2018 Vol. 33, No. 4

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Study of compression ratio distribution in two-stage series turbocharged aircraft piston engine
2018, 33(4): 769-775. doi: 10.13224/j.cnki.jasp.2018.04.001
Abstract PDF(1032KB) icon1479 icon943
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A calibrated simulation model was created in this research and the accuracy and applicability was verified by experiment. This model was used to match the appropriate turbochargers and explore how the diverse pressure ratio distribution affects the engine performance on the whole operating range considering the efficacy of the intercooler and pressure loss in inlet. Based on the objective of minimizing the energy consumption of turbocharging compressors, series of theories were revealed about how to distribute the pressure ratio between two turbochargers, and a matching method was got to optimize the pressure ratio distribution on a certain two-stage turbocharged aircraft piston engine by virtue of research. Results showed the turbochargers had high efficiency under all working condition and kept 7% away from the dangerous margin via the method. This also provided a foundation for further study on the safety of the two-stage turbocharged aircraft piston engine at service ceiling.
Influence of inlet distortion of different ranges on subsonic axial-flow compressor
2018, 33(4): 776-784. doi: 10.13224/j.cnki.jasp.2018.04.002
Abstract PDF(1623KB) icon1094 icon982
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By means of computational fluid dynamics commercial software NUMECA, the effects of inlet distortion of different range on the performance, stability and flow-field in a subsonic axial-flow compressor were investigated with ten-annulus unsteady numerical method. The research object was the single-stage subsonic axial-flow compressor rotor of Northwestern Polytechnical University. The inlet distortion of different range was generated by setting some barriers on the inlet ring surface. Results showed that inlet distortion didnt only degrade the performance of the compressor rotor, but also decreased the stability of the rotor. And the larger distortion range indicated the more obvious phenomenon mentioned above. The fundamental flow mechanism was obtained by detailed analysis of the flow-field in rotor passage. Inlet distortion led to serious blockage in tip blade passage and further induced rotor stall because of the influence of the tip clearance leakage flow and the separation of boundary layer in some tip blade suction. Like above, the larger distortion range meant the more serious influence.
Open-rotor engine noise analysis based on Mohring acoustic analogy
2018, 33(4): 785-791. doi: 10.13224/j.cnki.jasp.2018.04.003
Abstract PDF(856KB) icon1238 icon962
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In order to systematically study the characteristic of open-rotor noise in non-uniform flow field, CFD (computational fluid dynamics)/CAA (computing aerodynamic acoustic-s) hybrid method of the open-rotor aerodynamic noise was set up. The aerodynamic characteristics of open-rotor were analyzed through compressible modified Spalart-Allmaras turbulence model method, then, the open-rotor noise and noise transmission were numerically simulated by acoustic infinite element method. At the same time, the average flow field and its smooth interpolation were introduced to acoustic grid as a background flow field. Taking into account the noise distribution chart and the far field feature point spectrum curve, it was concluded that non-uniform flow field radiated noise in the open-rotor engine was mainly focused on low frequency range, the maximum noise was 142dB, and the maximum noise frequency was 115Hz. With the increase of frequency, multiple local peak sound pressures appeared alternately, and noise attenuation rate decreased. A method was proposed to predict the noise for the next generation open-rotor engine.
Numerical simulation of secondary flow control on high-speed compressor cascade with synthetic jets
2018, 33(4): 792-802. doi: 10.13224/j.cnki.jasp.2018.04.004
Abstract PDF(1515KB) icon1586 icon1051
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A simulation study on the use of synthetic jets for the corner separation control on a high-speed compressor cascade was conducted. The mechanisms of the active flow control on delaying flow separation and reducing flow losses were analyzed in detail. Results showed that the space-time structure of the flow field in the blade passage was improved significantly with a maximum loss reduction of 19.8%, while the static pressure rise increment of 8.8% was gained. By periodically adding or removing fluid to or from the boundary layer, synthetic jets could effectively control the corner separation. In the blowing phase, synthetic jets remarkably increased the energy level of the blade suction surface boundary layer and the corner region, while the accumulation of high entropy and low energy fluid towards the corner region was effectively suppressed in the suction phase. The above two contributed to enhancing the resistance to streamwise adverse pressure gradient and thus reducing flow separation. It was worthy to mention that the flow control effect during the suction phase was much more pronounced. Both jet angle and jet momentum were important parameters for the flow control effect. When the jet convected tangentially downstream, the flow control was very efficient in injecting high momentum fluid into the boundary layer, accounting mainly for the resistance to streamwise adverse pressure gradient. In addition, increasing the jet momentum could further improve the cascade time-averaged aerodynamic performance. Factorial design studies indicated that the effect of jet angle was much more pronounced than that of jet momentum, while there was no evidence to support the interaction between jet angle and jet momentum.
Effects of moving endwall on tip leakage flowin compressor cascade
2018, 33(4): 803-811. doi: 10.13224/j.cnki.jasp.2018.04.005
Abstract PDF(1124KB) icon1181 icon956
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A numerical simulation was carried out to investigate the effect of the moving endwall on the tip leakage flow structure in a compressor cascade with different tip clearances. A detailed comparison of leakage flows loss, trajectory, size and mass flow rate, as well as the tip leakage flow features and its interaction with the mainstream, was conducted. Results showed that the moving endwall led to more leakage mass flow through the gap. The region of the tip leakage vortex was flattened away from the blade suction side. The interaction between the mainstream and the tip leakage flow was weakened. The structure of the vortexes inside the passage was altered. In the small clearance cases, the main effects of the moving endwall were focused on the near wall region. While in the large clearance cases, the moving endwall clearly suppressed the tip separation vortex, thus influencing the velocity distribution of the whole clearance, and further impacting the shear between the mainstream and the tip leakage flow.
Influence of wet compression on performance of compressor rotor working at its stall boundary
2018, 33(4): 812-822. doi: 10.13224/j.cnki.jasp.2018.04.006
Abstract PDF(1999KB) icon997 icon777
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At the stall boundary condition of a transonic compressor rotor, numerical simulation method was used to study the effects on operation performance with different inlet fogging conditions. Correlation curves between performance parameters, including air flow, total pressure ratio, isentropic efficiency and specific work, and inlet fogging parameters, including water droplet diameter, velocity, temperature and water spray rate, were obtained. Results showed that inlet fogging can increase air flow, heighten total pressure ratio, enhance isentropic efficiency and reduce specific work. Through analyzing the flow field in detail, it was found that inlet fogging can inject high momentum water droplet to tip stagnation zone, decrease blockage region, move down the interface between main flow and leakage flow, weaken the strength of shock, reduce tip load and increase load at mainstream region. A parameter named water droplet area flow was taken to evaluate the effect on flow field with different fogging conditions quantitatively.
Flow characteristic of tip turbine of additional ducted-fan system
2018, 33(4): 823-831. doi: 10.13224/j.cnki.jasp.2018.04.007
Abstract PDF(1213KB) icon1174 icon971
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Based on a novel high bypass ratio turbofan propulsion system with additional tip-driving fan, the flow characteristics and working mechanism of the core components tip-turbine were investigated using numerical simulation, which laid a theoretical foundation for the further development of high bypass ratio propulsion. Research results indicated that tip-turbine was essentially the axial flow turbine with low solidity and low aspect ratio characteristic, solidity can be as low as 0.6, aspect ratio as low as 0.4. The reduction of solidity caused the migration of throat position, which led to a substantial decrease in the flow deflection and the acceleration of the expansion; the gap leakage loss caused by the low aspect ratio was 2 times larger than the original one, and the influence of leakage vortex was gradually extended to the vicinity of the root of the blade, which reduced the effective extraction of energy area and the energy extraction rate greatly; for the novel power system, effective extraction of energy region and energy extraction rate were proposed to measure the specific work of tip-turbine qualitatively and quantitatively.
Numerical investigation on three-dimensional flow-field characteristics of propeller slipstream
2018, 33(4): 832-840. doi: 10.13224/j.cnki.jasp.2018.04.008
Abstract PDF(1133KB) icon1195 icon963
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The flow-field characteristics of propeller slipstream under different advance-diameter ratios were carried out by three-dimensional numerical simulation. The induced coefficient was used to quantify the slipstream effect; the slipstream characteristics were deeply explored on the basis of the vortex theory. Result showed that the accelerated effect of the slipstream increased firstly, and then decreased along the radial direction, while the twisted effect of the slipstream decreased along the radial direction; both the accelerated effect and the twisted effect decreased with the increment of the advance-diameter ratio; the flow-field of the propeller slipstream can be well described by the axial induced coefficient and circumferential induced coefficient with circumferentially averaged treatment as well as the vortex system structure.
Plasma body force phenomenological model for different altitudes
2018, 33(4): 841-849. doi: 10.13224/j.cnki.jasp.2018.04.009
Abstract PDF(1095KB) icon1056 icon887
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In order to use the plasma body force phenomenological model to simulate the plasma flow control at different altitudes, if assuming the ion number density was directly proportional to plasma light emission, the relative light intensities of plasma emission and light intensity ratios i.e. ion number densities with gas pressure and applied voltage, were obtained by processing the plasma discharge image to gray scale at varying pressures 101.3-1.5kPa. It was found that when the pressure increased, the discharge became filamentary from diffuse, while the light intensity ratios decreased. The saturated relative light intensity in the atmospheric pressure can be used to represent one in other pressures. When the applied voltage increased, discharge occurred at the double-sides of actuator, and the light intensity increased. The total discharge had a linear relationship with applied voltage. Then the new charges boundary condition was built by theory analysis, and the relative light intensities of several plasma emissions at different gas pressure and applied voltage were fitted. Two typical cases were simulated, and the results were used to compare with the plasma jets at corresponding gas pressures measured by particle image velocimetry. It is shown that the new model can simulate the influences on the plasma jets by flight altitude and applied voltage.
Design of inward turning inlet with simultaneous control of intake axial and horizontal projections
2018, 33(4): 850-857. doi: 10.13224/j.cnki.jasp.2018.04.010
Abstract PDF(949KB) icon1442 icon851
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Based on the rotational method of characteristics and streamline tracing technique, the design method was developed for inward turning inlet with control of the axial and horizontal projections of intake, so as to meet the integrated design requirements of hypersonic vehicle waverider forebody and inlet with both sides intake layout. The inward turning inlet was designed utilizing this design method for the axial and horizontal projections of intake of super ellipse. Numerical simulation was conducted at design point Mach number of 6.0. The results indicate that the axial and horizontal projections of intake conform to the anticipated design. The inlet can retain the wave structure and pressure distribution of basic flowfield, and capture all of free incoming flow under the inviscid condition. Its performance of throat plane is almost equal to basic flowfield. Under the viscous condition, the inlet has high compression efficiency; the flow coefficient and total pressure recovery coefficient of exit plane are 0.96 and 0.56, respectively. This design method is feasible and effective.
Thrust characteristics test technique of vectoring nozzle in wind tunnel
2018, 33(4): 858-864. doi: 10.13224/j.cnki.jasp.2018.04.011
Abstract PDF(763KB) icon1293 icon973
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The research of the test technology includes the development of jet simulator and ground calibration system, the data correction method and the wind tunnel verification test. Jet simulator with built-in thrust balance and calibration system with ground measurement test capability were designed and manufactured. Three main factors affecting the thrust balance measurement results were analyzed, including additional stiffness effect, pressure effect, and flowing effect. Then, a correction method of measurement data was established through calibration system. After that, a ground experimental study on the thrust and thrust angle of the nozzle with four deflection angles at 0°,5°,10° and 15° was carried out under different nozzle pressure ratios. At last, the jet simulator was installed on the aircraft model in 8m×6m wind tunnel of CARDC, and an experimental study on the longitudinal aerodynamic characteristics with nozzle pressure ratio of 3 was carried out. According to the experimental result, this test technique can be used to measure the thrust, the vector angle and the aerodynamic characteristics of the nozzle. From the measurement result, the effective thrust angle of deflection was the maximum when the nozzle pressure ration was 2. At this point, the actual deflection angle was 10°, and the effective deflection angle can be increased by 3°.When the nozzle was deflected by 10°, the thrust had the biggest influence on the aerodynamic force of the model. The lift coefficient can be increased by 0.066.
Improving design and flow characteristics analysis ofan integrated wave-rider inlet
2018, 33(4): 865-873. doi: 10.13224/j.cnki.jasp.2018.04.012
Abstract PDF(1272KB) icon1147 icon989
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For improving the performance of an integrated wave-rider inlet, the arc tangent Mach number basal flow and enhanced osculating axisymmetric control lines were adopted in the design process. Numerical simulation was conducted at on-design and off-design conditions and flow characteristics were analyzed respectively. At the design condition, the wave-rider inlet had excellent performance and the characteristic of forebody/inlet integration, however, the cross-flow phenomena appeared in the boundary-layer of compression surface as a result of cutting off. At different angles of attack, the configurations of leading edge shock were nearly the same, helping to enhance the flow capture capability at off-design conditions. For different angles of sideslip, the regularity of the performance varying with angle of sideslip was nonlinear as a result of flows ununiformity. A low-pressure region appeared at the leeward side of inner compression part, leading to a stream wise vortex. As the gas flowed to the low-pressure region on different slices, a dual-vortex structure was formed at the inner compression part. As the gas flowed out of low-pressure region, the second vortex was dissipated.
Optimization of trailing edge angles of single expansion ramp nozzle based on Kriging method
2018, 33(4): 874-881. doi: 10.13224/j.cnki.jasp.2018.04.013
Abstract PDF(874KB) icon1152 icon587
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Based on three-dimensional numerical simulation, different trailing edge angles of single expansion ramp nozzle (SERN) were studied under the condition of nozzle pressure ratio of 25 as the design point, and nozzle pressure ratio of 15 as the undesign point. The trailing edge angles were optimized to get a higher axial thrust coefficient under the nozzle pressure ratio of both 25 and 15 by the Kriging method. Results showed that the trailing edge angles made plume expanding better and the three-dimensional effect of fluid was enhanced. This was propitious to improve the axial thrust coefficient. Whats more, within the design range, the axial thrust coefficient increased from 0.94 to 0.975, about 5% after optimization. So the optimization based on Kriging method is effective.
Analysis of parameteric influence on phase change heat transfer characteristics of 1-D and 2-D electrothermal deicing
2018, 33(4): 882-893. doi: 10.13224/j.cnki.jasp.2018.04.014
Abstract PDF(1245KB) icon1073 icon613
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Comparative analysis and numerical study on phase change heat transfer characteristics of one-dimensional (1-D) and two-dimensional (2-D) electrothermal deicing were presented. The effects of heating mode, cooling time, heater power and heater gap on phase change heat transfer characteristics were emphasized. The enthalpy model was applied due to the enthalpy-porous medium method. The computational domain was treated as a porous medium including multi-layer material, ice and water and mushy zone. The structured mesh topology was used to distribute the computational domain. The finite volume method was adopted to discretize the governing equations. The temperature was obtained by iteration of the energy equation coupled with the liquid volume fraction formula. The properties of the mushy zone can be obtained by linear interpolation. The interface temperature between different materials was obtained. The variation of interface temperature between ice and shield was emphasized. Due to the non-uniformity characteristics of ice-shield interface temperature of 2-D electrothermal deicing model, the temperature criterion to judge ice shedding coupled with hot point and cold point temperature of ice-shield interface was considered. It shows that periodic heating mode for high heater power is superior to continuous heating mode for low heater power. If reasonable combination of cooling time and heater power is adopted, better deicing performance and less energy consumption can be obtained. The reasonable distribution of heater gap can largely improve the deicing efficiency, but leading to the formation of hot point and cold point along ice-shield interface. It will make the ice-shield interface temperature uneven. The cold point at ice-shield interface is similar to stubborn anchor point, which will make the ice layer adhere to the shield surface firmly even if the ice at hot point melts completely. Therefore, the temperature characteristics of hot point and cold point at ice-shield interface for electrothermal deicer should be considered to judge the ice shedding from the shield surface.
Combustion characteristics of methane/n-decane mixed fuel
2018, 33(4): 894-902. doi: 10.13224/j.cnki.jasp.2018.04.015
Abstract PDF(1012KB) icon912 icon566
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In order to study the combustion characteristics and combustion stability of CH4/n-decane mixed fuel, the diffusion velocity of flame radius, the stretched flame speed, Marksein length, unstretched flame speed and laminar burning velocity of CH4/n-decane mixture were measured experimentally in a constant volume chamber at initial pressure of 0.1MPa, initial temperature of 420K, over the equivalence ratios range of 0.8－1.5 and over the methane concentration range of 0－0.8. The influences of methane concentration on Marksein length and laminar burning velocity were investigated. Results showed that at equivalence ratio of 1.3, the observed cellular structure at the end of flame propagation disappeared with the increase of methane concentration and the flame front became smooth, along with the increase of the flame stability; the decreasing trend of Marksein length with the growing equivalence ratio became slower due to the increase of methane concentration, and there was a critical equivalence ratio, with the methane concentration of 0.8 the combustion stability was poor when the equivalence ratio was smaller than 1.2 and to the contrary the combustion stability was better when the equivalence ratio was bigger than 1.2; over the equivalence ratio of 1.0－1.3 the laminar burning velocities of mixed fuel with methane concentration of 0.2 and 0.4 were faster, and over the equivalence ratios range of experiment conditions the laminar burning velocities of mixed fuel with methane concentration of 0.8 were slower.
Calculation method to separation efficiency of impeller type ventilation
2018, 33(4): 903-910. doi: 10.13224/j.cnki.jasp.2018.04.016
Abstract PDF(781KB) icon914 icon611
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In order to investigate its separation performance and establish the method for calculating the separation efficiency compliant with the requirements of engineering design, stress analysis of oil droplet was conducted, the trajectory model of oil droplet was established, and the minimum oil drop diameter under different conditions was calculated combining with the oil droplet capture condition, finally realizing the theoretical calculation of impeller ventilator separation efficiency. Research showed that in different speeds, mass flow rates, and inlet temperatures, the separation efficiency and the minimum separation diameter were in agreement with the experimental values; the maximum errors of separation efficiency were 0.95%, 1.01% and 1.01%, respectively; the maximum errors of the minimum separation diameter were 1.33%, 2.29% and 6.20%, respectively. The calculation results were in good agreement with the experimental results. This method can provide a reference for designing the impeller type ventilator and optimizing its separation performance.
Determination method for thermal test condition of the tip wedge structure heated by high temperature gas flow
2018, 33(4): 911-918. doi: 10.13224/j.cnki.jasp.2018.04.017
Abstract PDF(1032KB) icon855 icon563
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In view of to the traditional thermal test method based on the cold wall heat flux, a determination method based on the curves of the hot wall temperature and heat flux varying with time was proposed. Certain modification to the heat flux was made based on the deviation between the target and measured wall temperatures, in order to eliminate and remedy the test error during the earlier stage. And it determined the temperature of gas flow exactly and rapidly by interpolation in the database of the 300K cold wall heat flux, which solved one of the problems of transient thermal test along the trajectory. The heat flux database of tip wedge structure with variant gas flow rate and temperature was set up by CFD numerical analysis. The aerodynamic heating required for simulation of 9 typical conditions along a hypersonic flight trajectory was obtained and proved to be quite accurate for the requirements of engineering test by numerical simulation with the average error of heat flux of 4.5% in the stagnation point region, 4.6% in the rear plate region, and both maximum error of no more than 8%. The precision meets the requirements of engineering test. Moreover, transient thermal analysis results show that in forty-fifth seconds the biggest temperature gradient is 21K/mm and 18K/mm separately at 1mm and 10.1mm distance from the stagnation point, meeting the demand of high temperature gradient effect of aerodynamic heat.
Experiment on mass flow distribution of ground gas turbine single tube combustor
2018, 33(4): 919-927. doi: 10.13224/j.cnki.jasp.2018.04.018
Abstract PDF(1167KB) icon1074 icon621
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The air flow distribution in a designed single tube combustor (for 100kW ground gas turbine) was experimentally investigated. Under the atmospheric pressure condition, the flow distribution characteristics of the swirlers, primary holes and dilution holes were experimentally obtained using hole plugging method separately, and the discharge coefficients for different structures were analyzed based on the experimental data. Then the flow distribution and total pressure loss of the single tube combustor were studied under different inlet air flows. Results showed that the air flow distributions of different structures were stable when the inlet air flow varied, and agreed very well with the designed data. With the increase of the inlet air flow (inlet Reynolds number), the discharge coefficients for swirlers, primary holes and dilution holes decreased slightly, but the total pressure loss increased gradually. The experimental results of primary holes and dilution holes by using two flow distribution experiment methods were slightly different but basically the same for the air distribution, therefore, the two different experiment methods were reliable. The experimental results can provide an optimization basis for the design of 100kW ground gas turbine single tube combustor.
Experiment of particle rebound characteristics for impingement on resin coated/Al/alloy steel plate
2018, 33(4): 928-935. doi: 10.13224/j.cnki.jasp.2018.04.019
Abstract PDF(899KB) icon999 icon557
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In order to accurately explore the helicopter performance of inlet particle separator, the rebound characteristics of sand impinging different material surfaces were studied. The experiment of particle diameters of 600－800μm sands impacting resin coated/Al/alloy steel plate was conducted. A self-designed experimental system can realize the single particle impact plate. The velocity vectors before and after impact were measured with the high speed camera, within a range of impact angle from 10° to 80°. Results showed that the normal restitution coefficient with sand impinging resin coated plate was higher than that with sand impinging Al/alloy steel plate, while the tangential restitution coefficient was mainly affected by the friction. At the same time, due to the influence of the shadow effect, the normal restitution coefficient or tangential restitution coefficient was more than 1 respectively when the impact angle was close to tangential or normal impact direction. At different scavenging ratio(SCR), the separation efficiency of simulation using the restitution coefficient of sand impinging Al plate was closest to the result of recent experiment on inlet particle separator.
Ablation experimental method for insulation of secondarycombustion chamber in solid ducted rocket
2018, 33(4): 936-943. doi: 10.13224/j.cnki.jasp.2018.04.020
Abstract PDF(1250KB) icon1133 icon569
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Self-designed oxygen-kerosene ablation system was employed to propose an ablation test method for testing the ablation property of the thermal structure and material of secondary combustion chamber in solid ducted rocket (SDR). Particle-free erosion and alumina particle erosion tests were performed on the silicone rubber insulation by using this method to analyze the erosion of particles on the surface micro-morphology of the silicone rubber material. Experiment results revealed that the samples without particle erosion swelled and delaminated after ablation, whose mean linear ablation rate was -0.025mm/s; the mean linear ablation rate of the samples with particle erosion was 1.901mm/s, the ceramic layer in center of samples was destroyed by particle absolutely, only the ultrathin pyrolytic layer and the head of carbon fiber (CF) were left, while the rubber matrix, particle fillers and aramid fiber (AF) around CF were oxidized and stripped by high speed flame flow, then the intense erosion damage effect of particle on insulation had been proved. The research shows that the method can simulate the ablation thermal condition of secondary combustion chamber in SDR, meanwhile, this method can also be applied to selection of insulation formula and ablation performance test of insulating materials.
Simulation of inner flow characteristics of swirl injector during the filing process
2018, 33(4): 944-951. doi: 10.13224/j.cnki.jasp.2018.04.021
Abstract PDF(974KB) icon914 icon565
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Inner flow of fluid (VOF) method was applied in the simulation of inner flow characteristics of a swirl injector during the filling process. The resolution of the phase interface was investigated, the phenomena of the shrink of air core and the rotation of liquid film were discovered. The changes of liquid film thickness and spray angle at the injector outlet during the filling process were calculated by extracting the contours of gas phase volume fraction. The results showed that the liquid film thickness increased along with the outlet mass flow rate, while the spray angle decreased. Besides, the circumfluence in the swirl chamber was described, and the pressure field and velocity field were analyzed, indicating the change of the axial velocity was attributed to the co-effect of pressure drop and gas-liquid interaction.
Numerical simulation of flow field characteristics of UDMH/NTO rocket engine plume
2018, 33(4): 952-960. doi: 10.13224/j.cnki.jasp.2018.04.022
Abstract PDF(972KB) icon1145 icon587
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AUSM (advection upstream splitting method) spatial discretization scheme was used to explore the effects of three k-ε turbulence model and second，third orders upwind schemes on shock capturing. The simulation of unsymmetrical dimethyl hydrazine/nitrogen tetroxide(UDMH/NTO) rocket engine nozzle and exhaust plume flow-field at different altitudes and Mach numbers was modeled with standard k-ε two equation turbulence model and two order upwind scheme. Afterburning was modeled with 12-component 18-chemical reaction, and the entrance parameters of the nozzle were used by thermodynamic calculation. Results showed that with the increase of flow Mach number, the wave number decreased; the plume affected area gradually expanded with the rise of altitude; afterburning reaction caused the significant decrease of O2,N2 in the mix zone while the contents of O, OH, NO in the mixing zone increased.
Design method for the under platform damper of rotor blade
2018, 33(4): 961-968. doi: 10.13224/j.cnki.jasp.2018.04.023
Abstract PDF(922KB) icon1043 icon619
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The interactional relationship among the configuration parameters, contact parameters and dynamic parameters from the point of view for engineers and the need of engineering applications was discussed. The design process and optimization method of the under platform damper were further established to give the evaluation parameters for the rationality of the structure design. As the result of the optimization of a lamellar under platform damper, the uniformity of contact stress and the effective contact area can be improved by optimizing the section shape of the under platform damper. Meanwhile, the method can reach excellent damping effects for multiple operating rotation speeds.
Effect of dent-type foreign object damage on fatigue strength of TC4 material
2018, 33(4): 969-979. doi: 10.13224/j.cnki.jasp.2018.04.024
Abstract PDF(1322KB) icon1089 icon570
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For the dent-type foreign object damage (FOD) suffered by the concave side/convex side of fan/compressor blades, experimental study on FOD simulation, analysis of damage characteristics and stress concentration were carried out under different impact angles, while high cycle fatigue tests and fatigue strength prediction of as-impacted and residual stress relief annealed specimens were conducted. Conclusions were drawn as follows: damage depths and stress concentration factors increased with the increase of the impact angle. The range of the damage depth was 0.1-0.5mm, and the range of the stress concentration factors was 1.3-1.7. Under the condition of different impact angles, the fatigue strength of the dent-type damage declined relative to smooth specimens within the range of 50%-70%, and didnt show a monotonic decreasing relationship with stress concentration factors. The worst angle was 60°. The high cycles fatigue(HCF) behavior of residual stress relief annealed specimens was improved, which indicated that effect of residual stress could not be ignored. The HCF behavior of residual stress relief annealed specimens did not decrease with the increase of stress concentration factors, which validated the effect of microstructure damage. It showed that the damage depth as the single parameter to build allowable limits and repair limits had some limitations. Prediction accuracy of fatigue strength for dent-type damaged specimens was within ±20%.
Flutter boundary prediction research depending onsystem stability analysis methods
2018, 33(4): 980-988. doi: 10.13224/j.cnki.jasp.2018.04.025
Abstract PDF(934KB) icon1051 icon604
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A research was made on two flutter boundary prediction methods: flutter margin method and auto-regressive and moving average model (ARMA) stability analysis method. The influence of parameter identification errors on the accuracy of boundary prediction was analyzed through numerical simulation case. And the prediction effectiveness of these two methods was analyzed by two wind tunnel flutter tests. The results showed that: (1) the identified damping error had less influence on the criterion of flutter margin method than the identified frequency error, and ARMA stability analysis method was more sensitive to the identified damping error than flutter margin method contrastively, thus the flutter margin method had better robustness than ARMA stability analysis method. (2) For wing bending-torsion coupling flutter, these two criterions had approximate variation curves, which differed in order of magnitude approximating to the biquadrate of the sampling frequency, and good decreasing trends helpful to predict flutter boundary in advance were presented. (3) For in-plane bending flutter, ARMA stability analysis method may have better applicability than flutter margin method.
Vibration characteristic for face gear based onload equivalent installation errors
2018, 33(4): 989-998. doi: 10.13224/j.cnki.jasp.2018.04.026
Abstract PDF(1314KB) icon805 icon532
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Considering the support structure deformation and tooth elastic deformation of the face gear drive, a multi degree of freedom coupling vibration analysis model including equivalent installation errors was established by converting the transmission deformation into installation errors. Fourth order Runge-Kutta method was adopted to solve the differential equation of face gear drive. The influences of equivalent installation errors on the vibration acceleration and the dynamic meshing force of the face gear were obtained. The dynamic characteristics experiments for face gear drive were performed. The experimental results showed that the vibration acceleration along x direction caused by offset and shaft angle error was larger than that along z direction. The influence of vibration acceleration along x direction caused by offset error was larger than that caused by shaft angle error.
Thermal friction characteristics analysis of helical gears under non-Newtonian fluid and big slide-roll ratio
2018, 33(4): 999-1008. doi: 10.13224/j.cnki.jasp.2018.04.027
Abstract PDF(1123KB) icon972 icon556
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Considering the time-varying contact line length and gear tooth load during a meshing cycle, and combining the transient equivalent curvature and entrainment velocity along the contact line, a thermal elastohydrodynamic lubrication (TEHL) model of finite line contact for helical gears was developed under non-Newtonian fluid and big slide-roll ratio for the aerospace transmission. By employing multilevel method and sequential sweeping method, the complete TEHL numerical solution suitable to wide working conditions was obtained. The thermal friction performances for helical gears transmission were investigated in non-Newtonian fluid and big slide-roll ratio cases. Results showed that at the end of helical gears the obvious pressure and temperature peak existed, moreover the film thickness reduced slightly approaching to the engage-in position. Non-Newtonian characteristics exerted a significant influence on temperature, furthermore the temperature distribution behaved like a V-shape around the pitch point. The friction coefficient of the pitch point along the meshing line was close to zero, which gradually increased towards the two sides and with the going up of slide-roll.
Intake and exhaust system optimization of UAV piston engine
2018, 33(4): 1009-1016. doi: 10.13224/j.cnki.jasp.2018.04.028
Abstract PDF(842KB) icon887 icon529
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To solve the problem of torque decrease and high fuel consumption at 6500r/min flying speed of an unmanned aerial vehicle for plant protection, a multi-parameter optimization method using adaptive genetic algorithm(GA) based on 1-D thermodynamic simulation model was proposed to optimize the intake and exhaust system of single cylinder gasoline engine. GT-Power model of the UAV engine was built to simulate the engine working condition. The model was calibrated by experimental bench data to ensure the accurate simulation. Influence of structure parameters of intake and exhaust system on torque and fuel consumption was analyzed through design of experiment, and length of intake pipe and exhaust pipe, diameter of intake pipe and second cavity volume of air filter were chosen as optimization variables. The model was used to calculate the objective optimization function and limit conditions. Adaptive genetic algorithm was executed by Matlab which exchanged data by interface of Simulink/GT-Power to receive and feedback results. According to the optimization result of adaptive GA, new intake and exhaust system was mounted on the engine to verify optimization result. Result demonstrated that torque and fuel consumption at flying speed of 6500r/min of the engine was obviously improved by adaptive GA, and torque was raised by 5.51% and fuel consumption was improved by 6.31%.
Estimated method of target level of safety for unmanned aircraft system
2018, 33(4): 1017-1024. doi: 10.13224/j.cnki.jasp.2018.04.029
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Abstract:
Considering the differences of safety features between unmanned aircraft system(UAS) and manned aircraft, the ground impact was recognized as the critical risk condition by means of event tree analysis. Based on definition of equivalent level of safety, safety evaluation model for unmanned aircraft system was established by consideration of kinetic energy, configuration and flight environment factors. Target levels of safety evaluation approaches based on different UAS classes operating in different scenarios and Chinese population density were discussed. Results indicated that quantitative requirement for UAS operating in different scenarios may range between 4 or 5 orders of magnitude. On this basis, risk evaluation matrix of UAS operation based on ground impact condition was proposed, providing an effective approach to risk evaluation matrix for operations of UAS.

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