Search

With the help of an O 2 -pilot strut, the stable supersonic combustion was achieved in a scramjet combustor fueled by liquid kerosene. Objective of this work is to investigate the effect of different mass flow rate of oxygen on the cold flow and combustion characteristics to explain the role of the supplied oxygen in supersonic combustion. A series of numerical simulations and experiments were carried out under the high inflow enthalpy condition (Ma = 2.8). High-speed photography and pressure measurement were used to record the flame images and pressure of combustor. A skeleton kerosene mechanism was used for combustion simulation for more details of combustion. Simulation results indicated that by adjusting the mass flow rate of oxygen, the severe rich-kerosene zone around the ignitor could be improved. Under the low mass flow rate of oxygen condition, limited by over-high equivalence ratio the ignition delay time was too long and heat release was too small to establish the global flame. On the contrary, under the high mass flow rate of oxygen condition, limited by over-low equivalence ratio, only little kerosene could be ignited with most of the kerosene carried downstream by high-speed flow resulting in a sharp increase in flame lift distance. The boundary of mass flow rate of supplied oxygen for combustion stabilization was obtained. The results in this paper would be valuable for reveling the ignition and flame stabilization mechanism in supersonic airflow. • The flame stabilization boundary of lower and upper oxygen was obtained. • The effect of different mass flow rate of oxygen on the volume of subsonic zone was analyzed. • The influence of supplied oxygen on the least-reactive-kerosene-fraction was discussed. • The mechanism of oxygen on the flame stability and ignition delay time was investigated. [ABSTRACT FROM AUTHOR]

D-type asteroids are among the most primitive small bodies of the solar system. Believed to be formed in the outer solar system, a minor fraction of these faint objects can be found in the near-Earth region. Some were suspected to be extinct comets disguised as asteroids. If D-type near-Earth asteroids could represent extinct comets, they would offer us a unique opportunity to investigate the relationship between two classes of minor bodies in our solar system. To provide new insights into D-type asteroids' composition and dynamical evolution and the possible relation with comets, we introduce the mission concept CARINA (Comet Asteroid Relation INvestigation and Analysis). CARINA will visit and collect a sample from the asteroid 2002 AT4 and address key scientific questions related to our understanding of the early solar system evolution, and the origins of water and life on the early Earth. This paper outlines the scientific motivation and the means for the sample return. The spacecraft is equipped with a sampling ring to perform in-situ analysis and to collect, in a "touch and go" manner, samples from the surface and subsurface of the asteroid. A capsule is expected to return the samples to Earth in pristine conditions for detailed and extended analysis. These would represent a rare contribution to our meteorite collection since it would be the first time that material from a D class asteroid would be collected. • CARINA is proposed to be the first sample return mission to a D-type asteroid. • The mission would address the relationship between comets and asteroids. • CARINA's science case and objectives, subsystems engineering design, and mission analysis are presented. [ABSTRACT FROM AUTHOR]

The feasibility of utilizing Deep Neural Networks for generating guidance commands in spacecraft landing scenarios has garnered recent research interest. A primary challenge lies in the inefficiency of training sample generation. This paper proposes a new approach to address it by utilizing nonlinear homotopy methods, which involve embedding the homotopic parameter either into the performance index or the right-hand side of the differential equations. Unlike the existing approaches that overlooked the potential value of the intermediate solutions during homotopy process, the proposed method incorporates these discarded solutions as supplementary training data, which treats the homotopic parameter and state as inputs, and the control as the output. Two examples, namely a minimum-fuel lunar landing and a minimum-time asteroid landing, are provided to illustrate the effectiveness and advantages of the proposed method. • New sample generation method for Deep Neural Networks in spacecraft landing guidance. • Fully utilization of intermediate solutions within the nonlinear homotopy process. • Drastic reduction in the number of optimization problems requiring solving. • Validating the proposed method in lunar and asteroid landing scenarios. [ABSTRACT FROM AUTHOR]

Fixed-wing aircraft are particularly attractive for Mars surface exploration, but Martian low-density atmosphere poses several issues to airfoil aerodynamic design. In this framework, the paper deals with an optimization procedure aimed at maximize the aerodynamic efficiency of Low-Reynolds airfoils. The design procedure embeds airfoil parameterization, aerodynamic computation, and optimization algorithm for a maximum efficiency objective at R e ∞ = 2 × 1 0 5 and R e ∞ = 3. 4 × 1 0 4 . Two airfoil parameterization models are developed and adopted within the design optimization. The XFoil solver is considered to address airfoil aerodynamics; while its optimal shape is generated with a genetic algorithm to simulate a cruising condition. Design optimization developed two different airfoils with (L / D) m a x = 10. 4 and 16.4, achieved at R e ∞ = 3. 4 × 1 0 4 . Aerodynamic performances of these wing sections are discussed in the work. Optimal airfoils pointed out that maximization of aerodynamic efficiency is associated to both a thickness reduction and a curvature increase of the aeroshape. Finally, aerodynamics of the optimal airfoils are also investigated with computational fluid dynamics simulations. • Issues of airfoil aerodynamic design posed by Martian low-density atmosphere. • Optimization procedure of airfoils efficiency for flight conditions on Mars. • Design optimization created two airfoils with (L / D) m a x = 10. 4 and 16.4, at R e ∞ = 3. 4 × 1 0 4 . • Airfoils aerodynamic performances addressed by means of CFD and panel method. • Optimal airfoils exhibited a low ratio of thickness to chord and a camber increase. [ABSTRACT FROM AUTHOR]

The Moon has returned into the focus of human endeavors regarding human spaceflight, e.g., with NASA's Artemis program, ESA's Moon Village, and the Russian/Chinese International Lunar Research Station. In difference to the pathfinding missions of the Apollo -era, the goal for these future missions is to stay on the lunar surface for longer durations and inhabit the lunar environment (near-)permanently. This requires a different approach to be affordable, i.e., instead of resupply as mostly used on e.g., the International Space Station, resource management has to include recycling and in-situ utilization. The former especially calls for the application of so-called BLSS to allow providing essential life-support services to the crew without prohibitive resource consumption, which is economically not feasible to achieve with resupplies. Bio-regenerative-life-support systems have been researched for decades, yet the system complexity, technology advancements, and singular aspects as e.g., plant biology require more research, especially if combined as in a greenhouse. For instance, the understanding of how a microbiome develops in a closed environment and what implications the microbiome has on plant growth is still insufficient. Within the EDEN project, the German Aerospace Center built a lunar analogue greenhouse and operated it at the Neumayer-III research station in Antarctica for four years, testing the technology – which was not space hardware – and operations. Derived from this experience the next step in the project is to design and subsequently operate a ground test demonstrator for a lunar greenhouse, as close as possible to the actual space hardware and operations. This paper explains the current design and trade-offs that led to it. Furthermore, the concept of operations is shown to illustrate the demonstrator's utility for researching bioregenerative-life-support. Overall, the system presented is feasible and useful to close the gaps, currently still existing in this field of research, and thus a mission enabler for future long-duration human space exploration missions. • Design of an integrated greenhouse module test demonstrator. • Extensive science program definition. • Mission scenario definition for lunar greenhouse. • Discussion of benefits and open issues. [ABSTRACT FROM AUTHOR]

Propellantless propulsion with a greater thrust-to-power ratio than photon rockets and functioning independently from the Sun or Earth-based lasers promise to revolutionize interstellar space travel. Claims of micronewtons in force for a Mach-Effect-Thruster on a torsion balance in vacuum, along with theories predicting mass fluctuations in a device with large fluctuations in internal energy, have paved the way for experimental investigations. In this experiment originally conceived by Woodward, devices are subjected to high centrifugal acceleration for direct measurement of mass fluctuations using an embedded force transducer. The test devices included a capacitor stack, a toroid and a pre-stressed, multi-layer piezoelectric stack in which electromechanical energy was varied. This paper first describes the design of a centrifugal balance: a rotating apparatus with low eccentricity and maximal angular frequency of 3600 rpm, capable of applying an acceleration above 1000 g's to the test device. The piezoelectric force transducer was characterized and calibrated under quasi-static conditions using an electrodynamic test instrument. The second harmonic response of the force transducer was then examined at different rotational rates to identify the presence of mass fluctuations. The experiment revealed a maximal resolution of 40 μg over a 100 kHz measurement bandwidth. Results with the electrostatic and inductive devices did not show mass fluctuations, orders of magnitude below the prediction. Even if the measurement resolution with the piezoelectric stack was more strongly affected by the centrifugal force, electromagnetic interaction, and nonlinear behavior, the measured noise reached one order of magnitude below the mass fluctuation prediction as well. Despite some limitations with the calibration and with the piezoelectric test device, the experiments have successfully demonstrated the feasibility of the centrifugal balance for high-frequency and high-precision mass variation measurements and have ruled out the postulated Mach-Effect mass fluctuations. • Development of novel thrust balance to investigate mass fluctuations in high frequency regime up to 100 kHz • Direct measurement of mass fluctuation claim for Mach-Effect thruster concept • Null-measurement sets new limits for theory [ABSTRACT FROM AUTHOR]

This paper focuses on investigating ablation characteristics of different nozzle materials in hybrid rocket motors by firing tests and numerical simulations. Hybrid rocket nozzles with different ablation materials are designed and manufactured. Four firing tests with working time of 40 s are conducted, and 90% hydrogen peroxide and polyethylene are adopted. The test results indicate that erosion rates of T705 graphite, whole felt C/C composite, axial rod braided C/C composite, and carbon/ceramic composite are 0.045 mm/s, 0.022 mm/s, 0.03 mm/s, and 0.01 mm/s. Circumferentially uneven ablation is found in the nozzle throat of T705 graphite and C/C composite, and this is mainly due to asymmetric oxidizer injection, asymmetric combustion, and uneven material properties. Transient numerical simulations of 40 s on nozzle thermochemical erosion are carried out, and the regression process of fuel wall and nozzle wall is solved by a dynamic grid technique. The error of average throat erosion rate between simulation results and test results is only 1.82%. The evolution process of nozzle profile and flow field structure are exhibited, and coupling effects between the erosion rate and nozzle profile are obtained. Simulation results show that as the working time increases, protrusions and pits appear on the nozzle inner wall and gradually extend. The temperature of nozzle wall and nozzle erosion rate on the windward side is higher, while they are lower on the leeward side. Nozzle throat erosion rate history is divided into four stages, and time nonlinear characteristics of the throat erosion rate are revealed. • 40 s firing tests of different nozzle materials are conducted. • Ablation properties of graphite, C/C composite and carbon ceramic are obtained. • Transient simulations of 40 s on nozzle thermochemical erosion are performed. • Simulation results agree with test data and evolution of nozzle profile is shown. • Time nonlinear characteristics of nozzle throat erosion rate are revealed. [ABSTRACT FROM AUTHOR]

This paper investigates the trajectory design for landing on a small celestial body with a flexible lander. The flexible lander features a flexible structure that increases surface contact area and facilitates the dissipation of residual kinetic energy. Compared with rigid landing, flexible landing, which utilizes a flexible lander to execute the landing process, offers enhanced safety during the landing operation. However, the introduced nonlinear flexible force will degrade the convergence of landing trajectory optimization. To address this challenge, the homotopic approach is employed to smoothly connect the trajectory optimization problem from rigid landing to flexible landing. Then, the connection between the optimal solution of the rigid problem and the flexible problem is revealed in detail with the presented theorem. Based on this theorem, the flexible landing fuel optimal trajectory is generated by iteratively solving a sequence of convexified and discretized homotopic problems. The effectiveness of the proposed algorithm is demonstrated through 433 Eros flexible landing mission based numerical simulations. • A novel convex programming method for flexible landing trajectory design is proposed. • The proposed method is proved to converge to the optimal solution of the flexible landing problem. • The optimal thrust magnitude profile for the flexible landing problem is of bang–bang structure. [ABSTRACT FROM AUTHOR]

The paper is aimed at developing numerical models for simulation of hypervelocity collisions of small fragments with honeycomb shields relevant to Space structures protection from impact of debris fragments. Numerical modelling of hypervelocity collision within hydrodynamic approach reveals effects that are inconsistent with physical sense. In order to overcome these inconsistencies and to achieve justified description of the process of penetration and fracturing numerical model was developed assuming the impactor and projectile being viscous materials. Despite constant viscosity we investigate two phenomenological relationships for viscosity: depending on temperature and depending on deformation rate. The exponential dependence of viscosity on material state (temperature and second invariant of strain rate deviator) makes the material manifest both properties of fluid and solid state of matter. Taking into account viscosity allows to suppress jets at symmetry axis and small-scaled ejecta. The usage of non-constant viscosity results in a more physically based scenarios for fracturing of the shell obtained in numerical simulations. • A new numerical models for simulation of hypervelocity collisions is suggested. • A model was developed assuming the impactor and projectile being viscous materials. • Relationships for viscosity depending on temperature and deformation rate were used. • The model makes the material manifest both properties of fluid and solid state. • Realistic scenarios for fracturing of the shell were obtained in simulations. [ABSTRACT FROM AUTHOR]

Supersonic and hypersonic vehicles have complex shock wave/boundary layer interaction problems during flight, and it adversely affect the aerodynamic performance of the vehicle. In this paper, the bump passive control approach is introduced to control the shock wave/boundary layer interaction, and its configuration is optimized with the minimum separation area and the maximum total pressure recovery coefficient. The Kriging model is established by means of the Latin hypercube experimental design approach to obtain the sample points. The Non-dominated Sorting Genetic Algorithm-II (NSGA-II) algorithm is utilized to optimize the parameters of the bump configuration, and the flow field structures of the uncontrolled condition, the special basic condition and the optimized condition are compared and analyzed. The obtained results show that the surrogate-based optimized model shows better performances than the basic condition. In the parameter setting range considered in this study, the optimized Case 3c with the middle position in the optimization front owns a larger total pressure recovery coefficient and a smaller separation area, with its separation area and total pressure recovery coefficient being 0.80 mm2 and 67.39% respectively. • Bump was used to control the shock wave/boundary layer interaction. • Bump configuration was optimized firstly by surrogate-based optimization algorithm. • An optimized case was obtained with smaller separation area and larger total pressure recovery coefficient. [ABSTRACT FROM AUTHOR]

This paper examines options for orbit configurations for a space interferometer. In contrast to previously presented concepts for space very long baseline interferometry, we propose a combination of regular and retrograde near-Earth circular orbits in order to achieve a faster filling of (u , v) coverage. With the rapid relative motion of the telescopes, it will be possible to quickly obtain high quality images of supermassive black holes. As a result of such an approach, it will be possible for the first time to conduct high quality studies of the supermassive black hole close surroundings in dynamics. • A new configuration for the space interferometer is proposed. • This configuration uses near-Earth regular and retrograde orbits. • Calculated orbits are stable more than 10 years. • The supermassive black hole can be imaged within 20 h. • Synthetic simulations show observations of black holes are possible in dynamics. [ABSTRACT FROM AUTHOR]

Satellite operators struggle to communicate with their satellites over Europe in the UHF amateur band (430–440 MHz) due to high power in-orbit interference. Statistical characterisation of the interference beyond average interference levels using in-orbit measurements is useful for the design of suitable countermeasures. Some recent studies have started to investigate both the frequency and time behaviour of the interference, but more measurements are needed to cover the whole UHF amateur band. In this paper, we use the Local Mean Envelope (LME) method to analyse the time and frequency characteristics of in-orbit radio interference from the LUME-1 satellite, measured in the 437 MHz band. Satellite measurements were performed on a TOTEM Software-Defined Radio (SDR), in orbit over Europe and Antarctica and the results were downloaded for analysis on the ground. The average power spectrum and the variability of the LME for different time windows in the 437 MHz band are presented. The data analysis also includes the coefficient of variation of the LME to study the dispersion of the interference. The results are compared to previous measurements in the 435 MHz band using the same method. The new results are found to differ compared with previous measurements on the 435 MHz band. The bandwidth of the band-limited signals measured at 435 MHz was found to be 300 kHz, and at 437 MHz, 200 kHz wide for this dataset. Also, the interference power was found to be higher for 437 MHz. In both cases the interference behaviour is highly non-Gaussian, suggesting that typical communication countermeasures will not be sufficient. Although more measurements are needed for better spatial resolution, our findings may contribute to explaining why satellite operations are difficult on these bands. A better knowledge of the time and frequency variability of the interference can indicate which mitigation techniques are required and are efficient to improve satellite communication in that band. • Average interference power and the spread in envelope was higher in the 437.28 MHz band than in the 435 MHz band. • A 200 kHz band-limited interference signal was detected at 437.28 MHz in contrast with the 300 kHz bandwidth of the signal at 435 MHz. • The interference behaviour is highly non-Gaussian. • Countermeasures assuming Additive White Gaussian Noise (AWGN) will not be realistic in this band. • Interference measured explains challenges experienced by operators of small satellites. [ABSTRACT FROM AUTHOR]

Narrowband SETI algorithms have been in use for many decades. A key processing step in the usual SETI pipeline called "De-Doppler" involves non-coherent integration of linearly-drifting lines in a spectrogram. De- Doppler (DD) is needed because a transmitted constant-frequency tone will be received as a chirp, or a tone with a near-constant frequency rate, due to relative accelerations between the transmitter and receiver. The frequency drift may be an important discriminator against certain types of radio-frequency interference. This paper examines the standard Taylor DD algorithm and finds there are significant signal losses in the integration process. Part of the loss is caused by imperfect indexing in the algorithm. Additional loss is associated with the effect of an FFT filter bank on chirp signals; DD performance is intertwined with the filter bank. Filter bank improvements involving polyphase filter banks are examined. A new DD algorithm called "fastDD" is proposed, as well as extensions to the Taylor algorithm. They both offer significant performance gains and higher resolution in frequency and drift rate in the DD detection plane. The loss mechanisms are now understood. Computational cost trade-offs for various filter bank and DD parameters are presented based on C code functions. Several "best" options are given which offer significant SNR improvements ranging from 2 to 4 dB for modest levels of additional computation. These options have the potential to dramatically improve detection rates. • Goal: (i) Understand and improve the De-Doppler (DD) function for narrowband SETI. (ii) Taylor DD – the gold standard for over a decade – has not been examined in detail. (iii) Need to look at DD as part of an overall end-to-end narrowband SETI detection system. • Results: (i) Significantly improved detection performance is possible: 2–4 dB improved sensitivity, depending on computation and memory budget; Potential SETI detection rate improvements of 2–4 times. (ii) Gains arise from possible use of overlapped polyphase filter banks and higher resolution in the DD plane. (iii) New algorithms have been defined (fastDD and extended Taylor DD) and evaluated. (iv) SNR Gain vs. Computation cost trades have been examined using C language functions: Best options of filter bank and DD parameters are identified. [ABSTRACT FROM AUTHOR]

This paper presents an innovative prescribed-time moving path following control strategy for a stratospheric airship subject to time-varying error constraints and external disturbances. With the help of a new time-varying constraining function, the strict constraint requirements on initial conditions of the system are relaxed. A novel error transformation mechanism is incorporated into the guidance method to tackle the error constraint. The constructed prescribed-time disturbance observer can provide estimates of unknown disturbances and model dynamic couplings within the predefined time. The proposed scheme guarantees that the predetermined precision requirement can be implemented within the prespecified settling time regardless of initial conditions. Numerical simulations demonstrate the effectiveness of the presented control strategy. • A prescribed-time MPF control method is proposed for a stratospheric airship with time-varying error constraints and disturbances. • The initial constraint requirements are removed by a new prespecified error constraint function. • A novel error transformation mechanism is integrated into the control algorithm to address the error constraint. • Estimates of model dynamic couplings and disturbances are obtained from the PTDO. [ABSTRACT FROM AUTHOR]

Estimating the pose of an uncooperative spacecraft is an important computer vision problem for enabling the deployment of automatic vision-based systems in orbit, with applications ranging from on-orbit servicing to space debris removal. Following the general trend in computer vision, more and more works have been focusing on leveraging Deep Learning (DL) methods to address this problem. However and despite promising research-stage results, major challenges preventing the use of such methods in real-life missions still stand in the way. In particular, the deployment of such computation-intensive algorithms is still under-investigated, while the performance drop when training on synthetic and testing on real images remains to mitigate. The primary goal of this survey is to describe the current DL-based methods for spacecraft pose estimation in a comprehensive manner. The secondary goal is to help define the limitations towards the effective deployment of DL-based spacecraft pose estimation solutions for reliable autonomous vision-based applications. To this end, the survey first summarises the existing algorithms according to two approaches: hybrid modular pipelines and direct end-to-end regression methods. A comparison of algorithms is presented not only in terms of pose accuracy but also with a focus on network architectures and models' sizes keeping potential deployment in mind. Then, current monocular spacecraft pose estimation datasets used to train and test these methods are discussed. The data generation methods: simulators and testbeds, the domain gap and the performance drop between synthetically generated and lab/space collected images and the potential solutions are also discussed. Finally, the paper presents open research questions and future directions in the field, drawing parallels with other computer vision applications. • Deep Learning is used extensively for monocular spacecraft pose estimation. • Deployability is an important limitation of the current DL-based solutions. • Lack of space-borne images during training leads to the domain gap problem. [ABSTRACT FROM AUTHOR]

Hall thrusters, the most common type of electric propulsion system, typically use xenon as the propellant, given its inertness, its ability to be stored at a high density under pressure, and good thrust to power ratio coupled with a high specific impulse compared to chemical propulsion. However, the number of satellites utilizing electric propulsion units and particularly Hall thrusters is dramatically increasing, resulting in a strain on the availability of xenon propellant in the context of a volatile noble gas market. This phenomenon is seen with the dawn of large satellite constellations and the accelerated launch rate of satellite units, the majority of which now use a Hall thruster as their primary propulsion system. Alternatives to xenon are available in the form of other noble gases, molecular propellants and condensable elements. Such propellants offer certain advantages in terms of specific mission scenarios, or for certain propulsion system sizes. This paper represents a review of alternatives to the conventional xenon propellant for Hall thrusters, providing a comparative study of the most feasible alternatives. Various considerations of using alternative propellants are outlined, and a comprehensive database of experimentally measured Hall thruster performance is compiled to pair the measured performance using various propellants to the results of a theoretical propellant performance estimation. • Condensable propellants may replace xenon in Hall thrusters. • Power requirements for condensable propellants estimated. • Iodine shown to offer similar performance to xenon at all power levels. • Molecular propellants: air, carbon dioxide and water performance in Hall thrusters. • Extensive Hall thruster performance database. [ABSTRACT FROM AUTHOR]

Nuclear Thermal Propulsion (NTP) systems hold promise in reducing transit times for exploration of the solar system by both crewed and uncrewed missions. NTP systems currently under investigation include a once-through high-temperature gas-cooled fission reactor to provide thermal energy to heat the coolant, which also serves as the propellant. The fuel system consists of angular UN fuel particles dispersed in a matrix of W/Re, creating a ceramic and metallic composite or cermet, which has been irradiated in Idaho National Laboratory's (INL) Transient Reactor Test Facility (TREAT). This enabled evaluation of these materials under representative nuclear heating rates (∼95 K/s) and peak surface temperatures (∼2527 K). These surface conditions were indicative of achievement of the target peak internal temperature of approximately 2850 K. These tests named Sirius-1 (UN–W/Re), have been irradiated and this paper will present post-irradiation examination results. The Sirius-1 test produced minor, stable cracks in the fuel specimen and spalling of surface material. Volatilized uranium 'soot' was found deposited on the inner wall of the irradiation capsule indicating loss of some fissile material from the fuel specimen. Spalling from the surfaces was also noted upon visual inspection. Uranium diffusion from the fuel particles resulted in the formation of U/Re phases and the production of a laminar microstructure on the edge of the fuel system. Overall, the specimen was stable and remained within a coolable geometry upon conclusion of multiple thermal cycles to prototypical operating temperatures. • A cermet nuclear fuel of UN particles in a W/Re matrix was irradiated and examined. • The sample material is a candidate for a nuclear thermal propulsion system. • The irradiation used prototypic heating rates and peak temperatures. • Post examination revealed that the fuel specimen remained largely intact. • Some localized zones of intermetallic W–Re sigma phase existed. [ABSTRACT FROM AUTHOR]

Obtaining planetary soil samples and returning them to the Earth is the most effective solution to detect the distribution and reserves of planetary resources. In order to maintain the stratification and a high coring ratio of soil samples during the sampling process, China National Space Administration adopted a flexible tube coring (FTC) method for the upcoming Chang'e-6 mission. In this paper, the lunar regolith FTC characteristics, coupled with its spiral debris discharge were modeled. Orthogonal experiments were carried out with different drilling parameters using the prepared high-density lunar regolith simulant. The results show that the coring height and coring ratio are greatly determined by the penetration per revolution (P P R) index. The lunar regolith simulant's coring ratio was as high as 90% when the P P R was optimized. The physical and mechanical parameters of artificial lunar regolith were characterized with respect to the P P R index as well as the dynamic flowing of the internal core, with a resulting accurately tuned coring ratio prediction model. Furthermore, the drilling load analysis can be applied to the optimization of drilling parameters employed by Chang'e-6 when collecting lunar regolith. • FTC method analysis to maintain the information of lunar regolith bedding. • Modeling and test methods for coring ground flow characteristics. • Analysis of core disturbance due to shear dilation under flow field coupling. • PPR has significant influence on coring height and ratio. [ABSTRACT FROM AUTHOR]

Thermal analysis of spacecraft systems is a critical process for mission operations. However, knowing the temperature distribution of entire systems is not easy due to the uncertainty of the thermal mathematical model (TMM) and limited temperature sensors. This paper proposes a temperature estimation method using physics-informed machine learning (PIML). The PIML-based thermal analysis allows us to estimate the actual temperature distribution by seamlessly bridging the limited observations and the TMM. To evaluate the estimation accuracy of the proposed method, we conducted a numerical experiment using a pseudo small satellite model consisting of 100 nodes. The proposed method was applied to three different model error cases and was found to improve temperature estimation accuracy in all cases. In addition, the impact of the number of temperature sensors and their placement on estimation accuracy was investigated. • A data-driven thermal analysis that can bridge the uncertain model and limited measurements is proposed. • The physics-informed machine learning was applied to the thermal mathematical model. • A numerical experiment was conducted using a pseudo small satellite model. • The effect of the number and placements of sensors on the estimation accuracy for three model error scenarios was discussed. [ABSTRACT FROM AUTHOR]

The vibration suppression problem of the large space truss structure during on-orbit assembly process is investigated in this paper, and a distributed adaptive neural network control approach is proposed to solve this problem. The dynamic modeling process of large space truss structure assembly is firstly developed. It can be used to describe the dynamic characteristics of the incrementally increasing space structure. To achieve the vibration suppression during on-orbit assembly, a distributed control approach is then proposed. The controller is designed for each assembly substructure, and a coordination term is added into the controller to reduce the interactions between the assembled substructures, and the distributed control system is therefore formed along with the assembly proceeds. The distributed controller is then redesigned using neural networks, to compensate the effect of the modeling uncertainties and the system parameter variation in the assembly process. It also could be noted that the dependence of controller design on dynamic models is then reduced. The simulation results demonstrate that the vibration is effectively suppressed, and the dynamic performance of the closed-loop system is improved. The controller designed based on the proposed dynamic model has good expansibility, and the control system also has robustness to external disturbance, and good fault tolerance when controller fails. • A dynamic model of the LSTS during on-orbit assembly is proposed. • A DANNC, to achieve the vibration suppression during on-orbit assembly, is designed. • The proposed controller can improve dynamic performance of closed-loop system. • Model uncertainty and parameter variation are compensated by the neural network. • The closed-loop system has good robustness and fault-tolerant performance. [ABSTRACT FROM AUTHOR]

Astronauts are at risk for pneumothorax, a condition where injury or disease introduces air between the chest wall and the lungs (i.e., the pleural cavity). In a worst-case scenario, it can rapidly lead to a fatality if left unmanaged and will require prompt treatment in situ if developed during spaceflight. Chest tube insertion is the definitive treatment for pneumothorax, but it requires a high level of skill and frequent practice for safe use. Physician astronauts may struggle to maintain this skill on medium- and long-duration exploration-class missions, and it is inappropriate for pure just-in-time learning or skill refreshment paradigms. This paper proposes semi-automating tool insertion to reduce the risk of complications in austere environments and describes preliminary experiments providing initial validation of an intelligent prototype system. Specifically, we showcase and analyse motion and force recordings from a sensorized percutaneous access needle inserted repeatedly into an ex vivo tissue phantom, along with relevant physiological data simultaneously recorded from the operator. When coupled with minimal just-in-time training and/or augmented reality guidance, the proposed system may enable non-expert operators to safely perform emergency chest tube insertion without the use of ground resources. • System design of a desktop testbed that facilitates data-driven event labelling. • Use of a lightweight robot enables semi-automation during simulated surgical tasks. • Needle-tissue interaction forces are useful for surgical semi-automation. • Robotics and wearables allow for human-in-the-loop system performance monitoring. [ABSTRACT FROM AUTHOR]

Nomadic worlds, i.e., objects not gravitationally bound to any star(s), are of great interest to planetary science and astrobiology. They have garnered attention recently due to constraints derived from microlensing surveys and the recent discovery of interstellar planetesimals. In this paper, we roughly estimate the prevalence of nomadic worlds with radii of 100 km ≲ R ≲ 1 0 4 km. The cumulative number density n > > R appears to follow a heuristic power law given by n > ∝ R − 3 . Therefore, smaller objects are probably much more numerous than larger rocky nomadic planets, and statistically more likely to have members relatively close to the inner Solar system. Our results suggest that tens to hundreds of planet-sized nomadic worlds might populate the spherical volume centered on Earth and circumscribed by Proxima Centauri, and may thus comprise closer interstellar targets than any planets bound to stars. For the first time, we systematically analyze the feasibility of exploring these unbounded objects via deep space missions. We investigate what near-future propulsion systems could allow us to reach nomadic worlds of radius > R in a 50-year flight timescale. Objects with R ∼ 100 km are within the purview of multiple propulsion methods such as electric sails, laser electric propulsion, and solar sails. In contrast, nomadic worlds with R ≳ 1000 km are accessible by laser sails (and perhaps nuclear fusion), thereby underscoring their vast potential for deep space exploration. • Cumulative number density of nomadic worlds estimated as function of the size range. • Detectability and distance of nearest nomadic worlds in given size range calculated. • Novel class of deep space missions to nomadic worlds delineated. • Utility of light sails and nuclear fusion for flybys of nomadic worlds underscored. [ABSTRACT FROM AUTHOR]

Round combustors as a very desirable kind of supersonic combustor, face the difficulty of combustion organization. The typical cone-strut structure can accomplish some combustion organization, but it has a very high congestion ratio and is hence inefficient. To solve this problem, a method of combining cone-strut configuration and plasma/oxygen addition is proposed in this paper, and its scheme, experiment and combustion mechanism are studied. In the scheme of combustion organization, the fuel injection scheme and plasma arrangement scheme of the approach are projected, accomplished with the advantages of reducing congestion ratio from 20%-26%–13%. In the experiments of combustion organization, the effect of realizing combustion organization of such approach is proved, and three typical combustion organization states, which may be due to low congestion ratio, are displayed one by one, and its transformation speeds between three typical combustion organization states are investigated. In the speculations of mechanism, a set of simplified equations are provided for the first time, and these equations indicate that why the approach is effective and why the three kinds of combustion organization states display one by one. This research has theoretical and technical implications for the development of cone-strut configuration in round combustors. • Provide a combustion organization method of cone-strut/plasma torch with oxygen feedstock for round supersonic combustor. • Experiments confirm feasibility of this method and point out the key stage that affect combustion organization. • The wakeflow equations including combustion reaction is used to explain the feasibility and particularity of the method. [ABSTRACT FROM AUTHOR]

In recent years, the Moon, as the closest celestial body to the Earth, has become the foremost target of human exploration outward. Some lunar exploration missions and studies have shown that ice may exist in the polar regions of the Moon. Water is very important for human survival. The ice mining and utilization is based on the understanding of the sublimation mechanism. Simulation of ground-based tests is an effective means to explore the sublimation of ice. Therefore, in this paper, the thermal environment of Shackleton Crater and the particle size distribution of typical lunar soils are analyzed as the basis of parameter settings for the ground test. Four operating conditions are established for ground tests, including the influence of heat sink temperature, vacuum level, particle size distribution and dry soil density. The larger the radiation heat flow, the higher the heat sink temperature, the smaller the vacuum degree, the larger the average particle size, resulting in a larger sublimation rate. The decrease of density makes the sublimation process faster and then slower. The sublimation process proceeds fastest when the dry soil density is 1200 kg/m3. The heat sink temperature has the greatest effect on the sublimation process, and the vacuum has the least effect. When the radiation heat flow ranges from 39.01 W/m2 to 45.63 W/m2 at the heat sink temperature of 120 K, the maximum sublimation rate is 5.79E-15 kg/s, the maximum sublimation area is 1.93E-9 m2 and the maximum sublimation heat flow is 1.59E-8 W. However, the effect of vacuum degree and particle size distribution on the sample temperature variation is not significant, due to the small order of magnitude of the sublimation heat flow compared to the radiation and conduction heat flows. This paper gives a range of parameters for ground-based simulation experiments and points out the microscopic changes of the ice sublimation process, which provides ideas for the next ground-based experiments and the exploitation of water-ice resources on the Moon. • A ground-based environmental test setup that is closely linked to the real lunar surface environment. • Analysis of the effects of temperature, vacuum and particle size distribution on the sublimation process of water ice. • Calculation of sublimation area, sublimation percentage and sublimation heat flow. • Energy analysis of sublimation process. [ABSTRACT FROM AUTHOR]

Visual instance segmentation ability is one of the effective means to promote the autonomy and intelligence of space agents. However, due to the limited airborne computing capacity, current methods are difficult to deploy on space agents, because these methods are developed based on GPU. To this end, this paper proposes a novel framework of instance segmentation by introducing depth information and combining traditional computer vision techniques with an object detection method. This framework provides a new idea for the implementation of instance segmentation. The experiment results show that the proposed method achieves a real-time performance under a common laptop CPU platform. In addition, thanks to the introduction of depth information, the proposed method can obtain better segmentation results compared to Mask R–CNN and SOLOv2 in complex scenes (poor illumination and occlusion). Finally, because the semantic information is obtained by the object detection method in this paper, the model training adopts a weakly supervised manner from bounding-box annotations, which can reduce various costs of data labeling to a certain extent. • A depth-aided instance segmentation fuses mask extraction and object detection. • Real-time performance on consumer grade laptop CPU; friendly to airborne platforms. • Better segmentation results in complex scenes compared with Mask-RCNN and SOLOv2. • Training process adopts a weakly supervised manner from bounding-box annotations. [ABSTRACT FROM AUTHOR]

Lunar sun-synchronous roving is an affordable and promising manner of acquiring continuous power supply and warmth in robotic polar exploration. Sun-synchronous route planning is challenging to meet the spatiotemporal constraints of lunar topography and time-varying illumination. Moreover, in pursuit of spaceflight feasibility and safety, the safety-optimal planning method adapting to spatiotemporal maps remains to be addressed. This paper takes the first step to investigate this issue, in which the distribution of roving speed is considered a critical factor. Three-dimensional lunar spatiotemporal maps are constructed based on topography and illumination analysis. A spatiotemporal planning algorithm named MSST-A* is proposed, which takes a multi-speed node expansion mechanism, thus supporting speed distribution optimization. Afterward, route evaluation functions involving motion, energy, and integrated risk, together with the optimal heuristic function, are designed for safety-optimal planning. Two comparative tests are conducted, and results indicate that the proposed approach is practical and superior in terms of motion safety and energy reliability. Besides, an integrated-risk-minimal route circumnavigating the Shackleton Crater at the south pole in three lunar days is solved, demonstrating the existence of long-term lunar sun-synchronous routes with good engineering significance. • The paper is innovative in the following bullet points: • Safety-optimal lunar sun-synchronous route planning is firstly studied. • MSST-A* for optimizing the spatiotemporal distribution of speed is raised. • Polar sun-synchronous circumnavigating routes of fine safety are solved. • The existence of the sun-synchronous route of engineering reliability is proved. [ABSTRACT FROM AUTHOR]

This paper introduces a flexible framework to enable high-fidelity simulations of the fuel slosh dynamics for verification of the Guidance, Navigation and Control (GNC) algorithms. The sloshing phenomenon must be tackled during the control system design phase and the GNC software run and validated within appropriate simulation frameworks. Equivalent mechanical models can well approximate sloshing under specific assumptions, allowing the derivation of the multibody vehicle's equations of motion; however, while suitable for analysis and control synthesis purposes, this approach presents several limitations for implementation as a simulation model. In this work, the multibody plant embedding slosh dynamics is modeled by means of the DLR's Vertical Landing Vehicles Library (VLVLib) written using the object-oriented Modelica modeling language. The advantages of using Modelica are explored throughout the paper. Given that control synthesis, analyses and simulation campaigns are usually performed in the Matlab/Simulink environment, a core point of the proposed solution is to achieve a good synergy within the latter and the Modelica plant model with the least readaptation burden for the developer. The strategies to achieve a full integration and verification tool chain are illustrated. The potential of this approach is demonstrated via two different test cases: a lunar landing scenario of the ALINA spacecraft developed by PTS, and CALLISTO reusable rocket demonstrator. • A flexible framework to enable high-fidelity simulations of the fuel slosh dynamics for verification of the GNC algorithms is introduced. • The spacecraft model embedding the slosh dynamics is implemented exploiting Modelica language. • A GNC testing framework is obtained with the integration of the causal Simulink environment with the compiled Modelica model. • Two test cases are presented: the ALINA spacecraft lunar landing scenario and CALLISTO reusable Vertical-Takeoff–Vertical-Landing demonstrative mission. [ABSTRACT FROM AUTHOR]

Attention to the method of removing large space debris objects by impacting them with an ion beam has not waned over the past 10 years. The application of this method to the space debris removal from the protected area of the Geostationary Earth Orbit (GEO) seems to be the most effective. This is due to a low energy consumption for removing an object to a safe orbit with ΔV≃11 m/s, which allows the removal of several objects by a single spacecraft. A service spacecraft can be developed for this purpose based on the TelSat geostationary communication satellite technology. At the same time, one of the problematic issues remains that is the development of an injector of a weakly divergent ion beam. The ion beam divergence during injection into the vacuum occurs under the insuperable influence of the ambipolar electric field acting in the beam volume. At the same time, the beam behavior depends on the initial divergence angle stipulated by the conditions of beam formation by the electrostatic acceleration system. This paper considers the results of calculated and experimental studies for the peculiarities of the formation of an ion beam in an ion extraction system with slit and round apertures. The dependence of the initial divergence half-angle on the beam perveance for both geometries was defined. In the operation of the injector designed for the space debris removal, it is necessary to ensure stability of the geometry of the ion-extraction system, which is subject to uneven thermal loading. The resulting thermomechanical deformations of the grids in the technology of ion thrusters are parried by giving the grids a plate shape. Therefore, flat grids made of a carbon-carbon composite material of a new type were used in experiments, which avoided noticeable deformations. The results of experimental investigation of the prototype of ion injector with two configurations of carbon grids (with round and slit perforation) and with the initial divergence half-angle of 2°-4° are given in this paper. The thrust control of an ion beam by changing of the screen grid potential under the condition of a constant beam perveance is considered. It is shown that such regulation is associated with a change in the Mach number, which affects the beam divergence. As an example, the calculation of the dependence of the coefficient k on the distance to the removing object is given, which is considered to be a spent satellite SUPERBIRD-A. • The minimum angle of the ion beam divergence of the ion injector were achieved. • A comparison of ion injector with slit-type and round-type aperture perforation. • The value of coefficient of thrust transfer to the object to be removed. [ABSTRACT FROM AUTHOR]

This paper describes a technique for the sampling of condensed combustion products (CCPs) of metalized propellants. Experiments were carried out with model propellants containing 15.7% of a metallic fuels, ammonium perchlorate, and an inert binder under pressure levels of about 2, 4, and 8 MPa. We used microsized powders of aluminum, boron, aluminum diboride and aluminum dodecaboride as metallic fuels. For the initial fuels, we defined the granulometric composition and determined the content of the active (non-oxidized) metal by the cerimetric method. The quenching of the dispersed phase particles was performed near the burning surface by a co-current argon flow. The paper provides the burning rate of the propellants and CCPs characteristics – the particle size and morphological composition, and the combustion incompleteness of the metal fuels. We have identified the morphology features of the agglomerated particles of aluminum and boron. The paper also contains calculated data on the heat release efficiency of the tested propellants. The obtained experimental data provide a basis for the development and verification of theoretical models on the combustion of metalized propellants. For the studied propellant formulations, we have not identified the increase in the boron combustion completeness due to the adding of aluminum into the propellant composition. • Propellant, boron, aluminum boride, combustion completeness, cerimetric analysis. [ABSTRACT FROM AUTHOR]

Ensuring fire safety is one of the most important problems in the creation of habitable spacecraft. It is necessary to thoroughly study the issues of the reliability of the functioning of space systems, including those using chemical fuel. The paper studies experimentally the dynamics of an M-shaped premixed methane-air flame under normal and reversed gravity at various external acoustic excitation. The stable M-shaped flame is studied experimentally by means of high-speed video recording of flame chemoluminescence and PIV-methods. The paper shows that at low frequencies of acoustic excitation, the flame front remains practically undisturbed. At a 160 Hz frequency, the oscillatory motion is energetically more saturated than the main gas flow. With a further frequency increase, the main flow again becomes dominant in terms of energy, and the effect on the flame structure decreases. • Intense oscillations of the M-shaped flame occur when excited by acoustic oscillations of 160 Hz. • The intensity of forced oscillations rapidly decreases with an excitation frequency increase. • An external acoustics leads to a less regular flame front structure under reverse gravity. • The flame responds less intensely to external acoustic excitation under reverse gravity. [ABSTRACT FROM AUTHOR]

Safe operation of orbital platforms relies heavily on stability of fluid flows in microgravity, which are strongly affected by capillary forces. This paper describes experiments on capillary imbibition at a station in Earth orbit, as well as during parabolic flights: firstly, experiments on repeated imbibition of an artificial porous medium (glass balls of various diameters) during parabolic flights of the Airbus A300-ZeroG aircraft organized by the European Space Agency; secondly, investigation of capillary driven seepage in soil (80% quartz sand and 20% kaolinite) on the Discovery STS-91 orbiter carried out within the framework of the MIRROR GAS program. The features of the experiments, experimental equipment, post-processing of experimental data are described. The paper considers the features of the fluid flow through an inhomogeneous porous medium, as well as the case when the medium is repeatedly imbibitted. This is possible if the liquid is observed during several successive parabolas during parabolic flight. The paper describes the mathematical model developed by the authors. The results of numerical simulation are compared with experimental data. • The features of the flow of liquids under microgravity conditions are considered. • Flows are studied controlling safe operation of spacecraft fluid supply systems. • Multiple imbibition and drainage are discussed. • The mathematical model that takes into account hysteresis is described. [ABSTRACT FROM AUTHOR]

The processes occurring in near and far space have a significant impact on the safety of space flights. Processes such as explosions of supernovae and relativistic jets have a particularly great impact. They arise in quasars, black holes, protostars. Often such structures are observed during the formation of new stars. Such jets have a great impact on the safety of space flights. For example, there is currently no radiation shelter on the International Space Station, so astronauts are at serious risk of getting radiation sickness. In addition, the skin of the orbital station modules consists of aluminum, which accumulates secondary radiation. The formation process of new star systems is the result of complex processes that occur in interstellar gas. These processes include nonlinear interactions of turbulence, gravity, collisions of molecular clouds with each other, as well as rotation and several other factors. The evolution of the formation of superdense substance begins from the moment when it gathers in turbulent flows or is formed by supersonic collisions of molecular clouds with each other, or are formed during the intereaction of molecular clouds with supersonic waves, which are formed during the explosions of supernovae, until the moment when these superdense areas reach the prestellar density. Further, depending on several factors, these superdense formations either collapse or disintegrate and return to the interstellar medium. The analysis of observations allows to conclude that a significant part of the molecular clouds is not gravitationally bound. These conclusions are made based on the virial theorem, which expresses the connection between gravitational and kinetic energy. In this paper, a simulation of large molecular clouds collision in a three-dimensional formulation on ultra-high-resolution grids is carried out. Execution of such calculations requires a lot of computer capacity. The paper presents the results of computer modeling of large-scale processes of filament formation and superdense, gravitationally coupled structurally parallel computation clusters with hybrid architecture. Parallel simulation on supercomputers was conducted with author's software, which uses a modified second-order Godunov method of accuracy of the Total Variation Diminishion type. Calculations were conducted on grids that contain more than one billion cells (1024 × 1024 × 1024). The gravitational potential is calculated on graphic processors. To refine the calculations, the method of adaptive refinement of the AMR grid is used. Simulation results are presented for cases of frontal collision of two molecular clouds, which density is distributed along the radius according to certain laws. • The processes of formation of new stellar systems are studied in this work in 3D simulation. • Appearance of clumps and filaments is discussed. • Morphing of high compression clumps set in possible stars formation zone is discussed. • The calculation program is based on a modified scheme of a high order of accuracy. [ABSTRACT FROM AUTHOR]

Preliminary aerodynamic design of a new generation high-speed transport is characterized by epistemic uncertainty in some initial data required for optimization parametric synthesis of aircraft design. In the present paper Liu's uncertainty theory was applied to develop high-speed transport preliminary aerodynamic design under epistemic uncertainty. The theory introduces axiomatics based on measured uncertainty (so-called expert confidence level) and contains analytical expressions for all functions that depend on epistemically uncertain parameters. Such approach provides high computational efficiency when performing optimization procedure compared to statistical modeling. Multicriteria optimization models of aircraft design parametric synthesis, obtained based on uncertainty theory, are presented in the paper. Analytical expressions for numerical characteristics of target functions and constraints with uncertain input and optimized parameters are derived. These ones provide analytical tools for correcting optimized parameters in the iterative high-speed transport designing process that cannot be based on statistical data. Effectiveness of the proposed models is evaluated in the context of the supersonic cruise flight optimization for high-speed transport under epistemic uncertainty and additional requirements for aircraft flight safety and environment low impact. • High-speed transport flight safety. • Multicriteria design optimization. • Epistemic uncertainty optimization models. [ABSTRACT FROM AUTHOR]

This paper presents scientific merits and design challenges associated with a novel idea to perform rocket sounding of the Martian atmosphere. Limited data sets of the Mars atmosphere profile are available, while ambitious exploration plans for the next years and decades require achieving game-changing capabilities in terms of extensive atmospheric data acquisition. Planetary landers, orbiters or rovers have limited capabilities regarding gathering data of the atmosphere. Aircraft and rotorcraft are among other solutions for Mars exploration, but they enable conducting soundings only at relatively low altitudes. It is therefore of uttermost importance to ensure the capability for further data acquisition supporting safe and sustainable exploration of Mars. This paper proposes the concept of use of Mars-dedicated sounding rockets. Unlike other concepts, considering rocket propulsion on vehicles capable of doing tasks including i.e. descent and landing, bringing samples of Martian soil to orbit and serving as a long-range Mars surface transportation system, the proposed concept is focused only on conducting ballistic flights with the main purpose of atmospheric research, aiding exploration and potential future activities. Rocket design drivers are discussed and compared with the state-of-the-art approach regarding Earth atmospheric sounding. Means for propelling suborbital flights on Mars are discussed. Optimization of vehicles for a reference mission is described and effects of sizing rockets for Martian application instead of Terrestrial utilization are given. Finally, recommendations for future dedicated Martian sounding rocket developments and a technical scenarios outlook are provided. • Mars atmospheric research history and its current exploration-related objectives. • Growing need of Martian atmospheric research using in-situ measurements. • Rocket payload mass fractions and propulsion system trade offs. • Performance of dedicated sounding vehicles on Mars and on Earth. • Mars mission concept utilizing dozens of sounding rockets within a single lander. [ABSTRACT FROM AUTHOR]

Gaussian Mixture Models and State Transition Tensors are combined to propagate state uncertainty to demonstrate an analytical and accurate future uncertainty and conjunction assessment. The initial uncertainty distribution is split into a Gaussian Mixture Model to reduce the size of covariance associated with each mixture component. This reduction in size allows for a significantly more accurate nonlinear mapping of each individual component. For higher accuracy in the mixture component propagation, State Transition Tensors are used to capture the nonlinear effects due to higher-order dynamics, as shown already in work by Fujimoto and Scheeres. In this paper, this GMM-STT method is used to analyze collisions using different dynamical models. The first case study demonstrates the application of this method in two-body problem dynamics; the second also adds J 2 secular effects; and the third brings in full J 2 and Solar Radiation Pressure dynamics through a Simplified Dynamical System. This paper also includes a two-body problem repeating conjunction example to establish a relationship between the propagation time and the number of mixture components needed for an accurate probability of collision computation. In each example, as the number of mixture components increases, the probability of collision converges with the Monte Carlo truth, proving the feasibility of this method. • Orbit Determination models make accurate uncertainty propagation essential. • Propagation with nonlinear dynamics leads to non-Gaussian future uncertainty. • Conjunction models require accurately propagated future state and uncertainty. • Gaussian Mixture Model - State Transition Tensor model gives accurate uncertainty. • Model accuracy is confirmed with comparison to Monte Carlo Probability of Collision. [ABSTRACT FROM AUTHOR]

The diffusion characteristics of liquid kerosene in a supersonic combustor equipped with a thin strut under the effect of evaporation and combustion heat-release were discussed in this paper. A series of experiments and numerical simulations were carried out both in cold inflow (T t = 300 K) and high enthalpy inflow (T t = 1680 K). The experimental data of kerosene diffusion and combustion were captured by high-speed camera and schlieren equipment. Based on the Euler-Lagrangian method, the evaporation and diffusion process of liquid kerosene under different inflow and heat-release conditions were studied. Through the analysis of the basic data of the ground test and numerical simulation, the kerosene diffusion and penetration depth were obtained and analyzed. The results showed that the diffusion of kerosene vapor is affected by both the injection momentum and the evaporation momentum. The evaporation of liquid kerosene could effectively improve the penetration depth of kerosene vapor, and the maximum proportion could reach 60%. The combustion heat-release greatly enhanced the diffusion ability of kerosene vapor by increasing evaporation strength of liquid kerosene and pressure in the heat-release zone, and the diffusion and combustion of kerosene were positively coupled with each other. With the investigation in this paper, a mechanism on kerosene diffusion in the supersonic combustor was explained. • Kerosene diffusion characteristics were analyzed with evaporation characteristics. • Evaporation momentum was defined and effect on kerosene diffusion was discussed. • The enhancement of kerosene diffusion with combustion heat-release was discussed. [ABSTRACT FROM AUTHOR]

With the deepening of space exploration, the moon has significant value for scientific research. However, because of such research's costs and risks, international cooperation will be critical for the future lunar exploration. In 2021, China and Russia jointly initiated the International Lunar Research Station (ILRS) project and released the Guide for Partnership (V1.0). However, the basic regulatory framework for creating the ILRS project is not yet in place, posing a challenge to pursuing this complex and technically ambitious endeavor. The paradigm of international cooperation that was the International Thermonuclear Experimental Reactor (ITER) project set a good example of how scientific research could transcend geographic boundaries. This paper examines the ITER Agreement, the constitutive agreement of the ITER, and analyzes the successful experience of large-scale international cooperation on a complex scientific project. Then, in view of the characteristics of the ILRS project and the existing space law, the main elements of a regulatory framework for international cooperation on this project are proposed, which follows the ITER Agreement closely. The paper concludes that such a regulatory framework for international cooperation on the ILRS project may contribute to the peaceful exploration and use of the moon, thus promoting the well-being and interests of humankind. • The moon has significant value for scientific research. • International cooperation will be critical for the future lunar exploration. • A regulatory framework for the International Lunar Research Station (ILRS) is not in place. • The paradigm of international cooperation on the ITER set a good example. • The main elements of a regulatory framework for the ILRS are proposed. [ABSTRACT FROM AUTHOR]

This paper presents some initial results on the use of hyperspectral imaging technology and machine learning to characterise the surface composition of space objects and reconstruct their attitude motion. The paper provides a preliminary demonstration that hyperspectral and multispectral analysis of the light absorbed, emitted and reflected by space objects can be used to identify, with some degree of accuracy, the materials composing their surface. The paper introduces a high-fidelity simulation model, developed to test this concept, and a validation of the model against experimental tests in a laboratory environment. The paper shows how to unmix the spectra to provide an estimation of the materials composing the surface facing the sensor. A machine learning approach is then proposed to reconstruct the attitude motion from the time series of spectra. • Demonstration of characterisation of space object composition using hyperspectral technology. • Demonstration of attitude motion reconstruction from spectra with machine learning. • High-fidelity model of the spectra received by a sensor. • Validation of the simulation model in a laboratory environment. [ABSTRACT FROM AUTHOR]

The development of deep space exploration has led to new exploration demands that cannot be met by traditional planetary rovers. This paper proposes a new wheeled rover that combines the advantages of passive all-wheel attachment and vibration reduction at high speeds. This rover can achieve vibration reduction in the vertical direction from ground excitation, preventing the instability of pitch and roll. Flexible wheels were used to reduce the high-frequency excitation of the ground on the rover. The design concept and mechanical principle of the proposed rover were comprehensively analyzed. A quasi-static kinematic model was developed to analyze the overstepping performance of the rover with this new design. Further, the response characteristics of the system to the ground excitation were investigated by vibration equations established using the Lagrangian method. The optimal parameters were obtained using a genetic algorithm to optimize the system with multiple objectives. Furthermore, the overstepping performance, all-wheel attachment, and vibration reduction of the planetary rover were verified through multi-body simulation. The proposed rover has considerable potential for missions such as the Mars sample return and planetary station construction. • A planetary rover with adaptive suspension for high-speed driving is proposed. • The theoretical analysis of the climbing ability of a planetary rover is provided. • A vibration modeling and evaluation method for proposed rover is established. • A variety of objective functions are proposed and optimized accordingly. [ABSTRACT FROM AUTHOR]

In this paper, the effects of the Tesla valve intake structure with or without bypass manifold on suppressing pressure feedback and combustion product backflow in rotating detonation engine fueled by kerosene and enriched air with 29% oxygen content are investigated, with the configuration of convergent–divergent intake without bypass manifold being used for comparison. The working range, detonation wave propagation characteristics, and propulsion performance for the three configurations are evaluated. It is found that the Tesla valve intake structure can effectively improve the suppression of pressure feedback and combustion product backflow through the diversion function of the Tesla valve bypass channel and that the addition of the bypass manifold further improves the suppression of pressure feedback. The Tesla valve intake with bypass manifold configuration has the best ability to suppress pressure feedback, with a minimum value of pressure feedback percentage of about 6.5% and a minimum value of air plenum pressure fluctuation percentage of about 6.7%. By comparison, the minimum value of pressure feedback percentage of the convergent–divergent intake without bypass manifold configuration is about 29.8%, and its minimum value of pressure feedback percentage is about 7.8%. The Tesla valve intake without bypass manifold configuration is most effective in suppressing combustion product backflow, with the effective inlet area ratio of about 95.7%. Under the same conditions, the minimum effective inlet area ratio of the convergent–divergent intake without bypass manifold configuration is about 90.2%. Ultimately, optimizing the engine structure leads to a broader working range and higher propulsion performance. The Tesla valve intake structure broadens the working range by suppressing the generation of the longitudinal pulsed detonation mode. When the bypass manifold is incorporated, some combustible reactants directly enter the rear of the combustion chamber from the head of the chamber, and the working range is reduced. Compared with the convergent–divergent intake without bypass manifold configuration, the working range of the Tesla valve intake with bypass manifold configuration is widened by 383.3%, and that of the without bypass manifold configuration by 700%. Both the Tesla valve intake and bypass manifold increase the strength of detonation waves and improve propulsion performance. The Tesla valve intake with bypass manifold configuration has the highest specific thrust, at about 90.1 N s3/(kg·m). This is increased by an average of 10% compared with the Tesla valve intake without bypass manifold configuration, and by an average of 30.1% compared with the convergent–divergent intake configuration. • Tesla valve inlet and bypass manifold are used to suppress feedback. • Two quantitative evaluation parameters of pressure feedback are established. • The degree of combustion product backflow is evaluated. • The optimized structure can effectively improve the propulsion performance. [ABSTRACT FROM AUTHOR]

This paper proposes a framework for a satellite mission with the goal to perform Lunar Occultations within the context of the circular restricted three-body problem (CR3BP) of the Earth-Moon system. This is achieved by exploiting the geometrical relationships between the Earth, Moon, and a spacecraft orbiting around L 4. Additionally, it examines the viability of a linear state-feedback control policy to stabilize short-period orbits under the presence of the gravitational effect of the Sun, Solar radiation pressure, and the Lunar orbital eccentricity, which is not accounted for in the CR3BP dynamics. Given the unprecedented NUV angular resolution achieved by a spacecraft at L 4 , this research proposes a framework for a remote sensing observatory for Solar activity, thereby enabling the monitoring of the entire hemisphere of Solar radiation, and to study different astrophysical issues via Lunar Occultations. • Lunar Occultations are studied by leveraging short-period orbits about L 4. • Short-period orbits (SPO) are selected to increase observation time of the Solar corona. • A linear controller is used to account for unmodeled orbital perturbations. • SPO with lower angular velocity is preferred to observe LO and Kordylewski clouds. • LO from L 4 offer unprecedented NUV angular resolution. [ABSTRACT FROM AUTHOR]

By slotting an inlet at the head of the aerodisk, a channel through the spike, and the lateral jet outlet at the middle of the spike, the blunt body with spike-aerodisk can further reduce its aeroheating and drag. In this paper, the numerical simulation method was used to analyze the flow field structure of this spike-aerodisk-channel, the effects of the spike length, the aerodisk diameter, and the channel diameter on the drag and aeroheating performance of this spike-aerodisk-channel configuration were studied. In general, the lateral jet introduced by the channel from the stagnation region of the aerodisk leads to the shear layer separation in advance, pushes the separation shock away from the spike, enlarges the recirculation zone near the blunt body, avoids the interaction between the separation shock and the reattachment shock, increases the distance from the reattachment shock to the surface of the blunt body. The reattachment shock strength of the blunt body is weakened. With increased spike length and channel diameter, drag and aeroheating reduction are improved. With the increase of the diameter of the aerodisk, the aeroheating performance is improved, but the drag reduction performance continues to decline. Compared with the plain spike-aerodisk, the spike-aerodisk-channel is generally conducive to further drag and aeroheating reduction. In the range of research parameters, considering the effect of drag and aeroheating reduction and the structural feasibility, the spike-aerodisk-channel has the best drag and aeroheating reduction performance when the spike length is twice the blunt-body diameter, the ratio of the aerodisk diameter to the blunt-body diameter is 0.3, and the ratio of the channel diameter to the blunt-body diameter is 0.075. Compared with the plain spike-aerodisk of the same size, the aeroheating reduction efficiency is increased by 37.9%, and the drag reduction efficiency is increased by 16.0%. • Drag and aeroheating reduction mechanism of the spike-aerodisk and channel concept under hypersonic conditions was investigated by the numerical simulation. • Three influence factors were analyzed and compared comprehensively. • Combined spike-aerodisk and channel concept is a promising drag and aeroheating reduction technique. [ABSTRACT FROM AUTHOR]

The dynamic of hypervelocity projectile impacting inflatable multilayer flexible membrane structure is studied in this paper. The velocity increment of the inflatable multilayer flexible membrane structure under hypervelocity impact comes from two parts. One is the acceleration caused by the pulling of the membrane when the hypervelocity projectile is in direct contact with the membrane, and the other is the acceleration provided by the reverse impulse caused by the gas leakage in the inflatable structure. The hypervelocity projectile will be decelerated in the impact process, and finally fly with the accelerated multilayer flexible membrane structure at the same velocity, which means that the projectile is captured by the membrane structure. According to the above process, the influence of the mechanical property parameters of the flexible membrane on the structural parameters of the multi-layer flexible membrane can be calculated, which can be used to guide the engineering design. Finally, the hypervelocity impact process is verified by the numerical simulation. • Space soft capture method achieved through flexible inflatable membrane structure. • Hypervelocity impact dynamics of flexible inflatable membrane structure. • Verify the dynamic process of hypervelocity impact using numerical simulation. • The influence of membrane mechanical properties on the parameters of the structure. [ABSTRACT FROM AUTHOR]

This paper presents a new analytical procedure to model the umbra generated during an eclipse considering an oblate ellipsoid of rotation as occulting body and a conical shadow. The method is based on purely geometrical considerations and results in the analytical definition of the entry and exit points from the conical shadow starting from the knowledge of the Sun position vector, the occulting body position vector and the orbital elements of the spacecraft orbiting the occulting body. The conical shadow also permits analytical definition of the entry and exit points of the penumbra region, which cannot be defined by using the classic cylindrical approach. Some numerical applications are proposed to test the effectiveness of the analytical formulations and to check the error in the prediction of the time spent in the shadow by the satellite. Finally, a discussion between the new conical shadow model and the classic cylindrical eclipse is carried out to see the improvements introduced by the refined geometry and the effects on space missions focusing on the cumulative error when multiple revolutions are considered. • A new analytical method to compute umbra entry/exit points in an eclipse is derived. • The Earth's surface is modelled as an oblate ellipsoid of rotation. • The model permits the definition of the penumbra region and its entry and exit points. • The method is a higher-order refined model without loss of computational time. • Umbra/penumbra regions are modelled as the Cartesian equation of a conical surface. [ABSTRACT FROM AUTHOR]

The Common Heritage of Mankind (CHM) Principle is a political outcome from the interest game between spacefaring countries and non-spacefaring countries. The space CHM principle per se is an existence of international law and is a historical, dynamic, open and developing one. The CHM principle in the Moon Agreement should be interpreted with a strictly semantic interpretation standard ("no more, no less") in lex lata. The legal character of the celestial bodies, its natural resources in place and resources abstracted should be defined. The basic elements and connotation of the CHM principle in the Moon Agreement could be drawn from the analysis of the articles about non-appropriation, non-property and international regime for space resources exploitation. In respect of the prospects of the development of space resources exploitation, the paper comparatively analyzed res communes , res publicae , common property and condominium and subsequently propose the due common recognition of the legal concepts and principles of the CHM in space domain. • Common recognition of the CHM benefits future space activities regime. • Interpreting the CHM with "no more, no less" principle in lex lata. • The general space CHM principle lex ferenda is historical, dynamic, open and developing. • Differentiating general space CHM principle from res communes , res publicae , common property and condominium. [ABSTRACT FROM AUTHOR]

Space habitats, facing extreme conditions in space and on other planetary bodies, should have redundancy and versatility. Although the main challenges in the field have been identified as changes in gravity, radiation protection, extreme temperatures (fluctuations), changes in daylight patterns, and lack of (or less) atmosphere, more (unknown) challenges will inevitably emerge. Therefore, space habitats should extend beyond redundancy and adapt to their surroundings. There are many adaptability strategies for terrestrial habitats but none of these strategies were developed to work in the extreme conditions of outer space. This paper proposes application of a novel adaptability strategy, the Rhythmic Buildings strategy, which aims to develop buildings that adapt to the rhythms of the building's context (frequency, speed, and intensity of changes occurring on the case study location). The strategy is applied to the case study of a Martian habitat at Jezero Crater. The Rhythmic Buildings strategy includes the Rhythmic Framework and its 33 parameters (such as daylight and outdoor temperature) as well as analytical, technical, and design tools – collectively the Rhythmic Toolbox. First, the Framework tool was used to map the changes in the context following the environmental, economy, and society aspects. The rhythms of the case study context include local temperature, pressure, solar radiation, strong weather events (i.e. storms), but also comfort needs of the crew. These rhythms were then translated into a habitat design using an adapted version of the bubble diagram method. Lastly, novel technologies and materials were selected to address the rhythms which includes adaptive properties of materials. Results of the study show that the context's most important challenges are the daily temperature fluctuations and weekly crew schedule, while the most promising adaptability opportunity lies in the daily daylight rhythm. The design that followed the Rhythmic Buildings strategy directly addresses nine out of the 33 parameters and indirectly addresses a further five parameters. The proposed Martian habitat utilises the rhythms of daily temperature cycles and the 24 h daylight rhythm to reinforce the materials structures. The habitat is adaptive and responsive to its surroundings and the crew's needs. In the discussion, speculations are made on how space architecture can develop itself by intrinsically adapting to the rhythms of the environment and evolve into their own typologies, distinctly different than Earth architecture. • Introduces an adaptability strategy for Space Architecture called Rhythmic Buildings. • Using a novel analysis and design Toolbox to design a Martian habitat. • Design of a habitat that responds to the rhythms of the context. • Outlines criteria for the use of responsive materials and ISRU. [ABSTRACT FROM AUTHOR]

This publication examines the history of Project Orion and NERVA nuclear engine program within the Integrated Program Plan, the technical reasons for their study, their progress, and the political reasons for their cancellation. Whether pulsed or thermal, nuclear propulsion offers the best compromise in efficiency and thrust for the propulsion of future manned missions. Project Orion was born in the mind of Stanislaw Ulam, an atomic physicist who had worked on the Manhattan project. Despite theoretical performances superior to anything conceivable even with current technologies, and despite the mobilisation of some of the most brilliant physicists of the time (Freeman Dyson), the project was killed by the Treaty banning nuclear weapon tests in the atmosphere, in outer space and under water (Moscow, 5 August 1963). The Integrated Program Plan was Wernher von Braun's proposal to the Space Task Group (STG), chaired by Vice President Spiro Agnew, for the continuation of the Apollo program. This plan envisaged a lunar base in the 1970s, and a first manned mission to Mars in the early 1980s. All of this would have been powered by nuclear-powered orbital shuttles, using NERVA, a nuclear engine tested and operational since 1968. Here again, even though it performed better than anything else in the world, even today, the project failed politically. This retrospective publication therefore aims to answer the following question: How can nuclear space programs of today and tomorrow be protected from the uncertainties of this influence? The study of propulsion systems is very often left to engineers, and political and public relations issues to politicians. However, every technology is the result of a political system, and has a strategic impact, which is a subject for social and political sciences. That's why this paper adopts an original approach crossing social sciences through the historical study of decision-making dynamics, and engineering through the study of the technical characteristics of the projects involved. This publication is based on declassified technical reports, testimonies of the personalities of the time, and a document from the archives of the John F. Kennedy Presidential Library. History can be the academic bridge between political and technical studies. In a context of resurgence of nuclear propulsion projects throughout the world, such an approach is now necessary both at the historical level and in terms of political decisions making. Indeed, space programs involving nuclear energy are, by nature, at the crossroads of public relations issues very specific to both space and atomic subjects. [ABSTRACT FROM AUTHOR]

The vast majority of Earth-orbiting satellites carry an electrical power subsystem (EPS) which main components are solar panels and secondary batteries. During eclipse periods, satellites are powered only by rechargeable batteries which have a large energy density but a limited power density. This fact limits the power capabilities of small satellites during eclipse periods. Due to the large power density of Supercapacitors (SCs), ground and on-orbit tests have been conducted to verify their applicability on satellite EPSs. At the moment, no studies are underway on the issue of sizing SCs for eclipse operations. Hybrid configuration could reduce the mass and volume of EPS or maintain those reference values but increasing peak power capabilities. This paper deals with this issue. On the one hand, novel analytical expressions of a variable capacitance SCs are derived, including time–voltage dependence in constant power applications. Also an equivalent formulation for a SCs bank (SCsB) is derived. On the other hand, a SCsB sizing procedure is presented, considering the energy and power requirements of a particular space mission as an input. Finally, a simple mission is analysed, the results showing an improvement on the hybrid EPS design with respect to the traditional, being the mass reduction of 36%. • Implementation of a simple but accurate supercapacitor (SC) model for spacecraft applications. • Innovative analytical expressions of the performance of a SC in constant power applications. • Novel determination of the electrical parameters of a SCs bank, treated as an equivalent SC model. • Estimation of key variables in the operation of hybrid storage systems for space applications. • Preliminary design of a SCs bank devoted to supply high power during eclipse times. [ABSTRACT FROM AUTHOR]

Imitation learning leverages example demonstrations to teach a policy to replicate a desired behavior. In this investigation, the example demonstrations consist of open-loop optimal trajectories calculated off-line. This paper introduces the use of imitation learning to demonstrate optimal feedback control of two different high-fidelity 6-degree-of-freedom (6DOF) lander models. A loss of optimality is shown when disturbances are applied to policies trained only with nominal trajectories. Methodologies such as multi-phase optimal control and triple-single-phase optimal control are applied to include disturbances in the optimal trajectory generation, and the policies are trained to mitigate loss of optimality when disturbances are applied. Monte Carlo simulations show that loss of optimality can be mitigated by including disturbances in the optimal trajectory generation and training data. [Display omitted] • Imitation learning for optimal feedback controllers for the pinpoint landing problem. • Disturbances can cause loss of optimality for policies trained in nominal dynamics. • Including disturbances in optimal trajectories mitigates loss of optimality. [ABSTRACT FROM AUTHOR]

The first interstellar object to be discovered, 1I/'Oumuamua, exhibited various unusual properties as it was tracked on its passage through the inner solar system in 2017/2018. In terms of the potential scientific return, a spacecraft mission to intercept and study it in situ would be invaluable. As an extension to previous Project Lyra studies, this paper elaborates an alternative mission to 1I/'Oumuamua, this time also requiring a Jupiter Oberth Manoeuvre (JOM) to accelerate the spacecraft towards its destination. The difference is in the combination of planetary flybys exploited to get to Jupiter, which includes a Mars encounter before proceeding to Jupiter. The trajectory identified is inferior to previous finds in terms of higher Δ V requirement (15.6 km s−1), longer flight duration (29 years) and less mission preparation time (launch 2026), however it benefits from a feature absent from previous JOM candidates, in that there is little or no Δ V en route to Jupiter (i.e. a free ride) which means the spacecraft need not carry a liquid propellant stage. This is marginally offset by the higher Δ V needed at Jupiter, requiring either 2 or 3 staged solid rocket motors. As an example, a Falcon Heavy Expendable with a CASTOR 30B booster followed by a STAR 48B can deliver 102kg to 1I/'Oumuamua by the year 2059. Other scenarios with shorter flight durations and higher payload masses are possible. • A Mission to Interstellar Object, 1I/'Oumuamua is studied. • Chemical/High Impulsive Thrust Propulsion is assumed. • A VEEMJ trajectory to Jupiter is adopted as well as a Jupiter Oberth Manoeuvre. • A solar slingshot (Solar Oberth) is NOT utilized thus obviating the necessity for a solar shield. [ABSTRACT FROM AUTHOR]