The precise regulation of the hinge moment and pitch angle driven by the pitch-regulated mechanism is crucial for modulating thrust requirements and ensuring stable attitude control in Martian coaxial rotorcraft. Nonetheless, the aerodynamic hinge moment in rotorcraft presents time-dependent dynamic properties, posing significant challenges for accurate measurement and assessment for such characteristics. In this study, we delve into the detailed aerodynamic hinge moment characteristics associated with the pitch-regulated mechanism of Mars rotorcraft under a spectrum of control strategies. A robust computational fluid dynamics model was developed to simulate the rotor’s aerodynamic loads, accompanied by a quantitative hinge moment characterization that takes into account the effects of varying rotor speeds and pitch angles. Our investigation yielded a thorough understanding of the interplay between aerodynamic load behavior and rotor surface pressure distributions, leading to the creation of an empirical mapping model for hinge moments. To validate our findings, we engineered a specialized test apparatus capable of measuring the hinge moments of the pitch-regulated mechanism, facilitating empirical assessments under replicated atmospheric conditions of both Earth and Mars. The result indicates aerodynamic hinge moments depend nonlinearly on rotational speed, peaking at a 0° pitch angle and showing minimal sensitivity to pitch under 0°. Above 0°, hinge moments decrease, reaching a minimum at 15° before rising again. Simulation and experimental comparisons demonstrate that under Earth conditions, the aerodynamic performance and hinge moment errors are within 8.54% and 24.90%, respectively. For Mars conditions, errors remain below 11.62%, proving the CFD model’s reliability. This supports its application in the design and optimization of Mars rotorcraft systems, enhancing their flight control through the accurate prediction of aerodynamic hinge moments across various pitch angles and speeds.

A xenon micro flow control device (XMFCD) is the key component of a xenon feeding system, which controls the required micro flow xenon (µg/s–mg/s) to electric thrusters. Traditional XMFCDs usually have large volume and weight in order to achieve ultra-high fluid resistance and have a long producing cycle and high processing cost. This paper proposes a miniaturized, easy-processing, and inexpensive XMFCD, which is fabricated by low-temperature co-fired ceramic (LTCC) technology. The design of the proposed XMFCD based on complex three-dimensional (3D) microfluidic channels is described, and its fabrication process based on LTCC is illustrated. The microfluidic channels of the fabricated single (9 mm diameter and 1.4 mm thickness) and dual (9 mm diameter and 2.4 mm thickness) XMFCDs were both checked by X-ray, which proved the LTCC method’s feasibility. A mathematical model of flow characteristics is established with the help of finite element analysis, and the model is validated by the experimental results of the single and dual XMFCDs. Based on the mathematical model, the influence of the structure parameters (diameter of orifice and width of the groove) on flow characteristics is investigated, which can guide the optimized design of the proposed XMFCD.

During the operation of ADN (ammonium dinitramide, (NH⁴⁺N(NO₂)²⁻))-based thrusters, the ADN-based liquid propellant, a non-toxic green energetic material, tends to flow boil in the capillary tube due to heat transfer from the wall. A three-dimensional transient numerical simulation of the flow boiling of ADN-based liquid propellant in the capillary tube was carried out using the VOF (Volume of Fluid) coupled Lee model. The flow-solid temperature and the gas–liquid two-phase distribution and the wall heat flux at different heat reflux temperatures were analyzed. The results show that the magnitude of the mass transfer coefficient of the Lee model significantly influences the gas–liquid distribution in the capillary tube. The total bubble volume increased from 0 mm³ to 957.4 mm³ when the heat reflux temperature was increased from 400 K to 800 K. The bubble formation position moves upwards along the inner wall surface of the capillary tube. Increasing the heat reflux temperature intensifies the boiling phenomenon. When the outlet temperature exceeded 700 K, the transient liquid mass flow rate in the capillary tube was already reduced by more than 50%. The results of the study can be used as a reference for the design of ADN-based thruster.

This study investigates the impact of rotor spacing on the aerodynamic performance of a coaxialcopter and promotes an innovative regulated control strategy for the coaxial drone. The present research introduces a coaxialcopter with variable rotor spacing, and employing finite element numerical simulations, we assess the aerodynamic behavior of this novel configuration. Through comprehensive measurements and analysis of its aerodynamic performance across varying rotor spacings from 0.1 R to 1 R, we validate the effectiveness of a rotor-spacing control strategy for enhancing takeoff maneuvers. The numerical simulation and experiment results reveal that the performance characteristics of both the upper and lower rotors converge toward that of a single rotor as the space ratio increases, along with a reduction in their thrust fluctuations and aerodynamic performance periodicity. Considering stable power consumption patterns and endurance performance, we analyzed the interrelations binding the pitch distance of the rotors, rotational speed, and pitch angle, vis à vis the thrust coefficient and power coefficient. Through the parameter optimization method, we demonstrate that adjusting rotor spacing offers a practical means to enhance payload capacity without increasing the power input, thereby improving efficiency, which validates the practicality and efficacy of the parameter optimization approach. Furthermore, optimizing rotor spacing for specific operational scenarios enhances overall aerodynamic performance, suggesting a viable flight control strategy for takeoff and landing conditions for coaxial drones.

An experimental study was carried out on the ignition characteristics of the HAN/(Emim)(EtSO₄) (hydroxylammonium nitrate and 1-ethyl-3-methyl-imidazolium ethyl sulfate) dual-mode ionic liquid monopropellant in chemical propulsion mode in model thrusters. Firstly, a model thruster with a detachable convergent nozzle was designed and fabricated. Secondly, catalytic ignition experiments at different flow rates were carried out in atmosphere and in high chamber pressure environment, respectively, using a model thruster, with and without the convergent nozzle. During the catalytic ignition process, measurement methods such as thermocouple, pressure sensor, and flue gas analyzer were employed to obtain the temperature at different depths of the catalytic bed, the pressure of the combustion chamber, and the concentration variations of gaseous products CO, CO₂, CH₄, SO₂, NO, and NO₂. Then the three characteristic stages of water evaporation, HAN decomposition, and (Emim)(EtSO₄) combustion were analyzed at the initiation time, and the reaction characteristics in the process of the catalytic ignition were analyzed. In addition, the composition and concentration of the combustion products at equilibrium were theoretically calculated. The effects of temperature and pressure on the concentrations of five main gaseous products were studied. Finally, the exhaust gas of the three groups of catalytic ignition experiments under different pressure environments was separately collected and measured with gas chromatography (GC) when the experiments approached equilibrium, the result of which roughly agrees with the theoretical calculations. These results are of great significance for exploring the chemical propulsion of the dual-mode ionic liquid propellant and understanding its physical catalytic combustion mechanisms.