Final Orbit Research Articles

Cassini UVIS Detection of Saturn's North Polar Hexagon in the Grand Finale Orbits

AbstractCassini's final orbits in 2016 and 2017 provided unprecedented spatial resolution of Saturn's polar regions from near‐polar spacecraft viewing geometries. Long‐wavelength channels of Cassini's Ultraviolet Imaging Spectrograph instrument detected Saturn's UV‐dark north polar hexagon near 180 nm at planetocentric latitudes near 75°N. The dark polar hexagon is surrounded by a larger, less UV‐dark collar poleward of planetocentric latitude 65°N associated with the dark north polar region seen in ground‐based images. The hexagon is closely surrounded by the main arc of Saturn's UV aurora. The UV‐dark material was locally darkest on one occasion (23 January 2017) at the boundary of the hexagon; in most Ultraviolet Imaging Spectrograph images the dark material more uniformly fills the hexagon. The observed UV‐dark stratospheric material may be a hydrocarbon haze produced by auroral ion‐neutral chemistry at submicrobar pressure levels. Ultraviolet Imaging Spectrograph polar observations are sensitive to UV‐absorbing haze particles at pressures lower than about 10–20 mbar.

Optimization of space vehicle combined orbital plane change maneuver on the basis of pontryagin’s maximum principle

The task of optimizing the space vehicle combined orbital plane change maneuver with the aim of maximizing its final mass is considered in the paper. Burst of power is used for the vehicles exit from the initial low earth orbit and subsequent re-entry. Starting from atmospheric entry till the end of the orbital plane change with the entry in the final orbit the angle of attack, the air-path bank angle and the fuel-flow rate are used as controls. Limitations for the angle of attack, fuel-flow rate, adiabatic recovery temperature, longitudinal and vertical load factor are introduced. The successive linearization method is used to determine the optimal control programs. Solving the optimization task is exemplified by a hypothetical space vehicle. The results of modeling space vehicle motion are presented. The changes in the determined aerodynamic (angle of attack and air-path bank angle) and motion (fuel-flow rate) controls with increasing the angle of orbital plane change are discussed with and without account of the key limitation on the modes of motion maximum adiabatic recovery heating temperature.

Comparative analysis of lunar capture braking method based on particle swarm optimization

Purpose The purpose of this paper is to verify the feasibility of lunar capture braking through three methods based on particle swarm optimization (PSO) and compare the advantages and disadvantages of the three strategies by analyzing the results of the simulation. Design/methodology/approach The paper proposes three methods to verify capture braking based on PSO. The constraints of the method are the final lunar orbit eccentricity and the height of the final orbit around the Moon. At the same time, fuel consumption is used as a performance indicator. Then, the PSO algorithm is used to optimize the track of the capture process and simulate the entire capture braking process. Findings The three proposed braking strategies under the framework of PSO algorithm are very effective for solving the problem of lunar capture braking. The simulation results show that the orbit in the opposite direction of the trajectory has the most serious attenuation at perilune, and it should consume the least amount of fuel in theoretical analysis. The methods based on the fixed thrust direction braking and thrust uniform rotation braking can better ensure the final perilune control accuracy and fuel consumption. As for practice, the fixed thrust direction braking method is better realized among the three strategies. Research limitations/implications The process of lunar capture is a complicated process, involving effective coordination between multiple subsystems. In this article, the main focus is on the correctness of the algorithm, and a simplified dynamic model is adopted. At the same time, because the capture time is short, the lunar curvature can be omitted. Furthermore, to better compare the pros and cons of different braking modes, some influence factors and perturbative forces are not considered, such as the Earth’s flatness, light pressure and system noise and errors. Practical implications This paper presents three braking strategies that can satisfy all the constraints well and optimize the fuel consumption to make the lunar capture more effective. The results of comparative analysis demonstrate that the three strategies have their own superiority, and the fixed thrust direction braking is beneficial to engineering realization and has certain engineering practicability, which can also provide reference for lunar exploration orbit design. Originality/value The proposed capture braking strategies based on PSO enable effective capture of the lunar module. During the lunar exploration, the capture braking phase determines whether the mission will be successful or not, and it is essential to control fuel consumption on the premise of accuracy. The three methods in this paper can be used to provide a study reference for the optimization of lunar capture braking.

Leveraging Artificial Neural Networks to Systematically Explore Solar Gravity Driven Transfers in the Martian System

Current solar electric propulsion and launch vehicle technology allow sending multiple spacecraft to Mars simultaneously. For capture orbits with very high apoareion, the solar gravity perturbations deploy the spacecraft in vastly different areocentric orbits. This application requires knowledge of possible transfers and an efficient way to identify them; for example, a well-defined and easily accessed database of solutions. First, a method is developed to identify the required initial orbital elements and timing to target a final orbit, for a dynamical model ignoring the Martian eccentricity. This method can be applied to any planetary system with low heliocentric eccentricity. Second, the effect of Mars’ eccentricity and the corresponding time-dependence of solar distance are analyzed. This analysis requires a much larger number of numerical integrations. The research is therefore limited to placing bounds on the solution space. Third, this paper determines the way artificial neural networks can reduce the large number of required integrations by order(s) of magnitude while maintaining sufficient accuracy to enable preliminary transfer design. The neural networks enable the design of transfers between a low, near-polar orbit to Phobos and Deimos in an approximation of the eccentric model. These transfers are then validated in a higher fidelity ephemeris model.

Electron Density Distributions in Saturn's Ionosphere

AbstractBetween 26 April and 15 September 2017, Cassini executed 23 highly inclined Grand Finale orbits through a new frontier for space exploration, the narrow region between Saturn and the D Ring, providing the first opportunity for obtaining in situ ionospheric measurements. During the Grand Finale orbits, the Radio and Plasma Wave Science instrument observed broadband whistler mode emissions and narrowband upper hybrid frequency emissions. Using known wave propagation characteristics of these two plasma wave modes, the electron density is derived over a broad range of ionospheric latitudes and altitudes. A two‐part exponential scale height model is fitted to the electron density measurements. The model yields a double‐layered ionosphere with plasma scale heights of 545/575 km for the northern/southern hemispheres below 4,500 km and plasma scale heights of 4,780/2,360 km for the northern/southern hemispheres above 4,500 km. The interpretation of these layers involves the interaction between the rings and the ionosphere.

Uncertainty Analysis and Robust Design for a Hybrid Rocket Upper Stage

An uncertainty analysis is performed for a hybrid rocket engine, and a robust design method is proposed and applied to a liquid-oxygen/paraffin–based fuel hybrid rocket engine that powers the third stage of a Vega-like launcher. An indirect method is used to optimize a mission-specific objective function, which takes into account both the payload mass and the ability of the rocket to reach the required final orbit despite uncertainties. First, a screening of uncertain model parameters is performed using the Morris method. The screening shows that three groups can be distinguished: parameters with negligible or no effect on model outputs, parameters with small linear effect on model outputs, and parameters with large nonlinear effect on model outputs. The robust optimization design process is then carried out by a particle swarm optimization algorithm considering only the uncertainties of the latter group, which includes six parameters. A fractional factorial design of experiment techniques is considered to characterize the robust performance index. Average launcher performance is considered and the best solution is found by means of Box–Behnken’s array. Finally, the identified robust solution is checked against the second group of parameters uncertainties previously neglected in the robust optimization procedure. The scattering of the robust solution is small, and the robustness of the selected optimal solution is confirmed.

Viability of an Electrically Driven Pump-Fed Hybrid Rocket for Small Launcher Upper Stages

An electrically driven pump-fed cycle for a hybrid rocket engine is proposed and compared to a simpler gas-pressurized feed system. A liquid-oxygen/paraffin-based fuel hybrid rocket engine which powers the third stage of a Vega-like launcher is considered. Third-stage ignition conditions are assigned, and engine design and payload mass are defined by a proper set of parameters. Uncertainties in the classical regression rate correlation coefficients are taken into account and robust design optimization is carried out with an approach based on an epsilon-constrained evolutionary algorithm. A mission-specific objective function, which takes into account both the payload mass and the ability of the rocket to reach the required final orbit despite uncertainties, is determined by an indirect trajectory optimization approach. The target orbit is a 700 km altitude polar orbit. Results show that electrically driven pump-fed cycle is a viable option for the replacement of the conventional gas-pressurized feed system. Robustness in the design is granted and a remarkable payload gain is achieved, using both present and advanced technologies for electrical systems.

Search-based method optimization applied to bi-impulsive orbital transfer

In this paper optimal bi-impulse orbital transfer between coplanar and non-coplanar elliptical orbits has been considered. First, an efficient algorithm based on grid search method has been developed to find the global solution of the impulsive orbital transfer. There are three independent parameters in this algorithm: i) the angular position of the point on the initial orbit where the first impulse is applied to, ii) the angular position of the point on the final orbit where the second impulse is applied to and, iii) the argument of periapsis of the transfer orbit. The other parameters can be obtained from these three parameters in both coplanar and non-coplanar cases. The algorithm is able to obtain the global optimal solution for any arbitrary initial and final elliptical orbits. Second, the algorithm is employed to solve an extensive set of numerical examples, including coplanar and non-coplanar, co-axial and non-co-axial orbits of different shape and energy. This comprehensive set of examples allows disentangling the roles of each relevant orbital parameter in the evaluation of the optimal transfer. Finally, the numerical results are employed to achieve a very precise fitting function which is helpful onboard calculation and mission design.

Improving integrated precise orbit determination of GPS, GLONASS, BDS and Galileo through integer ambiguity resolution

The China Satellite Navigation Office announced on December 27, 2018 that the BDS-3 preliminary system had been completed to provide global services. Before this, GPS and GLONASS were the only two global navigation satellite systems (GNSS) supporting global positioning service, and have totally more than 50 satellites in normal operation. Furthermore, Galileo is intending to reach its full constellation around 2020. By that time, the number of available GNSS satellites will increase to more than 100, which brings both opportunities and challenges for high-precision positioning and orbit determination. The precision of orbits could be significantly improved through integer ambiguity resolution (AR), while AR is particularly difficult to achieve especially for GLONASS and BDS due to inter-frequency biases and satellite-induced code biases. Therefore, to address this limitation and further enhance the precision of multi-GNSS orbit determination, we try to fix the double-differenced intra-system ambiguities to integers and propose an integer AR method for multi-GNSS POD. To verify the contribution of AR, an experiment of 141 sites with global coverage is conducted. The results imply that the approach realizes an average fixing rate of 98.1%, 96.4%, 84.6% and 92.6% for GPS, GLONASS, BDS and Galileo over a whole year. The GNSS orbits are further improved with AR in terms of the precision compared with the International GNSS Service (IGS) final orbits, the discontinuity at overlapping day boundaries, and satellite laser ranging residuals. Thus, the integer AR improves the precision of multi-GNSS precise orbit determination, which can enhance integrated data processing of multi-GNSS and their applications in the future.

Consequences of planetary migration on the minor bodies of the early solar system

Pebble accretion is an efficient mechanism that is able to build up the core of the giant planets within the lifetime of the protoplanetary disc gas-phase. The core grows via this process until the protoplanet reaches its pebble isolation mass and starts to accrete gas. During the growth, the protoplanet undergoes a rapid, large-scale, inward migration due to the interactions with the gaseous protoplanetary disc. In this work, we have investigated how this early migration would have affected the minor body populations in our solar system. In particular, we focus on the Jupiter Trojan asteroids (bodies in the coorbital resonance 1:1 with Jupiter, librating around the L4and L5Lagrangian points called, respectively, the leading and the trailing swarm) and the Hilda asteroids. We characterised their orbital parameter distributions after the disc dispersal and their formation location and compare them to the same populations produced in a classical in situ growth model. We find that a massive and eccentric Hilda group is captured during the migration from a region between 5 and 8 au and subsequently depleted during the late instability of the giant planets. Our simulations also show that inward migration of the giant planets always produces a Jupiter Trojans’ leading swarm more populated than the trailing one, with a ratio comparable to the current observed Trojan asymmetry ratio. The in situ formation of Jupiter, on the other hand, produces symmetric swarms. The reason for the asymmetry is the relative drift between the migrating planet and the particles in the coorbital resonance. The capture happens during the growth of Jupiter’s core and Trojan asteroids are afterwards carried along during the giant planet’s migration to their final orbits. The asymmetry and eccentricity of the captured Trojans correspond well to observations, but their inclinations are near zero and their total mass is three to four orders of magnitude higher than the current population. Future modelling will be needed to understand whether the dynamical evolution of the Trojans over billions of years will raise the inclinations and deplete the masses to observed values.

Saturn's Dusty Ionosphere

AbstractMeasurements of electrons and ions in Saturn's ionosphere down to 1,500‐km altitudes as well as the ring crossing region above the ionosphere obtained by the Langmuir probe onboard the Cassini spacecraft are presented. Five nearly identical deep ionosphere flybys during the Grand Finale orbits and the Final plunge orbit revealed a rapid increase in the plasma densities and discrepancies between the electrons and ions densities (Ne and Ni) near the closest approach. The small Ne/Ni ratio indicates the presence of a dusty plasma, a plasma which charge carrier is dominated by negatively charged heavy particles. Comparison of the Langmuir probe obtained density with the light ion density obtained by the Ion and Neutral Mass Spectrometer confirmed the presence of heavy ions. An unexpected positive floating potential of the probe was also observed when Ne/Ni ≪ 1. This suggests that Saturn's ionosphere near the density peak is in a dusty plasma state consisting of negatively and positively charged heavy cluster ions. The electron temperature (Te) characteristics in the ionosphere are also investigated and unexpectedly high electron temperature value, up to 5000 K, has been observed below 2,500‐km altitude in a region where electron‐neutral collisions should be prominent. A well‐defined relationship between Te and Ne/Ni ratio was found, implying that the electron heating at low altitudes is related to the dusty plasma state of the ionosphere.

Оптимизация целевой орбиты и траектории апсидального импульсного выведения космического аппарата на нее с учетом сброса отработавших ступеней в атмосферу

The article considers the idea of reducing littering the near-Earth space by means of spent stage jettisons into the Earth's atmosphere. A solution to the problem of optimization of the apsidal impulse transfer between the reference circular orbit of an Earth’s artificial satellite and the final elliptical orbit is proposed. A parametric analysis of the obtained solutions is performed, a simple pulse choice scheme close in functionality to the optimal one is proposed. Questions about the optimal number and location of pulses and the optimal mass of the first stage are investigated. An assessment of the additional mass consumption associated with the stage jettisons into the atmosphere comparing with the simple separation of the stages is performed. It was found that in the case of optimal fuel distribution among the tanks such mass consumption is small when a final ascent pulse is 1.5 km/s.

REAL TIME PPP APPLIED TO AIRPLANE FLIGTHT TESTS

The availability in real time of GNSS satellites orbits, clock corrections and code and phase biases provided the possibility of application of Real Time Precise Point Positioning (RTPPP). This paper presents the methodology concerning RTPPP and application to kinematic trajectories of airplane flight tests, but without using the carrier phase bias. So, it is PPP float solution. It requires RT positioning estimation, task that most of time presents certain difficulties due to loss of communication or of satellites during maneuvers of the airplane. However, if the corrections become unavailable for a certain period of time, the system starts using the ultra-rapid IGS orbits. The experiments were accomplished taking into account a case simulating RT and another in fact RT, but storing data and corrections for post processing. The PPP solutions obtained either simulating RT or in RT were compared against the PPP post processed solution that uses the final clock and orbit corrections. Then, statistics were generated to analyze the quality of both results. They were applied to kinematic trajectory that on average was 360 km/h, reaching about 600 km/h. The results provided accuracy better than the requisites for such cases which is of about 80 cm in height.

Validation of Geostationary Earth Orbit Satellite Ephemeris Generated from Satellite Laser Ranging

This study presents the generation and accuracy assessment of predicted orbital ephemeris based on satellite laser ranging (SLR) for geostationary Earth orbit (GEO) satellites. Two GEO satellites are considered: GEO-Korea Multi-Purpose Satellite (KOMPSAT)-2B (GK-2B) for simulational validation and Compass-G1 for real-world quality assessment. SLR-based orbit determination (OD) is proactively performed to generate orbital ephemeris. The length and the gap of the predicted orbital ephemeris were set by considering the consolidated prediction format (CPF). The resultant predicted ephemeris of GK-2B is directly compared with a pre-specified true orbit to show 17.461 m and 23.978 m, in 3D root-mean-square (RMS) position error and maximum position error for one day, respectively. The predicted ephemeris of Compass-G1 is overlapped with the Global Navigation Satellite System (GNSS) final orbit from the GeoForschungsZentrum (GFZ) analysis center (AC) to yield 36.760 m in 3D RMS position differences. It is also compared with the CPF orbit from the International Laser Ranging Service (ILRS) to present 109.888 m in 3D RMS position differences. These results imply that SLR-based orbital ephemeris can be an alternative candidate for improving the accuracy of commonly used radar-based orbital ephemeris for GEO satellites.

Energetic Neutral and Charged Particle Measurements in the Inner Saturnian Magnetosphere During the Grand Finale Orbits of Cassini 2016/2017

AbstractWe give an overview of the charged and neutral particle measurements during the Grand Finale of Cassini performed by the Magnetosphere Imaging Instrument, MIMI, on board. The first in situ measurements of MeV particles between the D‐ring and the atmosphere of the planet showed protons with energies above 300 MeV with higher count rates when magnetically connected to the D‐ring which was less dense than expected, while more material was present at the inner edge. The variability in the count rates was a combination of sensor look direction, pitch angle coverage, changes in the L‐range, and latitude. Enhanced emissions of energetic neutral atoms during ring current activity was detected and confirmed the existence of a low‐energy ion radiation belt discovered by Cassini in 2004. No detection of low‐energy ions was indicative of a source population from a region closer than 1,800 km above the 1 bar cloud level.

Electromagnetic Energy Harvester for Vibration Control of Space Rack: Modeling, Optimization, and Analysis

AbstractA space rack suffers from shock and broadband random loads during shipment and launch, which may lead to irreversible damage until the spacecraft reaches final orbit. This paper presents a ...

Diving within Saturn's rings.

![Figure][1] A diagram of Cassini's final orbits (blue curves), showing the last ring-grazing orbit and the 22 Grand Finale orbits. IMAGE: NASA JET PROPULSION LABORATORY Launched in 1997, the Cassini spacecraft arrived at Saturn in 2004 and spent the following 12 years orbiting far outside the planet's rings. Among the mission's many discoveries were potentially habitable environments on Saturn's moons, including a subsurface ocean of water on Enceladus. To avoid accidental contamination of the moons with any Earth bacteria that might have hitched a ride on the spacecraft, mission planners decided that when Cassini ran out of fuel it would be deliberately crashed into Saturn. As fuel reserves ran low, the spacecraft embarked on a series of maneuvers that took it through previously unexplored regions and ended with its destruction. The first phase was a series of 20 “ring-grazing” orbits in which Cassini passed near the outer edge of the rings. Then the trajectory was shifted to fly through the gap between the planet and the inner edge of the rings, a phase dubbed the “Grand Finale.” Cassini flew through this region 22 times, in various orientations designed to optimize the scientific measurements made by its numerous instruments. Finally, in 2017 Cassini dove into the planet itself at 35 km s−1, leading to the spacecraft's fiery disintegration in the upper atmosphere. The papers in this section present results from those final phases of the mission. They include studies of Saturn's puzzling magnetic field, an isolated inner radiation belt, auroral radio emission, the mass and composition of solid grains falling from the rings toward the planet, and the composition of Saturn's atmosphere. Many more results will follow from Cassini's last period of exploration in the Saturn system. [1]: pending:yes

Dust grains fall from Saturn's D-ring into its equatorial upper atmosphere.

The sizes of Saturn's ring particles range from meters (boulders) to nanometers (dust). Determination of the rings' ages depends on loss processes, including the transport of dust into Saturn's atmosphere. During the Grand Finale orbits of the Cassini spacecraft, its instruments measured tiny dust grains that compose the innermost D-ring of Saturn. The nanometer-sized dust experiences collisions with exospheric (upper atmosphere) hydrogen and molecular hydrogen, which forces it to fall from the ring into the ionosphere and lower atmosphere. We used the Magnetospheric Imaging Instrument to detect and characterize this dust transport and also found that diffusion dominates above and near the altitude of peak ionospheric density. This mechanism results in a mass deposition into the equatorial atmosphere of ~5 kilograms per second, constraining the age of the D-ring.

Dust Observations by the Radio and Plasma Wave Science Instrument During Cassini's Grand Finale

AbstractDust particles in the Saturn system can be detected by the Radio and Plasma Wave Science (RPWS) instrument on board Cassini via antenna voltage signals induced by dust impacts. These impact signals have been simulated in the laboratory by accelerating dust particles onto a Cassini model with electric field antennas. RPWS dust measurements have been shown to be consistent with the Cosmic Dust Analyzer. During the Grand Finale orbits, Cassini flew through the gap between the D ring and Saturn's atmosphere 22 times. In situ measurements by RPWS helped quantify the hazards posed to the spacecraft and instruments on board, which showed a micron‐sized dust density orders of magnitude lower than that observed during the Ring Grazing orbits. Close inspection of the waveforms indicated a possible dependence of the impact signal decay time on ambient plasma density.

Enceladus Auroral Hiss Emissions During Cassini's Grand Finale

AbstractCassini's Radio and Plasma Wave Science (RPWS) instrument detected intense auroral hiss emissions during one of its perikrone passes of the Grand Finale orbits. The emissions were detected when Cassini traversed a flux tube connected to Enceladus' orbit (L‐shell = 4) and at a time when both the spacecraft and the icy moon were in similar longitudes. Previous observations of auroral hiss related to Enceladus were made only during close flybys and here we present the first observation of such emissions close to Saturn. Further, ray‐tracing analysis shows the source location at a latitude of 63°, in excellent agreement with earlier UVIS observations of Enceladus' auroral footprint by Pryor et al. (2011, https://doi.org/10.1038/nature09928). The detection has been afforded exclusively by the Grand Finale phase, which enabled sampling of Enceladus' high‐latitude flux tube near Saturn. This result provides new insight into the spatial extent of the electrodynamic interaction between Saturn and Enceladus.