 2002-2009. doi: 10.13224/j.cnki.jasp.2016.08.027
Abstract PDF(1028KB) icon1246 icon562
Abstract:
The key technologies associated with propellant on-orbit refilling was summarized whilst the system construction and the refilling procedures were introduced. Moreover, a suggesting research project was proposed for guiding the domestic researches. Several valuable conclusions were drawn as follows: (1) On-orbit gas-liquid separation technology is the basis of the propellant refilling. Membrane and vane tank can be used to separate the gas-liquid phases in a normal-temperature propellant tank, while screen channel liquid acquisition devices (LAD) may be the best choice for the cryogenic propellant tank. (2) Space thermal-insulation and boil-off control technologies must be applied in the on-orbit refilling operation of cryogenic propellant. (3) Venting approach can be used for normal-temperature propellant, and no-vent filling is suggested for the cryogenic propellant. Also, use of thermodynamic venting system (TVS) is beneficial to a larger refilling rate of cryogenic propellant. (4) A research scheme concerning the normal-temperature propellant first and then the cryogenic propellant is suggested in the domestic researches. Simultaneously, the present research platform and the latest research achievements should be referenced.
Coupling of dynamics and elastohydrodynamic lubrication for spur gear
ZOU Yu-jing, CHANG De-gong
2016, 31(8): 2010-2020. doi: 10.13224/j.cnki.jasp.2016.08.028
Abstract PDF(1321KB) icon1084 icon609
Abstract:
A six-degree-of-freedom tribo-dynamics model for spur gear pair was proposed based on the load sharing theory and dynamics theory. Time-varying mesh stiffness and film stiffness, surface roughness as well as the friction torque were considered for establishing the model. A decoupling method was used to solve the tribo-dynamics model. In the iterative loop, the tribo-dynamic model predicted the dynamic tooth forces and surface velocities and applied into the lubrication model. In return, the gear lubrication model provided the tribo-dynamic model with friction coefficient and film stiffness. An example of gear pair was analyzed using the proposed model and method to demonstrate the coupling relationship between tribological and dynamic behaviors. The results show that total gear stiffness decreases slightly with consideration of coupling effect.Sliding friction force has significant effect on dynamic response that's perpendicular to action line.Dynamic load will affect oil film thickness, portion of load carried by film, film temperature rise and the friction coefficient greatly. Moreover, the effect of dynamic behavior on film temperature rise depends on the relative sliding velocity.
Studies on the effects of structural parameters on jetcharacteristics of nozzlesin in lubricating aviation gears
ZHANG Rui-qiang, LIU Shao-jun, HU Xiao-zhou
2016, 31(8): 2021-2028. doi: 10.13224/j.cnki.jasp.2016.08.029
Abstract PDF(1208KB) icon1426 icon613
Abstract:
In order to study the effects of the supply pressure and temperature of oil on the diffusion angle and volume flux of jet oil sprayed by nozzles of gear cases, both the two phase and single phase flow models of the jet flow field of three dimensional nozzles were simulated by employing the Large Eddy Simulation based on the FLUENT. The rationality of the simulation results of the two phase and single phase flow models of jet flow field of three dimensional nozzles was demonstrated by the diffusion angle and volume flux of jet oil measured by the method of high speed visualization and weighting jet oil, respectively. The diffusion angle of jet was approximately zero within the scope of 100 millimeters along the axis apart from the exit of nozzle when the ratio of length to diameter was greater than 2 and the angle between the axis of nozzle orifice and the axis of upstream flow channel changed in the range of 0 to 90 degree. The volume flux of jet oil increased as the orifice diameter, the angle between orifice axis and upstream axis increased and as the orifice axial length decreased. The result indicate that the diffusion angle of jet is not affected in actual working conditions when the ratio of length to diameter is greater than 2 and the angle between the axis of nozzle orifice and the axis of upstream flow channel changes in the range of 0 to 90 degree. Both the local loss and the linear loss are influenced by the change of the orifice diameter, the orifice axial length and the angle between orifice axis and upstream axis, finally leading to variation of volume flux of jet.
Research on model-based safety analysis for aero-engine
BAO Meng-yao, LI Guo, DING Shui-ting
2016, 31(8): 2029-2039. doi: 10.13224/j.cnki.jasp.2016.08.030
Abstract PDF(1000KB) icon1272 icon645
Abstract:
By introducing and summarizing the traditional safety analysis method, a model-based safety analysis process and method for aero-engine were proposed. From the property of the aero-engine, unified safety analysis and performance design achieved the goal of balance between them. This could guarantee the consistency, integrity and objective of the analysis results. The three key technical breakthroughs during the safety analysis process and method are:(1) the description method of failure modes; (2) the classification and location methods of the key impact factors; (3) safety analysis verification methods.
Intermediate state control of turbofan engine in full envelope based on switched polytopic LPV approach
WU Bin, HUANG Jin-quan
2016, 31(8): 2040-2048. doi: 10.13224/j.cnki.jasp.2016.08.031
Abstract PDF(1201KB) icon1026 icon549
Abstract:
The parameter range of aeroengine in full flight envelope is so wide that a single controller is difficult to ensure the performance. To avoid this problem, a switched polytopic linear parameter varying (LPV) approach was proposed for the turbofan engine intermediate state control. The full envelope was firstly divided into several locally overlapped sub-regions according to the inlet conditions. The state reset method and the polytopic gain scheduling technique were used to derive the robust stability conditions of this switched LPV system, which was depicted as linear matrix inequality (LMI). Then, the resulting Lyapunov matrices were used to design a family of single LPV controllers which can be applied in the sub-region control combined with the scheduling strategy based on geometric position; Hysteresis switching strategy based on the overlapped sub-regions was used for switching, meanwhile the stability of this closed-loop switching system was also proved. Simulation results with a turbofan engine model show that the steady state error is less than 0.1 percent, and the maximum overshoot is less than 0.5 percent.

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

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