Bijiao He

Experimental and Numerical Research on the Diversion Effect of a Conic Flame Deflector for a Lunar Module Ascent Stage

AbstractThe diversion effect of a conic flame deflector for a lunar module ascent stage is investigated experimentally and numerically. Eighteen tests were performed in a vacuum chamber mounted with a liquid helium cryogenic pump using a scaled engine and models. The experimental results indicate that the conic flame deflector can guide the exhaust gas away from the ascent stage model more effectively than a plate; meanwhile, the conic flame deflector can create a reverse torque, which helps to correct the deflection of the ascent stage model. Finally, the experimental results show that the reverse torque changes with the distance and angle between the ascent stage model and the conic flame deflector model. Numerical simulations have been done with the computational fluid dynamics and direct simulation Monte Carlo (CFD/DSMC) method, and the numerical results are in good agreement with the experimental results. The numerical results reveal that the reason for the above phenomena is shock waves. An expansio...

Improving the viability and versatility of the E × B probe with an active cooling system

A thermostatic E × B probe is designed to protect the probe body from the thermal effect of the plasma plume that has a significant influence on the resolution of the probe for high-power electric thrusters. An active cooling system, which consists of a cooling panel and carbon fiber felts combined with a recycling system of liquid coolants or an open-type system of gas coolants, is employed to realize the protection of the probe. The threshold for the design parameters for the active cooling system is estimated by deriving the energy transfer of the plasma plume-probe body interaction and the energy taken away by the coolants, and the design details are explained. The diagnostics of the LIPS-300 ion thruster with a power of 3 kW and a screen-grid voltage of 1450 V was implemented by the designed thermostatic E × B probe. The measured spectra illustrate that the thermostatic E × B probe can distinguish the fractions of Xe+ ions and Xe2+ ions without areas of overlap. In addition, the temperature of the probe body was less than 306 K in the beam region of the plasma plume during the 200-min-long continuous test. A thermostatic E × B probe is useful for enhancing the viability and versatility of equipment and for reducing uneconomical and complex test procedures.

A momentum flux measuring instrument with the variable-range for exhaust plume

A momentum flux measuring instrument (MFMI), which ranges from 1 mN∼300 N, is designed to measure the parameters related to the momentum flux of exhaust plume of attitude and orbit control thrusters covered electric and chemical thrusters. A MFMI with the variable-range provides an efficient and economy way to study the exhaust plume of both electric and chemical thrusters. The flexural pivots, replaceable strain gauge, replaceable target plate, and force arm with the variable length make it possible. The designed MFMI with the non-displacement measurement can reduce the influence of pipelines and test lines, especially the elastic force and friction due to displacement. The designed MFMI system with the range of 0∼100 mN and 0∼15 N was calibrated separately. The target indirect measurement method is a simple and economical option for measuring the thrust of electric thrusters, and is employed to measure the thrust of the LIPS-200 ion thruster using the designed MFMI. Research indicates that the sputtering contributes to the momentum transfer of plasma-surface interactions. In addition, the high-speed charged ions from the thruster will experience a charge-exchange (CEX) collision with the slow neutral background. Therefore, the influence of the sputtering and CEX on target indirect measurement method is derived in detail, and then the measured thrust is corrected. The measured average thrust, which is 40.1 mN, demonstrates that the designed MFMI is effective. The designed MFMI can also be employed as a thrust stand by directly mounting the thrusters on torsional arm, and a maximum of three thrusters can be installed on the MFMI at the same time.A momentum flux measuring instrument (MFMI), which ranges from 1 mN∼300 N, is designed to measure the parameters related to the momentum flux of exhaust plume of attitude and orbit control thrusters covered electric and chemical thrusters. A MFMI with the variable-range provides an efficient and economy way to study the exhaust plume of both electric and chemical thrusters. The flexural pivots, replaceable strain gauge, replaceable target plate, and force arm with the variable length make it possible. The designed MFMI with the non-displacement measurement can reduce the influence of pipelines and test lines, especially the elastic force and friction due to displacement. The designed MFMI system with the range of 0∼100 mN and 0∼15 N was calibrated separately. The target indirect measurement method is a simple and economical option for measuring the thrust of electric thrusters, and is employed to measure the thrust of the LIPS-200 ion thruster using the designed MFMI. Research indicates that the sputterin...

Development of a coupled NS-DSMC method for the simulation of plume impingement effects of space thrusters

A coupled NS-DSMC method possessing adapted-interface and two-way coupling features is studied to simulate the plume impingement effects of space thrusters. The continuum-rarefied interface is determined by combining KnGL and Ptne continuum breakdown parameters. State-based coupling scheme is adopted to transfer information between continuum and particle solvers, and an overlapping grid technique is investigated to combine structured-grid NS code and Cartesian-grid DSMC code to form the coupled solver. Flow problem of a conical thruster plume impinging on a cone surface is simulated using the coupled solver, and the simulation result is compared with experimental data, which proves the validity of the proposed method. Plume flow while the ascent stage of lunar module lifting off in lunar environment is also computed by using the present coupled NS-DSMC method to demonstrate its capability. The whole flow field from combustion chamber to the vacuum environment is obtained, and the result reveals that special attention should be paid to the plume aerodynamic force at the early stage of launching process.

A stable Runge-Kutta discontinuous Galerkin solver for hypersonic rarefied gaseous flows

A stable high-order discontinuous Galerkin scheme which strictly preserves the positivity of the solution is designed to solve the Boltzmann kinetic model equations. The stability is kept by the accuracy of the velocity discretization, the conservation of the collision terms and a limiter. By requiring the time step smaller than the local mean collision time and forcing positive values of velocity distributions on certain points, the limiter can preserve the positivity of the solutions of the cell average velocity distributions. Verification is performed with a normal shock wave at Mach number of 2.05 and a supersonic flow about a 2D cylinder at Mach number of 6.0.

Modeling of Lunar Dust Spray Using the DSMC method

To study the interaction of nozzle plume and lunar dust, the plume‐dust two‐phase flow simulation is implemented on the self‐developed DSMC workstation. The momentum and energy transfer to lunar dust particle from the rarefied gas flow was computed. Particle‐particle collision is considered as hard sphere collision. The dust particles are accelerated by the gas and a lunar crater is formed.