Reentry Point Research Articles

Design of Entire-Flight Pinpoint Return Trajectory for Lunar DRO via Deep Neural Network

Lunar DRO pinpoint return is the final stage of manned deep space exploration via a lunar DRO station. A re-entry capsule suffers from complicated dynamic and thermal effects during an entire flight. The optimization of the lunar DRO return trajectory exhibits strong non-linearity. To obtain a global optimal return trajectory, an entire-flight lunar DRO pinpoint return model including a Moon–Earth transfer stage and an Earth atmosphere re-entry stage is constructed. A re-entry point on the atmosphere boundary is introduced to connect these two stages. Then, an entire-flight global optimization framework for lunar DRO pinpoint return is developed. The design of the entire-flight return trajectory is simplified as the optimization of the re-entry point. Moreover, to further improve the design efficiency, a rapid landing point prediction method for the Earth re-entry is developed based on a deep neural network. This predicting network maps the re-entry point in the atmosphere and the landing point on Earth with respect to optimal control re-entry trajectories. Numerical simulations validate the optimization accuracy and efficiency of the proposed methods. The entire-flight return trajectory achieves a high accuracy of the landing point and low fuel consumption.

Shooting stars on schedule: synchronising re-entry of particles launched from a satellite

The ALE-3 mission plans to create artificial ‘shooting stars’ in the upper atmosphere by launching pellets sequentially from a satellite in low Earth orbit. How does one arrange for all pellets to re-enter over the intended location simultaneously? Starting with conservation of energy and angular momentum, I derive a version of Kepler’s equation that gives time as a function of radial distance in an orbit, which can be used to find the transfer time from the pellet’s ejection to its re-entry as a function of its launch velocity. I show that for a given pellet ejection speed, there is a launch angle from the satellite that results in the fastest transfer time. I use these results to determine the pellets’ launch times and velocities for simultaneous arrival at a desired re-entry point. These results can be applied to de-orbiting any set of objects launched from a satellite, and use concepts that can be covered in an advanced undergraduate course in physics or aerospace engineering.

Developing community-based criteria for sustaining non-timber forest products: A case study with the Missanabie Cree First Nation

Non-timber forest products (NTFPs) are crucial to the spiritual and cultural identity of Indigenous communities, while playing a significant role in traditional economies. In Canada, current forest policies and commercial management systems do not adequately protect and conserve these natural resources. Decisions makers have yet to fully and effectively use local and traditional knowledge in forest policy and planning. This research used a participatory approach and 26 community interviews to document local criteria, elements and critical local values surrounding NTFP development and conservation within the traditional territory of the Missanabie Cree First Nation. Applied ethnobotany and policy analysis were used to develop a locally specified framework and case study to guide decision making. The analysis establishes key community-based objectives to promote diverse socio-economic benefits, while advancing local forest management systems. In particular, the co-research process identified six criteria (1. productive capacity; 2. health; 3. conservation and protection; 4. long-term management benefits; 5. community participation; 6. capacity of institutions) and associated 18 elements and 239 critical local values, which span the maintenance and enhancement of NTFPs. The resulting framework represents one possible re-entry point into the forest planning process for Missanabie Cree who are working to reconnect with their land following a long period of forced exclusion. Moving forward, leadership can continue to revise this framework to accommodate changing needs, further engage their membership, and in doing so refine their roles in forest governance. The research process and framework documented here can inform other forest land managers looking to create more inclusive decision making processes and document local values.

Non-locality of ion reflection at the shock front: dependence on the shock angle

In typical heliospheric collisionless shocks most of the mass, momentum and energy are carried by ions. Therefore, the shock structure should be most affected by ions. With the increase of the Mach number, ion reflection becomes more and more important, and reflected ions participate in shaping the shock profile. Ion reflection at the collisionless shock is a non-local process: the reflected–transmitted ions re-enter the shock front far from the reflection point. The direction and the magnitude of this shift depend on the shock angle. The distance between the reflection point and the re-entry point is of the order of the upstream ion convective gyroradius and exceeds the shock width. The non-locality of ion reflection may have implications for shock rippling since reflected ions may carry perturbations along the shock front.

Practical use of dual-lumen catheter-facilitated reverse wire technique for nonhighly angulated side branch

Dual-lumen catheter (DLC)-facilitated reverse wire technique is considered a method of last resort for inserting a guidewire into a markedly angulated side branch. Moreover, this technique can be practically applied to other types of anatomical variations around the bifurcation. Case 1 was that of a 53-year-old man with a tight stenosis of the proximal left anterior descending artery at the diagonal bifurcation with angiographically apparent coronary dissection. We successfully achieved guidewire insertion into the targeted branch using the DLC-facilitated reverse wire technique. Case 2 involved a 78-year-old man with total occlusion of the mid-portion of the right coronary artery. The guidewire reentry point in the atrioventricular branch was slightly distant from the true distal end of the occlusion. We successfully used the DLC-facilitated reverse wire technique to pass a second guidewire to the posterodescending artery. Case 3 was that of an 80-year-old man whose coronary artery had an aneurysm with severe stenoses at both entry and exit of the aneurysm. We adopted the DLC-facilitated reverse wire technique and easily advanced the guidewire by matching the guidewire advancing path with the direction of the sequential conduit constituted by the coronary aneurysm and stenosis. The timing of decision-making to try the DLC-facilitated reverse wire technique is important. The lesion for which reverse wiring is suitable is usually difficult to be treated with conventional guidewire crossing. We have to promptly judge the validity of applying this technique based on the angiographic findings of targeted lesions and take immediate action to implement this procedure to reduce the procedural time and irradiation dose.

Integrated Design of Moon-to-Earth Transfer Trajectory Considering Re-Entry Constraints

The exploration of the Moon has always been a hot topic. The determination of the Moon-to-Earth transfer opportunity and the design of the precise transfer trajectory play important roles in manned Moon exploration missions. It is still a difficult problem to determine the Moon-to-Earth return opportunity for accurate atmospheric re-entry and landing, through which the actual return trajectory can be easily obtained later. This paper proposes an efficient integrated design method for Moon-to-Earth window searching and precise trajectory optimization considering the constraints of Earth re-entry and landing. First, an analytical geometry-based method is proposed to determine the state of the re-entry point according to the landing field and re-entry constraints to ensure accurate landing. Next, the transfer window is determined with the perilune heights, which are acquired by inversely integrating the re-entry state under the simplified dynamics as criterion. Then, the precise Moon-to-Earth trajectory is quickly obtained by a three-impulse correction. Simulations show the accuracy and efficiency of the proposed method compared with methods such as the patched-conic method and provide an explicit reference for future Moon exploration missions.

Precise re-entry and landing of propellantless spacecraft

• Propellant-free spacecraft ground impact point targeting algorithm. • Guidance and feedback control targeting strategy using exclusively aerodynamic drag. • Monte Carlo simulation results proving algorithm robustness and performance. Spacecraft returning scientific samples or humans from space must be capable of surviving re-entry and landing in a desired location. Traditionally, this has been accomplished via a propulsive de-orbit burn. However, it is not always possible to mount a propulsion system on board a satellite or a capsule. In the case of small satellites deployed from the International Space Station, for example, on-board propulsion systems are forbidden for safety reasons. Our work proposes a new technological solution for re-entering and landing a spacecraft in a desired location from a low Earth orbit using exclusively aerodynamic drag and eliminating the need for chemical propulsion. First, an iterative procedure is utilized to compute the desired state at the re-entry interface (100 km) such that a propagation of the vehicle dynamics in the nominal re-entry drag configuration from this initial state leads to a landing at a desired latitude and longitude on the surface of the Earth. Next, a re-entry point targeting algorithm is utilized to determine the on-orbit ballistic coefficient profile necessary to target the desired re-entry point. Finally, the ballistic coefficient profile during the final hours of the trajectory before the re-entry interface is iteratively modified to correct any remaining along-track error in the landing location. The proposed solution is applied to a small satellite system that is jettisoned from the ISS and is equipped with a deployable heat shield that also serves as a drag device.

I Stink, Therefore I Mink: A Manifesto

The recent mass culling of mink in Denmark and elsewhere, following the animals’ contamination by a COVID-19 variant, is taken as a re-entry point into Derrida and Lacan’s mink-mediated conversation in The Beast and the Sovereign. Out of the etymological ‘stink’ attached to the mink emerges an animot gifted with (unlimited) ink, with a potential to disturb philosophies of language, to write back or strike back, as it has recently done in the form of alignments of dead yet resurfacing animals. In the wake of Derrida’s verbal disseminations around the vison, and of Lacan’s attribution of a ‘sort of language’ to the animal in The Formations of the Unconscious, this essay follows an animal pack with includes the 17 million mink programmed for (double) extinction by inhumation and cremation. A hauntology follows, adumbrated by Lacan’s interest in the ‘secretion’ of fur, mink oil and (psychoanalytic) sense, and by Derrida’s encounter with the neoliberal, crypto-vison Alain Minc in 1994.

Preliminary Determination of Footprint Area of Uncontrolled Space Debris: Case Study of Tiangong-1 Space Station

Indonesia is an archipelagic country consisting of 16,056 islands and covering a vast area around 5,120km x 1,760km. With the largest coastline in the world, Indonesia is vulnerable to the fall of human-made objects from space. Furthermore, the space objects placed at polar and equatorial regions pass over the equatorial region, including Indonesia, more frequently around 4 and 9 times a day, successively depending on their altitudes. Due to the significant probability of the passages, determining the footprint of falling space objects (debris) is mandatory. Therefore, this study examines the demise of Tiangong 1 as a case study. First, trajectory propagation was carried out to track the re-entry point resulting in an estimated footprint area of around 2,632 km x 2,698 km over the Sothern Pacific Ocean. Second, a mathematical formulation in Astrodynamics was applied to engage a series of assumptions, which led to a more cramped footprint area of around 193km x 12km over a small portion of the South Pacific Ocean. Since the orbital prediction is fraught with great uncertainty, it was very likely that the Tiangong-1 debris fell over the Southern Pacific Ocean of the order of thousands of kilometers.

Terminological and Metrological Aspects in Siepser-Type Iris Suture Procedures

Instead of contemporary plethora, with tenths of different terms for only 7 elements, I suggest a system in which the main criterion is the length of the 2 filament ends just before the loop is externalized: L: long end, and S: short end. The name of the 2 corneal pathways, and 2 iris bites will rely on the name of the corresponding filament end that passes through, just before the loop is externalized: Sp/SP: short-end puncture/paracentesis, Lp/LP: long-end puncture/paracentesis, SIB: short-end iris bite, LIB: long-end iris bite. When the loop is externalized, the “intermediary segment”: (I) is created from L. The limits of (I) have been adjusted: instead of generally admitted limits from the middistance between SIB and LIB, to the middle of the externalized loop, I suggest that the S/(I) limit be situated at 1mm after the exit point of the loop from SP, and the (I)/L limit be situated at the re-entry point of the loop in SP. The suggested terminology and limits of (I) allow a clear and unequivocal description of techniques, simplify the repartition of functions per filament segment, avoid misunderstanding as cause of failure, facilitate the establishment of rules for success valid in any Siepser-type suture, and allow a new systematization of all iris suture procedures in a structure with a common stem, 2 branches (McCannel, Siepser) and leaves represented by technical variants.

The Fate of Anthropogenic Nanoparticles, nTiO2 and nCeO2, in Waste Water Treatment

Wastewater treatment is one of the main end-of-life scenarios, as well as a possible reentry point into the environment, for anthropogenic nanoparticles (NP). These can be released from consumer products such as sunscreen or antibacterial clothing, from health-related applications or from manufacturing processes such as the use of polishing materials (nCeO2) or paints (nTiO2). The use of NP has dramatically increased over recent years and initial studies have examined the possibility of toxic or environmentally hazardous effects of these particles, as well as their behavior when released. This study focuses on the fate of nTiO2 and nCeO2 during the wastewater treatment process using lab scale wastewater treatment systems to simulate the NP mass flow in the wastewater treatment process. The feasibility of single particle mass spectroscopy (sp-ICP-MS) was tested to determine the NP load. The results show that nTiO2 and nCeO2 are adsorbed to at least 90 percent of the sludge. Furthermore, the results indicate that there are processes during the passage of the treatment system that lead to a modification of the NP shape in the effluent, as NP are observed to be partially smaller in effluent than in the added solution. This observation was made particularly for nCeO2 and might be due to dissolution processes or sedimentation of larger particles during the passage of the treatment system.

Using a Three-Impulse Maneuvering Scheme for Returning from the Lunar Orbit to the Reentry Point of the Earth’s Atmosphere

A practically useful algorithm for constructing the trajectory of returning from a Lunar orbit to the Earth using one-impulse and three-impulse maneuvering schemes is presented. This algorithm forms the return trajectory to the reentry point of the atmosphere satisfying the given constraints, taking into account the current ephemeris and the fact that the gravitational fields of the Moon and Earth are noncentral. The computer simulation results make it possible to draw conclusions about the advantages and expedience of using a particular maneuvering scheme.

Multi-Process Integrated Planning and Guidance for Return of Reusable Aircraft Under Time Constraint

As the launch frequency of reusable launch vehicle (RLV) increases, the requirement of landing site management is increasing. Especially, the return flight time of RLV should be constrained to improve the utilization of landing site. In this paper, a multi-process integrated (MPI) planning and guidance algorithm for deorbit phase and reentry phase is proposed to meet the mission requirement of RLV. Firstly, the states of RLV at reentry point and deorbit point are given by MPI planning algorithm through inverse calculation. To be specific, by taking the terminal state as the initial value, the reentry point is obtained by the backward calculation based on a cosine-quadratic bank angle profile. According to the relationship among the reentry point, the deorbit time and the deorbit range, the parameters of deorbit phase can be calculated by efficient interpolation based on deorbit ability analysis. The bank angle profile is adjusted to make estimated total flight range approximate the predefined total range by particle swarm optimization algorithm. After that, guidance algorithm considering the terminal position and time constraints of the whole flight phase is given, and the trajectory optimization problem related to brake angle and start-up time is transformed into a series of convex optimization problems in deorbit phase, which can be solved by primal-dual interior-point method with better computational performance and satisfactory accuracy. In the reentry phase, the bank angle profile is designed and updated with improved predictor-corrector algorithm to correct the error during the flight. Finally, the feasibility and robustness of the propose algorithm are verified by multi-mission planning guidance simulation and Monte Carlo simulation.

Hardware and GNC solutions for controlled spacecraft re-entry using aerodynamic drag

Traditionally, controlled spacecraft re-entries have been conducted using propulsive de-orbit burns which are risky, expensive, and may not be possible for all vehicles. Recently, the miniaturization of technology has ushered in a new class of small satellites (such as CubeSats) that are too small to host thrusters but may require a controlled de-orbit if they contain materials capable of surviving re-entry. For all space vehicles requiring a controlled re-entry, the ability to harness the naturally occurring aerodynamic drag force for orbit control provides a cheaper and more reliable alternative to chemical propulsion.This paper discusses a comprehensive method for drag-controlled re-entry that is applicable to any vehicle capable of modulating its ballistic coefficient. First, a novel guidance generation algorithm efficient enough to run onboard a CubeSat outputs a desired ballistic coefficient profile and corresponding numerically propagated trajectory that if followed, will lead the spacecraft to a desired de-orbit location. This guidance generation algorithm is based on an analytical solution that provides convergence guarantees, ensures rapid performance, and facilitates a controllability analysis. Next, the guidance tracking algorithm utilizes an extended Kalman filter and GPS measurements to estimate the position and velocity of the satellite relative to the guidance. A full state feedback linear-quadratic-regulator (LQR) control strategy is then used to drive the relative position and velocity to zero using solely aerodynamic drag. This paper also discusses a novel retractable drag de-orbit device (D3) that can be attached to existing CubeSat structures and can easily be scaled up for larger satellites. The D3 provides passive three-axis attitude stabilization using aerodynamic and gravity gradient forces and can be repeatedly modulated to perform aerodynamically-based orbital maneuvering and controlled re-entry. The design of the planned 2U CubeSat to test the D3 and control algorithms in flight is also discussed.The re-entry point targeting algorithms were validated through extensive Monte Carlo simulations which included realistic GPS measurement errors and drag force uncertainties. The algorithms were able to guide the satellite to a desired de-orbit location with an average error below 25 km and in all cases, the targeting error was low enough for debris mitigation purposes. The accuracy and reliability of these algorithms coupled with the D3 device that has successfully undergone thermal vacuum, vibration, and fatigue testing provide a cheap, reliable, and comprehensive attitude, orbit, and de-orbit control solution that can be used on large and small space vehicles, possibly replacing conventional propulsion and attitude control systems and making space more accessible to everyone.

Построение траектории возврата с орбиты ИСЛ к точке входа в атмосферу Земли

Построение траектории возврата с орбиты ИСЛ к точке входа в атмосферу Земли

Guidance, navigation, and control solutions for spacecraft re-entry point targeting using aerodynamic drag

As large numbers of increasingly smaller spacecraft continue to be launched, means of efficient and reliable orbital maneuvering and orbit disposal have become increasingly necessary. For spacecraft that do not contain thrusters, aerodynamic drag modulation using a retractable drag device or attitude changes presents itself as an efficient way to perform orbital maneuvers and control the re-entry location.This paper introduces an aerodynamically based re-entry guidance generation algorithm for low Earth orbit spacecraft that exhibits significant accuracy, robustness, and efficiency. The paper also presents a novel guidance tracking algorithm whereby the drag device of a spacecraft is deployed or retracted relative to a nominal deployment profile (given in the guidance) based on the difference between the actual and desired state of the spacecraft. A full state feedback linear-quadratic-regulator control scheme is utilized with the Schweighart Sedgwick equations of relative motion to drive the relative position and velocity between the spacecraft and the guidance trajectory to zero. A problem-specific Extended Kalman Filter implementation is also introduced to remove noise from the GPS-derived relative motion estimate.One thousand Monte Carlo simulations of the guidance generation algorithm with randomized initial conditions and desired re-entry locations are conducted, resulting in an average guidance error of 12.5 km and a maximum error below 106 km. The tracking of these aerodynamic decay guidances with the aforementioned algorithms is also simulated with drag force uncertainties up to a factor of two and navigation errors (noise and bias) comparable to that expected from a CubeSat GPS unit. Despite these simulated errors and uncertainties, this approach provides guidance tracking down to a re-entry altitude of 120 km with a final position error under 6 km for all cases. The algorithms detailed in this paper provide a way for any spacecraft capable of modulating its drag area to autonomously perform orbital maneuvers and execute a precise re-entry.

Orbit state deviation prediction model with second-order correction due to the J2 term

The analytic orbit state deviation prediction model derived based on the State Space Perturbation Method (SSPM) has proven effective and efficient to predict the impact error of ballistic missile caused by the J2 term. However, the accuracy of this analytic model declines when the prediction point is located in the ascending portion of an orbit with large eccentricity. To deal with this problem, a second-order correction algorithm is proposed in this paper by deriving the partial derivatives of the analytic integral function with respect to the reference two-body orbit parameters. Three application scenarios are adopted to verify the accuracy of the proposed orbit state deviation prediction model, namely orbit injection point deviation prediction of carrier rocket, impact deviation prediction of ballistic missile and re-entry point deviation prediction of space shuttle. The accuracy of the proposed model is compared with the one without second-order correction as well as the Kozai’s solution. Simulation results show that the proposed model has high precision in each case.

Numerical-experimental analysis of a carbon-phenolic composite via plasma jet ablation test

Materials used in space vehicles components are subjected to thermally aggressive environments when exposed to atmospheric reentry. In order to protect the payload and the vehicle itself, ablative composites are employed as TPS (Thermal Protection System). The development of TPS materials generally go through phases of obtaining, atmospheric reentry tests and comparison with a mathematical model. The state of the art presents some reentry tests in a subsonic or supersonic arc-jet facility, and a complex type of mathematical model, which normally requires large computational cost. This work presents a reliable method for estimate the performance of ablative composites, combining empirical and experimental data. Tests of composite materials used in thermal protection systems through exposure to a plasma jet are performed, where the heat fluxes emulate those present in atmospheric reentry of space vehicles components. The carbon/phenolic material samples have been performed in the hypersonic plasma tunnel of Plasma and Process Laboratory, available in Aeronautics Institute of Technology (ITA), by a plasma torch with a 50 kW DC power source. The plasma tunnel parameters were optimized to reproduce the conditions close to the critical re-entry point of the space vehicles payloads developed by the Aeronautics and Space Institute (IAE). The specimens in study were developed and manufactured in Brazil. Mass loss and specific mass loss rates of the samples and the back surface temperatures, as a function of the exposure time to the thermal flow, were determined. A computational simulation based in a two-front ablation model was performed, in order to compare the tests and the simulation results. The results allowed to estimate the ablative behavior of the tested material and to validate the theoretical model used in the computational simulation for its use in geometries close to the thermal protection systems used in the Brazilian space and suborbital vehicles.

Study on command attitude law for refracted starlight observation in SINS/RCNS integrated navigation

To ensure the validity of refracted starlight observation, a command attitude law for maneuver is built for strapdown star sensor. The constraint relations between the relative attitude angles and the constant boresight apparent height are first established, by which attitude control command is generated, and the following attitude maneuver enables the star sensor FOV (Field of View) to keep covering the stratosphere during the midcourse. The simulating environment is set up, which involves the schema trajectory designing, the simulating algorithm of refracted star map and the modeling of the complete autonomous SINS/RCNS (Strapdown Inertial Navigation System/Refraction Celestial Navigation System) integrated navigation system. In the SINS/RCNS integrated navigation, the star sensor boresight is directed to the stratosphere in a suitable height with the command attitude law. The test simulation results show that the SINS/RCNS system prominently increases the precision of position and velocity at reentry point by 87.8% and 39.5% respectively than the sole SINS, and using of wide FOV helps to increase the positioning accuracy.

Satellite de-orbiting via controlled solar radiation pressure

The goal of the present research was to study the use of solar radiation pressure to place a satellite in an orbit that makes it to re-enter the atmosphere of the Earth. This phase of the mission is usual, since the orbital space around the Earth is crowded and all satellites have to be discarded after the end of their lifetimes. The technique proposed here is based on a device that can increase and decrease the area-to-mass ratio of the satellite when it is intended to reduce its altitude until a re-entry point is reached. Equations that predict the evolution of the eccentricity and semi-major axis of the orbit of the satellite are derived and can be used to allow the evaluation of the time required for the decay of the satellite. Numerical simulations are made, and they show the time required for the decay as a function of the area-to-mass ratio and the evolution of the most important orbital elements. The results show maps that indicate regions of fast decays as a function of the area-to-mass ratio and the initial inclination of the orbit of the satellite. They also confirmed the applicability of the equations derived here. The numerical results showed the role played by the evection and the Sun-synchronous resonances in the de-orbiting time.