吸气式多级电流体推进器性能研究

Performance investigation of air-breathing multi-stage electrohydrodynamic thruster

吸气式电流体动力(Electrohydrodynamic,EHD)推进器通过多级加速的方式可以有效提高推力,缓解航空器在空间环境下燃料不足的问题。为了探索不同参数下推进器的性能变化规律,分别研究了海拔、电压、电极间距和推进器级数等对电流体推进器的影响,通过求解电流体力学基本方程获得了放电、流动及推进特性相关的数据。研究发现,推进器在海拔0~20 km存在较大的性能差异,当海拔增大时,电极可施加的极限电压明显减小,电流及空气流速均有不同程度的降低。单级加速腔在海拔0 km处产生的最大推力是海拔20 km处的18.4倍,对应的推功比是海拔20 km处的3.1倍。多级推进器的推进性能并不等同于多个单级加速腔的代数相加,其前后两级电极的相互干扰会导致电场强度和电荷密度均发生变化,产生的反向电场将阻碍气体的流动。当推进器级数由1提升至5时,总推力由1.7 mN提升至6.53 mN,增大3.84倍;而推功比由0.94 N/kW降至0.58 N/kW,减小61.7%。

Air-breathing electrohydrodynamic(EHD)thrusters offer a solution to the issue of limited fuel for aircraft in space.Researchers typically employ a multi-stage acceleration approach to enhance the thrust of these thrusters.To identify the performance variation of thrusters under different parameters,the effects of alti-tude,voltage,electrode spacing and thruster stages on EHD thruster are investigated,by solving the basic EHD equations to obtain data on the characteristics of discharge,gas flow and propulsion.The results indicate clear disparities in the performance of thrusters at altitudes of 0 km and 20 km.As the altitude increased,the maxi-mum voltage that could be supplied to the electrode reduced significantly,resulting in varied reductions in cur-rent and air velocity.The maximum thrust generated by a single-stage accelerating cavity at 0 km altitude was 18.4 times that at 20 km altitude,and the corresponding thrust-to-power ratio was 3.1 times that at 20 km alti-tude.The propulsive performance of a multi-stage thruster did not exhibit an algebraic relationship to the propul-sive performance of single stage accelerating cavities.The interaction between the front and rear electrodes caused alterations in both the intensity of the electric field and the density of charge,resulting in a hindrance to the gas movement due to the opposing electric field.When the thruster stage increased from 1 to 5,the overall thrust increased from 1.7 mN to 6.53 mN,resulting in 3.84 times.Additionally,the thrust-to-power ratio re-duced from 0.94 N/kW to 0.58 N/kW,representing a reduction of 61.7%.