Target Satellite Research Articles

Mission Planning Method for Dense Area Target Observation Based on Clustering Agile Satellites

To address the mission planning challenge for agile satellites in dense point target observation, a clustering strategy based on an ant colony algorithm and a heuristic simulated genetic annealing optimization algorithm are proposed. First, the imaging observation process of agile satellites is analyzed, and an improved ant colony algorithm is employed to optimize the clustering of observation tasks, enabling the satellites to complete more observation tasks efficiently with a more stable attitude. Second, to solve for the optimal group target observation sequence and achieve higher total observation benefits, a task planning model based on multi-target observation benefits and attitude maneuver energy consumption is established, considering the visible time windows of targets and the time constraints between adjacent targets. To overcome the drawbacks of traditional simulated annealing and genetic algorithms, which are prone to local optimal solution and a slow convergence speed, a novel Simulated Genetic Annealing Algorithm is designed while optimizing the sum of target observation weights and yaw angles while also accounting for factors such as target visibility windows and satellite attitude transition times between targets. Ultimately, the feasibility and efficiency of the proposed algorithm are substantiated by comparing its performance against traditional heuristic optimization algorithms using a dataset comprising large-scale dense ground targets.

Development and research testing of a low thrust on gaseous-propellant rocket engine chamber

The article presents preliminary results of the development and research testing of a rocket engine chamber with a thrust of 100 N operating on fuel components: gaseous oxygen-gaseous hydrogen, designed to be used as the main engine of a small upper stage for launching a payload weighing up to 150 kg into target orbits. The main features of the engine chamber are its fabrication by selective laser melting from 12X18N10T steel powder and its regenerative cooling with gaseous hydrogen. Calculation and experiments confirmed the possibility of regenerative cooling of the chamber in the nominal mode. In addition, the current results of the development of a technique for optical recording of the process of structural material removal from the chamber during tests using different optical filters are presented. It is shown that there is a correlation between the brightness of the obtained frames and the hydrogen flow rate. It is also shown that the afterburning of the removed material particles, in contrast to high thrust engines, occurs mainly in the tail of the jet.

Simulation case study of precise orbit determination for IGSO satellites using onboard Ka/GNSS observations

Abstract Inter-satellite links can provide a more accurate space reference for IGSO satellites than onboard GNSS side-lobe signals. In this study, POD quality is analyzed based on onboard Ka/GNSS observations by utilizing a simulated IGSO satellite orbit. Considering the similar orbit altitude, we first select three IGSO satellites from BDS-3 as target satellites. These satellites have actual Ka-band measurements available, along with other BDS-3 satellites and ground anchor stations, allowing us to assess the POD performance. The Ka-band availability of these IGSO satellites can serve as a means to evaluate the maximum utilization of BDS Ka-band resources. In terms of the POD results, the root mean square (RMS) of the Ka-band residuals can reach 10.9 cm, and the orbit consistency in the 3D direction is better than 0.35 m. In addition, simulations are conducted for onboard GNSS, inter-satellite link (ISL), and ground anchor-satellite link (GSL), considering the different constraints on the Ka-band resources. With only GNSS measurements, the 3D orbit accuracy is approximately 1.5 m, but when additional ISL and GSL measurements are considered, the POD accuracy can surpass 0.28 m. These results illustrate that incorporating Ka-band measurements can effectively enhance the POD accuracy for high-orbit satellites.

A General On-Orbit Absolute Radiometric Calibration Method Compatible with Multiple Imaging Conditions

On-orbit absolute radiometric calibration is not only a prerequisite for the quantitative application of optical remote sensing satellite data but also a key step in ensuring the accuracy and reliability of satellite observation data. Due to the diversity of imaging conditions for optical remote sensing satellite sensors, on-orbit absolute radiometric calibration usually requires a large number of imaging tasks and manual labor to calibrate each imaging condition. This seriously limits the timeliness of on-orbit absolute radiometric calibration and is also an urgent problem to be solved in the context of the explosive growth of satellite numbers. Based on this, we propose a general on-orbit absolute radiometric calibration method compatible with multiple imaging conditions. Firstly, we use a large amount of laboratory radiometric calibration data to explore the mathematical relationship between imaging conditions (row transfer time, integration level and gain), radiance, and DN, and successfully build an imaging condition compatibility model. Secondly, we combine the imaging condition compatibility model with cross calibration to achieve a general on-orbit absolute radiometric calibration method. We use cross calibration to obtain the reference radiance and corresponding DN of the target satellites, which calculates the general coefficient by using row transfer time, integration level, and gain, and use the general coefficient to calibrate all imaging conditions. Finally, we use multiple imaging tasks of the JL1GF03D11 satellites to verify the effectiveness of the proposed method. The experiments show that the average relative difference was reduced to 2.79% and the RMSE was reduced to 1.51, compared with the laboratory radiometric calibration method. In addition, we also verify the generality of the proposed method by using 10 satellites of the Jilin-1 GF03D series. The experiment shows that the goodness of fit of the general coefficient is all greater than 95%, and the average relative difference between the reference radiance and the calibrated radiance of the proposed method is 2.46%, with an RMSE of 1.67. To sum up, by using the proposed method, all imaging conditions of optical remote sensing satellite sensor can be calibrated in one imaging task, which greatly improves the timeliness and accuracy of on-orbit absolute radiometric calibration.

Concept of a Green Propulsion System for Ioshex, Designed to Perform In-Orbit Randezvous and Docking

ABSTRACT Recent challenges in space access include cost optimization and space debris reduction. The European Space Agency has envisioned the evolution of in-orbit transportation involving re-usable service vehicles, moving payloads from high parking orbits to their target orbits. This concept for orbital infrastructure requires de-risking activities and the development of building blocks, such as standardized interfaces and communication. Mastery of close-proximity operations and docking is essential for the new transportation system. This new branch of space operations represents a good opportunity to introduce more sustainable propulsion solutions, especially given the high costs and uncertain future of hydrazine and its derivatives. IOSHEX, a service spacecraft to be equipped with a green propulsion system utilizing 98% hydrogen peroxide, serves as a reference. This paper presents a design concept for this propulsion system, including trade-off analyses, calculations, and a three-dimensional model integrated with IOSHEX.

Test Method for Single Satellite’s Inter-Satellite Link Pointing and Tracking via Ground Station

An inter-satellite link is a key technology that improves control accuracy, transmission efficiency, and autonomous capability of constellations. A satellite’s pointing and tracking abilities mainly determine the inter-satellite link’s performance, which should be validated through an in-orbit test. However, during the construction of the constellation, the distribution of satellites does not satisfy the constraints of establishing the inter-satellite link. A test method for inter-satellite link pointing and tracking is developed with respect to a single satellite. A practical mission scenario for testing inter-satellite links’ performance is constructed. A virtual satellite is introduced as the target satellite to establish an inter-satellite link with the local satellite. The orbit of the virtual target satellite between two ground stations is characterized based on the Newton–Raphson method. By comparing the predicted and actual time differences between two ground stations receiving the signals from the local satellite, the inter-satellite link pointing and tracking abilities are evaluated independently. Numerical simulations verify the design of the virtual satellite. The single satellite test method for inter-satellite link pointing and tracking abilities is available.

Real-time fault-tolerant guidance for launch vehicle ascending flight under thrust drop failure

This paper presents a real-time fault-tolerant guidance method, which solves the autonomous rescue problem in the event of thrust drop failure during the final stage flight of launch vehicles. The core of the method is the online and optimal reconstruction of both the degraded target orbit and the subsequent flight trajectory. Aiming at achieving satisfied onboard computing performance and making the best use of the launcher’s residual energy, a unique relaxation-penalization model transformation method for the nonlinear orbit injection conditions is proposed, and a successive convexification algorithm is developed to solve the problem. Furthermore, both the two classic thrust drop failure situations, that is the mass flow rate drop and the specific impulse drop problems, can be resolved by the proposed algorithm. The average run-time of this orbit-trajectory joint optimization algorithm is below 150ms in an embedded ARM processor @1.5 GHz, which is adequate for onboard use. Using the above algorithm as a central infrastructure, a receding-horizon-strategy-based fault-tolerant guidance method is proposed. By performing the optimization repeatedly at a high frequency, the effects of parameter deviations and external disturbances can be absorbed for the most part, for example, the guidance algorithm is verified to be capable of tolerating at least ±5% thrust deviations. Comprehensive numerical and hardware-in-the-loop simulations are performed to demonstrate the computational performance, accuracy, and reliability of the proposed algorithms.

Study of the payload extraction trajectory heavy class carrier rocket

Relevance. As the weight and complexity of the payload that needs to be launched into orbit increases, the relevance of rational trajectory selection to ensure maximum efficiency and minimum costs for delivering the payload to a given orbit increases. Rational choice of the trajectory of a heavy-class launch vehicle has a number of important practical applications. Firstly, it allows you to increase the payload capacity of the launch vehicle and reduce the cost of delivering payload to the target orbit. This is especially important in the context of the development of the space industry, when more and more companies and organizations are showing interest in launching their own satellites and other spacecraft in conditions of fierce economic competition. Choosing a rational trajectory for launching a payload into orbit will significantly reduce the cost of launches and make them available to a wider range of potential customers. Secondly, the choice of launch vehicle trajectory parameters is important for ensuring safety and minimizing risks during spacecraft launches. Thanks to the rational choice of trajectory, it is possible to reduce adverse impacts on the environment and eliminate the possibility of emergency situations associated with loss of control over the flight of the launch vehicle. Rational selection of launch vehicle trajectory parameters is a complex task that requires comprehensive research and consideration of various factors, such as aerodynamic parameters of the atmosphere, mass and characteristics of the payload (spacecraft), engine operating parameters, characteristics of the target orbit, features of the launch of the launch vehicle and many other factors. A more thorough and systematic study of the influence of these parameters will significantly improve the efficiency and reliability of launching spacecraft into orbit. Thus, the choice of rational parameters for the launch vehicle trajectory is a relevant and important topic for scientific research. Increasing the rocket's payload capacity, reducing the cost of delivering a spacecraft to a given orbit, and ensuring launch safety are tasks that depend on the chosen shape and parameters of the rocket's trajectory. Such research has important practical significance and can become the basis for the development of new technologies and methods in the space industry. The purpose of the study is to study and select rational parameters for the trajectory of a heavy-class launch vehicle when launching a payload. The main task is to determine the flight path parameters that will allow achieving maximum efficiency and accuracy in delivering the payload to a given orbit. To achieve the goal of the study, an analysis of various factors influencing the launch parameters of the spacecraft is required, such as the structural and aerodynamic characteristics of the rocket, the influence of aerodynamic factors and the Earth’s gravitational field on the flight path. Taking these factors into account, numerical calculations were carried out on the basis of a system of differential equations of motion using a computer program created in the MAPLE software package. Based on calculations, modeling of the shape and parameters of the launch vehicle flight path was carried out. Research results. During the study, the rational parameters of the trajectory of a heavy-class launch vehicle were selected. Calculations were carried out using numerical modeling of the parameters of payload launch trajectories, and an analysis of the resulting trajectories was carried out. Minimizing the rocket's flight time was identified as the main criterion for the rational choice of trajectory, which allows increasing launch efficiency and saving energy resources. An increase in payload mass and minimization of fuel consumption were adopted as additional criteria. Conclusion. The procedure for choosing rational parameters for the trajectory of a heavy-class launch vehicle proposed in this work will improve the delivery accuracy and reliability of spacecraft launches at the stage of ballistic analysis when designing rockets. The results of the study have practical significance for the development of future heavy-duty launch vehicle missions and improving the efficiency of space launches.

Research on relative reachable domain in target orbit for maneuvering spacecraft

PurposeLots of successful space missions require that the maneuvering spacecraft can reach the target spacecraft. Therefore, research on relative reachable domain (RRD) in target orbit for maneuvering spacecraft is particularly important and is currently a hot-debated topic in the field of aerospace. This paper aims at analyzing and simulating the RRD in target orbit for maneuvering spacecrafts with a single fixed-magnitude impulse and continuous thrust, respectively, to provide a basis for analyzing the feasibility of spacecraft maneuvering missions and improving the design efficiency of spacecraft maneuvering missions.Design/methodology/approachBased on the kinematics model of relative motion, RRD in target orbit for maneuvering spacecraft with a single fixed-magnitude impulse can be calculated via analyzing the relationship between orbital elements, position vector and velocity vector of spacecrafts, and relevant studies are introduced to compare simulation results for the same case and validate the method proposed in the paper. With analysis of the dynamic model of relative motion, the calculation of RRD in target orbit for maneuvering spacecraft with continuous thrust can be transformed as the solution of the optimal control problem, and example emulations are carried out to validate the method.FindingsFor the case with a single fixed-magnitude impulse, simulation results show preliminarily that the method is in agreement with the method in Ref. (Wen et al., 2016), which treats the same case and thus is plausibly correct and feasible. For the case with continuous thrust, analysis and simulation results confirm the validity of the proposed method. The methods based on relative motion in this paper can efficiently determining the RRD in target orbit for maneuvering spacecraft.Originality/valueBoth theoretical analyses and simulation results indicate that the method proposed in this paper is comparatively simple but efficient for determine the RRD in target orbit for maneuvering spacecraft swiftly and precisely.

Lunar flyby transfers to TianQin configuration

The TianQin mission for gravitational wave detection consists of three drag-free spacecrafts deployed in geocentric orbits, and forms an equilateral triangle configuration. In this paper, we investigate lunar flyby transfers to the TianQin configuration. Firstly, a preliminary design method of lunar flyby transfers is proposed in the patched-conic model. Then, lunar flyby transfers in the ephemeris model are constructed through the multiple shooting method. Numerical computation shows that compared to the traditional Hohmann transfer, the lunar flyby transfer can save about 12% of fuel, but its time of flight increases to about 11 days. For the initial deployment of the TianQin configuration, we propose two methods, the phasing method and the direct insertion method. The phasing method uses the phase adjustment when the three spacecraft reach the target orbit. Numerical calculation indicates that the fuel costs of the three spacecraft are identical to that of the nominal transfer, and the total time of flight is about 17 days. Then, the Jet transport tool is applied to find appropriate separation impulses to design the direct insertion orbits. Numerical computation shows that compared to the phasing method, the fuel cost of the direct insertion method is slightly larger than that of the phasing method, but the time of flight of the former is about 5 days shorter than that of the latter.

Adaptive track approach for multiple sources scenarios during radar survey for space surveillance applications

The increasing population of resident space objects is currently fostering many space surveillance initiatives, which rely on the use of on-ground sensors. In particular, survey radars allow to first characterise the target orbit from a single transit, through measurements which are Doppler shift, slant range and angular profile. In this framework, the Music Approach for Track Estimate and Refinement (MATER) algorithm was developed to compute the angular track through the MUltiple SIgnal Classification (MUSIC), by solving possible ambiguities which may arise because of the receiver array geometry.This work presents the MATER extension to the case in which multiple sources are simultaneously detected by the sensor. For each detected source, the signal Direction of Arrival (DOA) is computed through MUSIC, and, if no prediction is exploited, possible ambiguous solutions arise. The computation is repeated for the entire observation, and all the estimations related to a specific target are grouped based on the angular sequence shape and the detection epochs. Finally, the possible ambiguity problem is solved, and the angular profile is obtained through a quadratic regression in time.The algorithm is numerically tested on a survey observation simulation. The detection length depends on the impinging signal intensity, and the angular accuracy is in the order of 1e-03 deg. A sensitivity analysis highlights that a transmitted power decrease shortens the detection length, with no remarkable angular accuracy deterioration. An additional simulation shows that in proximity operation monitoring MATER performance depends on the mutual geometry between target and chaser, as it may bring down the reciprocal angular distance under the resolution level. Nevertheless, it is always possible to identify the presence of both sources through the eigenvalues analysis of the signal correlation matrix. Finally, the simulation of a fragments cloud observation highlights that MATER performance depends both on the size of the observed fragment, as this is strictly linked to the signal detected by the receiver, and on the simultaneous detection of other fragments with a more effective signal.

GEO satellite on-orbit refueling and debris removal hybrid mission planning under uncertainty

This article proposes two optimization models to solve the multi-satellite on-orbit hybrid service mission planning problem under uncertainty. During the process of on-orbit refueling and debris removal, two categories of uncertainties are taken into account. These uncertainties have an impact on the mass of the service satellite as well as its fuel consumption during the orbit transfer. The two-stage orbit transfer strategy is given to save fuel for maneuvering, which also determines the directed graph between target orbits of satellites or debris. The stochastic expectation model is to get the service sequence with the optimal average benefit, while the robust model considers the worst situation and gives the optimal minimum benefit. Simulations of GEO satellites in orbit are conducted to demonstrate the impact of various models on handling uncertain tasks.

Discrete-Event Drag Modulation Aerocapture for Mars and Titan Missions

Aerocapture is a spaceflight maneuver that uses the atmosphere of a planet or moon to slow down a spacecraft after a single atmospheric pass and capture it into an elliptical orbit around the celestial body. This paper presents a discrete-event drag modulation method for Mars and Titan aerocapture missions. A single-stage jettison is used as a control strategy to modulate the aerodynamic drag for an aerocapture scenario. Entry corridor analysis is performed under worst-case trajectory dispersions for determining the entry flight-path angle. The jettison timing of the high-drag skirt is the single free guidance parameter that adjusts the spacecraft’s trajectory during an aerocapture maneuver. A computationally simple jettison-switching guidance scheme is developed to determine when to jettison the high-drag skirt, and the result is an analytical function that represents the ballistic-coefficient switching curve. The proposed discrete-event drag-modulation strategy is extensively tested using four entry vehicles and three target orbits for Mars and Titan aerocapture scenarios. Monte Carlo simulations show that the simple switching-curve guidance provides performance metrics (apoapsis targeting errors and postexit impulsive maneuvers) that are essentially equal to the performance of more computationally burdensome guidance methods such as an onboard numerical predictor–corrector scheme.

The Impact of GCP Chip Distribution on Kompsat-3A RPC Bias Compensation

The vast potential of high-resolution satellite images, including Kompsat-3A, has been demonstrated across diverse applications, such as mapping and disaster monitoring. However, these images can only be utilized as reliable GIS (geographic information system) data when they possess precise geographical information. To achieve this, sensor model information, represented by RPCs (rational polynomial coefficients), requires bias compensation through GCPs (ground control points). Though having a substantial number of well-distributed GCPs across satellite images is ideal, the acquisition process is often restricted due to cost and inaccessibility. The uniform distribution of GCP chips is not guaranteed, necessitating an investigation into the impact of GCP distribution on the bias compensation process, which is the focus of this study. Experiments were meticulously conducted using Kompsat-3A data using dense GCP information. The dense GCP information was automatically generated from aerial orthoimages through a three-step process. Firstly, the GCP chips were extracted from the aerial images, focusing on feature points. Secondly, these chips were projected onto the target Kompsat-3A data to align them accurately. Lastly, precise satellite image coordinates of the chips were obtained through image matching between the chips and the target Kompsat-3A image. The dense GCPs enabled detailed bias analysis that exhibited skewness in most Kompsat-3A data. This necessitates the implementation of an affine model for proper bias compensation over the entire image space. Next, the study delved into the influence of GCP distribution on RPC bias compensation. To this end, each target satellite image space was divided into nine zones, with the dense GCPs assigned accordingly. The accuracy of bias compensation was analyzed across nine experimental cases, ranging from GCPs occupying only one zone to GCPs covering all nine zones. It was observed that GCPs covering at least four or five zones should be utilized for reliable RPC bias compensation of Kompsat-3A, especially when aiming for a high level of accuracy with an RMSE of one pixel. Finally, it was concluded that GCPs covering three zones yielded satisfactory results as a minimum GCP requirement, but this was contingent upon their distribution not following a straight zone pattern.

Density-based in-orbit collision risk model valid for any impact geometry

Neglecting small fragments in space debris evolutionary models can lead to a significant underestimation of the collision risk for operational satellites. However, when scaling down to the millimetre range, the debris population grows to over a hundred million objects, making deterministic approaches too computationally expensive. On the contrary, probabilistic models provide a more efficient alternative, which however typically works under some simplifying assumptions on the dynamics, limiting their field of applicability. This work proposes an extension of the density-based collision risk models to any orbital dynamics and impact geometry. The impact rate with a target satellite is derived from a multi-dimensional phase space density function in orbital elements, which discretely varies over both phase space and time. The assumption of a bin-wise constant cloud density allows for the analytical transformation of the six-dimensional distribution in orbital elements into the three-dimensional spatial density function, guaranteeing an efficient and accurate evaluation of the fragments flux. The proposed method is applied to the assessment of the collision risk posed by occurred fragmentation events in different orbital regions on a high-risk target object. The effect on the impact rate of the modelling improvements, compared to previous probabilistic formulations, is discussed.

CORBES: Radiation belt survey with international small satellite constellation

The COnstellation of Radiation BElt Survey program (CORBES) is designed to deploy small satellites into a highly elliptic orbit for multi-point exploration of the Earth’s radiation belts. Its scientific objective is to achieve unprecedented high-time-resolution dynamics measurements within the regions of Earth’s outer radiation belts. The CORBES program initiative comprises satellites equipped with three types of payloads: the Magnetometer (MAG), the Search Coil Wave Detector (SCWD), and the High Energy Electron Detector (HEED). The energy interval of HEED is suggested as 0.1–4 MeV, logarithmically divided into 12 channels. To ensure extensive coverage of the outer radiation belts, a highly eccentric and inclined orbit is suggested, featuring a perigee of 280 km, an apogee of 7 Earth-Radius (Re), and an inclination of approximately 11°, resulting in an orbital period of approximately 13.5 h. Within a single orbital period, it takes roughly 10 h to traverse the outer radiation belts (3 Re to 7Re). All satellites are expected to operate within the same orbit, maintaining a spin-stabilized with sun-pointing spinning axis, and a spinning speed of approximately 8 RPM. Each satellite’s mass should not exceed 30 kg. For telecommand, either S-band or X-band will be utilized, while X-band is designated for data downlink. The satellites are scheduled for launch by one or two rockets, with the equipped upper stage placing them into the target orbit, and the attached dispenser releasing them individually according to the required separation sequence. Key aspects of the program include cross-calibration, radiation shielding, assembly integration and testing (AIT). Prior to launch, the cross-calibration is optional for the payloads. The payloads will be tested in the same environment to calibrate the technical specifications. Post-launch, in orbit cross-calibration becomes necessary to maintain data consistency and comparability. Specifically for HEED, this involves selecting electrons with the same energy range during the magnetospheric quiet period (Kp < 3), and comparing the observation results of different HEEDs under the same L,B conditions. A similar method applies to MAG and SCWD comparing observations during selected quiet period. Given that the satellites will operate within radiation belts characterized by high-energy protons at low altitudes and electrons at high altitudes, all on-board electronic components must meet fundamental requirements, including shielding geometry structure design, and thickness calculation to mitigate the Total Ionizing Dose Effect (TID) to a level of 200 krad [Si] over a one-year mission cycle. Lastly, system-level AIT before launch could be performed.

Transformer Network-Aided Relative Pose Estimation for Non-cooperative Spacecraft Using Vision Sensor

The objective of the proposed work is to perform monocular vision-based relative 6-DOF pose estimation of the non-cooperative target spacecraft relative to the chaser satellite in rendezvous operations. In this work, the convolutional neural network (CNN) is replaced by the high-resolution transformer network to predict the feature points of the target satellite. The self-attention mechanism inside the transformer provides the advantage of overcoming the inadequacies of the translation equivariance, 2D neighborhood awareness, and long-range dependencies in CNN. First, the 3D model of the target satellite is reconstructed using the inverse direct linear transform (IDLT) method. Then, the pose estimation pipeline is developed with a learning-based image-processing subsystem and geometric optimization of the pose solver. The image-processing subsystem performs target localization using CNN-based architecture. Then, the key points detection network performs regression to predict 2D key points using the transformer-based network. Afterward, the predicted key points based on their confidence scores are projected onto the corresponding 3D points, and the pose value is computed using the efficient perspective-n-point method. The pose is refined using the non-linear iterative Gauss–Newton method. The proposed architecture is trained and tested on the spacecraft pose estimation dataset and it shows superior accuracy both in translation and rotation values. The architecture has shown robustness against the drastically changing clutter background and light conditions in the space images due to the self-attention mechanism. Moreover, this method consumes less computation resources by using fewer floating-point operations and trainable parameters with low input image resolution.

Assessment of small satellite aerocapture with a morphable entry system at Mars

A significant challenge for interplanetary small satellites is fully propulsive orbit insertion due to limited total velocity change capabilities. At destinations with significant atmospheres, this challenge can be circumvented via aerocapture, a technique that uses a single atmospheric pass to convert a hyperbolic approach trajectory into a captured elliptical orbit. This paper investigated small satellite aerocapture at Mars with the morphable entry system, a deployable entry vehicle concept that uses shape morphing for trajectory control. A Monte Carlo dispersion analysis was conducted to assess the vehicle’s robustness with respect to uncertainties in entry state, atmospheric density, and vehicle aerodynamics. Under the combined effects of these uncertainties, the morphable entry system achieved a 100% aerocapture success rate. Furthermore, the total velocity change required to place the vehicle on the target orbit for 99% of the simulated trajectories was 166.5 m/s, which is within the current capabilities of small satellite propulsion systems. The results suggest that the morphable entry system is a viable option for smallsat aerocapture at Mars.

Orbit Transfer Design and Analysis for the SMILE Satellite

Solar Wind Magnetosphere Ionosphere Link Explorer (SMILE) aims to understand the interaction between the solar wind and the Earth’s magnetosphere. The initial orbit of the SMILE satellite is a low Earth orbit, while the target orbit is a highly elliptical orbit. This paper presents the design and analysis of orbit transfer for the SMILE satellite with the constraints of ground stations. According to the maneuver constraints, design principles and strategies have been determined. The orbit transfer strategy is to raise the apogee first and then the perigee, and no active control is carried out for the argument of perigee and the right accession of ascending node. Besides balancing the number of orbit maneuvers and the fuel loss, the orbit transfer is decomposed into 11 maneuvers. To increase the solvability, the idea of multistep approaching the target has been adopted. After a three-step design, the orbit transfer scheme satisfying the observational constraints of the maneuvers by the ground stations has been achieved. Results reveal that the fuel consumed during the orbit transfer is estimated to be 1440 kg, within the fuel budget. Furthermore, the orbital lifetime is calculated to be 7.08 years, which can fully meet the mission’s requirements.

Feasibility study of a Solar Electric Propulsion mission to Mars

Traditionally, space missions to outer planets have relied on large amounts of chemical propellant or flybys to achieve their objectives. This large mass consumption may be significantly reduced using electric propulsion (EP), since it allows for larger specific impulses. However, this improvement is counteracted by the lower thrust per power levels achievable with EP engines compared to chemical thrusters. Here we discuss how to find a balance for a purely solar EP (SEP) uncrewed mission to Mars. We present system engineering and mission analyses focused on the orbital mechanics. The result is the conceptual design of a realistic mission to Mars, in which any target orbit around Mars can be reached using solely present SEP technology. In particular, for a 2000 kg spacecraft propelled with a Hall effect thruster (HET), the interplanetary transfer can be performed in 363 days with less than 400 kg of fuel consumption, provided that the vehicle leaves the Earth orbit with a suitable specific energy that is attainable using current launchers. Subsequently, 300 kg of propellant are sufficient to allow for the planetary capture, plane change, and circularisation manoeuvres, finally inserting the spacecraft into a polar orbit of height between 300 km and 1000 km above the surface of Mars.