Advances in Astronautics Science and Technology最新文献

The interstellar medium (ISM) represents the next frontier in space exploration, with many new discoveries to be made. The challenge, being so far away from Earth, the ISM requires many decades to reach. To advance our knowledge of what exists beyond our solar system, new approaches for rapid access are required. One such approach is solar thermal propulsion (STP). The approach uses several Venus and Earth gravity assists to fly to Jupiter and use its gravity well to dive towards the Sun. Approaching within three solar radii a perihelion burn would be performed, maximising the spacecraft’s ΔV to achieve high solar system escape velocities. A unique aspect of the STP mission concept is that the Sun is used not only as a gravity well for an Oberth manoeuvre, but also to heat the fuel to ultra-high temperatures (> 3000 K), enabling a monopropellant burn with high specific impulse (I_sp). Prior preliminary studies indicated escape velocities of over 20 astronomical unit (AU)/year would be possible. An in-depth modelling exercise was undertaken to determine how such a system would perform. The model in this paper showed the current STP design is capable of providing just under 9 ± 1 AU/year, but there are many technology developments that could increase escape velocity. The technologies vary from items that could be implemented in the near term, like turbo-pumps driven by the hydrogen, to items requiring more extensive development programs like thin coatings which do not erode in superheated hydrogen. After reviewing the STP approach, and comparing it to a solid rocket motor (SRM), it was found that with currently available technology, SRM outperforms STP with an escape velocity of approximately 10–12 AU/year. However, future advances in heat exchanger lining materials, turbo pumps, and advanced heat exchanger geometries may enable solar thermal propulsion to provide higher escape velocities, providing one of the fastest ways to exit the solar system. Ultimately, if all technology paths could be implemented with minimal side effects, the performance in a best-case scenario could reach up to 16 AU/year.

One of the biggest open questions in the field of asteroid mining are the unknown characteristics and composition of more than 99% of the discovered objects. In order to minimize the risks of false investment, this uncertainty has to be diminished. One strategy to overcome this is to realize space missions that characterize the composition of asteroids in an efficient way. The above described issue is addressed in this paper. A target accessibility analysis is presented, that is tailored for asteroid characterization missions departing from a main belt parking orbit. To enable and facilitate a flexible, non predefined and autonomous object selection, adequate database constraints based on the orbital elements of the targets at departure are presented that enable the filtering of non-reachable objects without excluding potential targets. The constraints are applied to all asteroids currently listed in the JPL Small-Body Database Browser. With the derived constraints, the used database can be reduced to 1% of its original size.

In this paper, a new method of state stability, transition and tracking control for dynamics system based on model prediction and inversion is introduced. The estimation of system state bias in the future is obtained by model prediction. According to the dynamics evolution law of the controlled object, the control command to eliminate the future deviation is obtained by dynamics inversion. The compensation control is designed for the current and historical state deviations, and a state control method based on dynamics model of the controlled object is constructed. This method can be widely applied to various forms of control, and can better realize the state transition, tracking and stability control of dynamics system under the time constraint. The controller can be self-organized according to the model of the controlled object, and the parameters of the controller can be adjusted adaptively. It has good robustness to load/disturbance and deviations of model parameter, state measurement and control execution. The characteristics of this method are discussed, a simulation of rocket attitude control is given and the future research focus of this method is prospected.

In recent years, the mega-constellation programs represented by StarLink have been proposed successively. Mega-constellations compose of far more satellites than traditional constellations and possess more complex constellation configurations, which leads to a sharp increase in the calculation of probability distribution of interference. In this paper, we propose a new method to reduce the computational amount. For the satellite communication system whose operating mode and communication parameters have been determined, the interference value is only related to the position distribution of satellites in different constellations relative to the earth station. Hence, the probability distribution of inter-constellation interference value can be solved by deriving the probability of diverse constellations satellites’ position relative to the earth station. On basis of this, we calculate the occurrence probability of different constellation snapshots and the interference value between snapshots of different constellations in turn. Furthermore, we can use the joint probability of constellations’ satellite distribution to evaluate the probability distribution of interference between mega-constellations. Simulation results demonstrate that the proposed method achieves the same accuracy as the traditional method with higher computation efficiency, and the occurrence probability of maximum interference value can be obtained deterministically, which is suitable for the interference evaluation involving mega-constellations.

Active debris removal (ADR) is promising methods for ensuring safe space activities, free from the danger of debris. To carry out an ADR mission, the attitude and motion of the target must be determined precisely. Developing methodology to extract these values using only the target’s light curve would be a great step forward. We started the light curve observations of the ADR candidates, H2A rocket bodies (R/Bs), 2nd stages of Japanese H2A rockets using the 60 cm telescope, and the CMOS (complementary metal-oxide semiconductor) sensor. We developed an optical simulator in the laboratory to mimic observed light curves. The simulator can reproduce the exact light curve using a scale model of the H2A R/B. It considers the attitude, motion, and lighting conditions of the H2A R/Bs. On March 19, 2019, two extremely strong peaks were observed in the light curve of one of H2A R/Bs (satellite number: 39771). Simulations showed that the observed light curve is explained by the attitude of the gravity gradient stabilization where the PAF (payload attach fitting) of the H2A R/B was directed toward the earth. We found a few degrees’ tilt of the target causes shifts of the timings of the peaks. This means that the attitude of the target can be ascertained using the peak timing in some cases. Although this is one case out of countless situations, simulating exactly the same light curve is the one step toward total understanding of ADR targets’ attitude and motion from light curve observations. We also developed a light curve simulation tool using the 3-D (three-dimensional) model of H2A R/B that can estimate the overall tendency of the light curve, which will dramatically reduce experimental times for simulating light curve using the optical simulator.

Cubesats with imaging payloads face unique challenges in terms of stringent pointing accuracy and stability requirements. Team Anant is a student-run technical team working to build a 3U Cubesat. This paper discusses the implementation, validation and integration of an attitude estimation algorithm as part of the satellite’s Attitude Determination System (ADS). The ADS hardware usually comprises sensors such as an IMU, magnetometer, and sun sensors. Validation methodology and architecture design, which aims to satisfy the allocated pointing budget, are also discussed. The paper introduces the motivation to choose Murrell’s version Kalman Filter and a comparison with popular alternatives. This is followed by some prerequisites, after which, the paper describes the top level overview and testing framework developed for the Kalman Filter. This requires emulating the in-orbit environment and tracking the true state to establish the performance limit with a predefined performance metric. The verification procedure adopted by the team is discussed in detail. Apart from analysing the expected trend of the filter parameters over time, a quasi-Monte Carlo approach was also followed. Furthermore, the Cramer–Rao bound is used to establish a lower bound on the error covariance matrix. Lastly, an approach for fine sensor selection is provided based on emulating its integration with the ADS. The paper concludes by discussing the lessons learnt and the important stages in the development and testing of an attitude estimation algorithm.

Hybrid rocket motors have several attracting characteristics such as simplicity, low cost, safety, reliability, environmental friendliness. In particular, hybrid rockets can provide complex and flexible thrust profiles not possible with solid rockets in a simpler way than liquid rockets, controlling only a single fluid. Unfortunately, the drawback of this feature is that the mixture ratio cannot be directly controlled but depends on the specific regression rate law. Therefore, in the general case the mixture ratio changes with time and with throttling. Thrust could also change with time for a fixed oxidizer flow. Moreover, propellant residuals are generated by the mixture ratio shift if the throttling profile is not known in advance. The penalties incurred could be more or less significant depending on the mission profile and requirements. In this paper, some proposed ways to mitigate or eliminate these issues are recalled, quantitatively analysed and compared with the standard case. In particular, the addition of energetic additives to influence the regression rate law, the injection of oxidizer in the post-chamber and the altering-intensity swirling-oxidizer-flow injection are discussed. The first option exploits the pressure dependency of the fuel regression to mitigate the shift during throttling. The other two techniques can control both the mixture ratio and thrust, at least in a certain range, at the expense of an increase of the architecture complexity. Moreover, some other options like pulse width modulation or multi-chamber configuration are also presented. Finally, a review of the techniques to achieve high throttling ratios keeping motor stability and efficiency is also discussed.

The case for modern solid rocket motor (SRM) is a complex compose structure with fiber reinforced resin matrix composite, insulation layer and metal connections, the case forming process is a multi-physical–chemical process which involving heat transfer, chemical reaction and structure deformation. During the cure process, temperature determines whether the case cure completely and the uniformity of temperature field is an important factor in causing residual thermal stress and shrinkage stress, so temperature field is the key to the mutual coupling effect of each physical and chemical process, and the basis to analyze the cure process of the case. During the cure process of case, fluctuation of cure temperature, cure time or heat transfer of hot air in the furnace may occur, which make the actual forming process of case deviate from the ideal cure process. In order to investigate the sensitive degree of thermal cycle, convective heat transfer coefficient and thermal properties to cure uniformity during cure process of composite case, the influence rule of the three factors on uniformity of temperature and cure degree fields were analyzed by numerical simulation. A thermal-chemical model was built for a simplified composite SRM case firstly, and the model was verified. Then, the influence degree of the three factors on cure uniformity of composite case was analyzed and quantized by the Morris global sensitivity analysis method. The results show that he sensitivity order of the uniformity of temperature field for the four parameters is: thermal diffusion coefficient > heat transfer coefficient > duration time > cure temperature. Besides, he temperature and duration time of the fourth dwell stage have less effect on the cure uniformity of composite case than that of heat transfer coefficient. Therefore, it is a challenge to design a thermal cycle that can not only guarantee the vulcanization of EPDM insulation layer, but also improve the cure uniformity.

One of the important challenges to be addressed during development of a new liquid rocket stage for a launch vehicle is the suppression of pogo. Pogo is a harmful dynamic phenomenon arising from the interaction of vehicle structural axial modes with propulsion fluid system modes. The vehicle structural axial mode frequency generally increases during the course of the stage operation. The likelihood of experiencing pogo phenomenon increases when the vehicle structural axial mode frequency crosses the propulsion system frequency during the flight time. To avoid this harmful effect generally a pogo suppression device (PSD) is installed in feedlines which helps to decouple these two frequencies by lowering the propulsion system frequency to a safe value which is called the targeted design frequency. The propulsion system frequency is lowered by providing the required compliance at the pump inlet. The present work describes the design of a gas filled bellow PSD for a semi-cryogenic stage of ISRO’s new launch vehicle. A gas filled bellow PSD utilizes the stiffness of the bellows and the pneumatic stiffness offered by the compressed gas in the bellows cavity to provide the required system compliance. In general, a PSD is positioned at the pump inlet. Bellows get compressed when the pump inlet pressure increases and undergoes expansion when the pump inlet pressure reduces. One of the challenges in designing a gas filled bellow PSD is achieving the system compliance maintaining bellow deflection within permissible limits and at the same time achieving the targeted design propulsion frequency. Since the stiffness of gas is a function of its pressure and volume at a particular instant, the system is a variable stiffness system which brings in further challenges in designing the pogo corrector. The system is to be designed considering the variation in pump inlet pressure, thermodynamic process the gas undergoes, the instantaneous volume and pressure of the gas and various bounds for each of these parameters. This paper illustrates how the inputs, system requirement and constraints are formulated mathematically so that the entire design is transformed into an optimization problem with envelope of the system as the minimization function. The choice of gas and initial gas pressure and volume to be maintained are also discussed. Finally, the performance of the pogo corrector under various thermal conditions is also looked into.

Pulsed electrothermal thrusters use kiloampere discharge currents for polytetrafluoroethylene (PTFE) ablation and sublimation. However, higher arc currents cause higher electromagnetic interference (EMI). A 10 J surface arc thruster (SAT), which adopts current regulating diodes (CRD), was developed that enables significant reduction in EMI. A CRD limits the discharge currents to 5 A in spite of the applied voltage. A low-melting-point sulphur propellant has been used that enables low-discharge currents to efficiently ablate it and accelerate it electrothermally. In this paper, the near-term potential for elemental sulphur propellant in SAT is investigated. The advantages of sulphur with respect to PTFE are presented. First, we measured the pulse width of the main discharge. Then sulphur propellant proved superior to PTFE propellant in discharge duration. The mean pulse widths of PTFE and sulphur are 3.38 and 22.1 ms, respectively. Second, we measured the pressure rise in the vacuum chamber after each discharge. The mean pressure rises of PTFE, sulphur powder and sulphur solid are 0.43, 0.94 and 1.9 mPa, respectively. Sulphur powder experienced the least intensity of discoloration in comparison with other propellants. The discharge initiator misfired frequently during PTFE ablation, but it ignited successively during sulphur ablation. Experimental results indicate sulphur is a suitable propellant for surface discharge propulsion, and for low arc currents is superior to PTFE.