Earth Orbit Applications Research Articles

A background suppression detector array for fast neutron measurement in space science study

A compact detector array with a symmetrical structure was designed and developed for the measurement of fast neutron flux in space radiation environments by the recoil proton method. It consists of seven solid-state detectors and a hydrogen-containing conversion layer, which can effectively eliminate the background from the charged particles and gamma rays. The performances for fast neutron measurements in the low Earth orbit and on the lunar surface have been simulated by Geant4 software. The signals from recoil protons can be well distinguished in a complex radiation environment. The accuracy of fast neutron flux measurement can be improved with a higher signal-to-noise ratio. Utilizing this array, a neutron spectrometer for low Earth orbit applications was constructed and was launched on board the “Weiming-1” CubeSat at the beginning of 2024.

A 33V to 1V Ripple-Less Buck Converter With the Inverted AC Current Replica Circuit and Sub-0.5% Output Ripple for 5G Low Earth Orbit Application

A 33V to 1V Ripple-Less Buck Converter With the Inverted AC Current Replica Circuit and Sub-0.5% Output Ripple for 5G Low Earth Orbit Application

Advanced analytical model for orbital aerodynamic prediction in LEO

The growing interest in low earth orbit (LEO) applications demands for accurate modeling of orbital aerodynamics. But classical analytical models of aerodynamic coefficients in free molecule flow, such as the Sentman’s model, Schamberg’s model and Schaaf-Chambre model, were built upon over simplistic gas-surface interaction models, which degrade the fidelity of aerodynamic prediction. This work presents a new analytical model of orbital aerodynamic coefficients based on the state-of-the-art Cercignani–Lampis–Lord (CLL) gas-surface interaction model, where lobular quasi-specular scattering pattern and separate accommodation degree for different velocity components can be well captured. A key component of the new model is a rigorous function approximation solution of the reflected normal momentum flux based on the CLL model which is derived for the first time and is validated within 1% for any hypothermal flow and surface accommodation conditions. Closed-form analytical solutions of aerodynamic coefficients for simple convex geometries are obtained and exhibit high accuracy (within 0.1%) in typical LEO scenarios. The new analytical model surpasses the classical models in some important aspects, such as overcoming the diffuse scattering hypothesis constraint, considering the variation of normal momentum exchange with the surface incidence angle and being applicable in any hypothermal flow situation. In virtue of the advanced CLL model and feasibility of coupling with the panel method technique, the new analytical model is promising to provide more accurate predictions on the orbital aerodynamic coefficients for LEO applications.

An analysis of X-ray pulsar navigation accuracy in Earth orbit applications

At present, studies on X-ray signal processing and the navigation filtering algorithm in X-ray pulsar navigation are largely independent; thus, navigation accuracy analysis must be based on the analysis of a sound and complete navigation system. In this study, a navigation system simulation initiated from the numerical generation of the observed signals to navigation filtering solutions was established on the basis of the requirements of the authenticity and integrity of a demonstration of Earth-orbit spacecraft navigation. The influences of the navigation observation time, time delay estimation algorithms, and different pulsar combinations imposed on navigation accuracy were therefore analyzed. The simulation results showed that an optimal observation time existed with the highest navigation accuracy in Earth-orbit applications when three pulsars were observed in parallel, thereby obtaining the optimal observation time and its corresponding navigation accuracy. Using the short time high-accuracy algorithm of time delay estimation, the navigation accuracy of low-orbit spacecraft was greatly improved by 1010 m when the observation time was 60 s. Among three combinations of pulsars analyzed in this study, the pulsar combination of PSR B0531+21, PSR B1821-24, and PSR B1937+21 achieved the highest navigation accuracy on a low orbit, whose result was 560 m. The corresponding simulation results were based on a complete and near-real navigation algorithm, thereby offering a theoretical foundation for determining the navigation accuracy and selecting the orbit and observation time parameters in Earth-orbit pulsar navigation applications.

Analysis of ionization in air-breathing plasma thruster

The primary focus of this work is to study the ionization inside an air-breathing plasma thruster (ABPT) in low earth orbit applications. For this high-speed technology to work, a high degree of ionization needs to be achieved. This paper focuses on plasma chemistry simulation for air in low earth orbits (80–110 km) to explore the possibility of high ionization of the incoming air. The results of plasma chemistry simulation showed the variation of ionization degree and species densities concerning the mean input energy that contributed to the chemical reactions. This research is essential to understand ionization processes to develop a low earth orbit ABPT design. Our results have indicated the possibility of building ABPT without an external neutralizer. The neutralization is created by extracting negative and positive ions to obtain neutralization, thereby eliminating existing design complexity.

An FPGA-Based Hardware Accelerator for CNNs Inference on Board Satellites: Benchmarking with Myriad 2-Based Solution for the CloudScout Case Study

In recent years, research in the space community has shown a growing interest in Artificial Intelligence (AI), mostly driven by systems miniaturization and commercial competition. In particular, the application of Deep Learning (DL) techniques on board Earth Observation (EO) satellites might lead to numerous advantages in terms of mitigation of downlink bandwidth constraints, costs, and increment of the satellite autonomy. In this framework, the CloudScout project, funded by the European Space Agency (ESA), represents the first time in-orbit demonstration of a Convolutional Neural Network (CNN) applied to hyperspectral images for cloud detection. The first instance of this use case has been done with an INTEL Myriad 2 VPU on board a CubeSat optimized for low cost, size, and power efficiency. Nevertheless, this solution introduces multiple drawbacks due to its design not specifically being for the space environment, thus limiting its applicability to short-lifetime Low Earth Orbit (LEO) applications. The current work provides a benchmark between the Myriad 2 and our custom hardware accelerator designed for Field Programmable Gate Arrays (FPGAs). The metrics used for comparison include inference time, power consumption, space qualification, and components. The obtained results show that the FPGA-based solution is characterized by a reduced inference time, and a higher possibility of customization, but at the cost of greater power consumption and a longer Time to Market. As a conclusion, the proposed approach might extend the potential market of DL-based solutions to long-term LEO or interplanetary exploration missions through deployment on space-qualified FPGAs, with a limited cost in energy efficiency.

Space Environment Evaluation Test of Solid-State-Ceramic Battery Advanced Energy Storage Under Vacuum and Thermal Vacuum

The desired capabilities of small satellites to enable their applications in communication, earth observation, and new scientific instruments require advanced energy storage to face the design’s challenges with the constraints of volume and mass. Small batteries with high energy density may be the solution. New lithium-ion battery technologies areimproved in order to meet these requirements by bringing higher energy density and a wide temperature range than the commercially available ones as well as a lower risk of explosion. In this paper, the ability of solid-state-ceramic batteries to withstand the vacuum and thermal vacuum for low earth orbit applications has been demonstrated, with a minimum safety issue. So far, this technology has never been flown in space. This paper also provides a guideline for the battery evaluation test where the main lines are represented. Batteries are tested under vacuum and thermal vacuum, where they are discharged and charged during several cycles between two temperatures limits. The evaluation focuses on analyzing the physical degradation, the discharge capacity, and the internal resistance before and after each test. Batteries have showed promising results regarding their survivability to thermal vacuum. After several cycles, they have kept almost the same performances, with the same internal resistance and 98% of capacity.

Design and Fabrication of a Planar Inverted-F Antenna (PIFA) for LEO Satellite Application

The paper is mainly emphasizes on the microstrip patch antenna design for satellite applications. The specific objective of this study is to solve the research problem on finding the simple and compact design for small satellite application in real world. A new design of a Planar Inverted-F Antenna (PIFA) with the two height of 12mm and 10mm are proposed for low Earth orbit (LEO) applications in S band. The antenna is a single form antenna with the coaxial probe fed is used. CST Microwave Studio student version was used for the simulation of the antenna and matching design parameters. The antennas are obtained an efficient high return loss -18.426dB and -18.169dB at S11, the return loss bandwidth of 471MHz and 466MHz. The radiation pattern results are analyzed beam width angle 85.1° and 95.1°. Moreover, the peak directivity gain results 4.7dBi and 4.79dBi, the absolute E-filed radiation effect are also presented. The actual measurement results for S11 parameter of return loss -13.54dB and -14.2dB, return loss bandwidth of 510MHz and 425MHz are also fabricated with the comparison of simulation and fabrication process of PIFA antenna characteristics. The measurement of the return loss (S11) and return loss bandwidth were almost identical with the simulation.

Computational Design of Metal–Fabric Orbital Debris Shielding

Hybrid particle–finite element methods, developed specifically to simulate hypervelocity impact physics, may be used to compliment experimental and analytical research on micrometeoroid and orbital debris shielding. Recent research has applied a hybrid particle–finite element method to model the hypervelocity impact response of an enhanced metal–fabric orbital debris shield developed for the Soyuz Orbital Module, composed of both thermal insulation and orbital debris protection layers. The numerical model was employed to simulate impact effects in a high-velocity impact regime outside the capabilities of current light-gas guns. The simulation results suggest that momentum scaling may be applied to estimate the ballistic limit performance of multilayer composite shields, for spherical projectiles over an impact velocity range important in low Earth orbit applications but not accessible to experiments.

A smooth and robust Harris-Priester atmospheric density model for low Earth orbit applications

The modified Harris-Priester model is a computationally inexpensive method for approximating atmospheric density in the thermosphere and lower exosphere – a vital step in low Earth orbit trajectory propagation. This work introduces a revision, dubbed cubic Harris-Priester, which ensures continuous first derivatives, eliminates singularities, and adds a mechanism for introducing smooth functional dependencies on environmental conditions. These changes increase the accuracy, robustness, and utility of the model, particularly for preliminary orbit propagation, estimation, and optimization applications in which fast, reasonably accurate force models and sensitivities are desirable. Density results and computational efficiency are compared to other density models. The Fortran code used to implement the model is provided as an electronic supplement.

Link budget analysis for future E-band gigabit satellite communication links (71–76 and 81–84 Ghz)

The paper shows the feasibility of a satellite-earth downlink in E-band (60–90 GHz) taking into account both the expected channel attenuation and hardware characteristics. For low earth orbit (LEO) applications, the attenuation and the link are studied as a function of the elevation angle seen at the ground station. Link budget calculations for a realistic scenario indicating the feasibility of broadband and multi-gbit/s E-band communication between LEO satellites and ground stations with available receive antenna sizes and availabilities above 98% are presented. Additionally, a possible E-band geostationary earth orbit (GEO) link budget with 5 Gbit/s of data rate is presented, including atmospheric losses.

Radiation-Tolerant Superfluorescent Fiber Sources for High-Performance Fiber-Optic Gyroscopes Working Under Gamma Irradiation Higher Than 200 krad

Two effective methods, bidirectional-pump doublepass (BPDP) configuration and green light photo-annealing (GLPA) technique, are proposed to obtain radiation-tolerant superfluorescent fiber sources (SFSs) for high performance interferometric fiber-optic gyroscopes ('FOGs). The results show that the SFS prepared by each method could maintain the same meanwavelength under γ-irradiation higher than 200 krad, and still have a broad linewidth larger than 34 nm. The lowest output power loss was experimentally demonstrated to be as low as 0.6 dB for 40-mW output power. Such radiation-tolerant SFSs have better performance than those reported previously, and can be used for an 'FOG in both the low-Earth orbit (LEO) and the geostationary Earth orbit (GEO) applications.

Survey of Spacecraft Trajectory Design in Strongly Perturbed Environments

A S KEPLER pointed out centuries ago, the motion of a spacecraft governed by a gravitating point mass is a wellbehaved conic section. In the low-energy regime, the conic section transcribes an ellipse and the motion is periodic and stable. Furthermore, the gravitational potential of a spherical mass with radially symmetric density is identical to that of a simple point mass located at the center of the sphere. Despite the fact that many perturbations are important when considering high-precision Earthorbiting applications, all perturbations combined are generally less than one part in a thousand compared with the dominant two-body term due to a spherical Earth. Therefore, the motion of a low-Earthorbiting ballistic spacecraft is very near periodic, and changes to the qualitative motion are slowly time-varying. For these reasons, the near-Earth environment indeed can be considered dynamically benign when compared with a great number of other locations of interest for space missions in our solar system. For the purposes of this paper, we will therefore consider strongly perturbed environments as those exotic locales beyond the near-Earth environment where perturbations to the Kepler two-body problem are no longer small. Notable examples of space missions in such highly perturbed dynamical environments include lunar missions; grand tours of the solar system and planetary satellite systems propelled by gravityassisted flybys; orbiters around major planetary satellites such as Jupiter’s Galilean moons and Saturn’s Titan and Enceladus; and missions to very small and irregular-shaped bodies such as comets, asteroids, and small planetary satellites. Drawing from common models such as patched conics of interplanetary missions and the restricted three-body problem (RTBP) of the libration point missions, this paper is organized in terms of the dominant perturbation forces on a spacecraft. The focus of this survey includes the wide-reaching perturbations that permeate globally the spacecraft environment over the course of a mission. Included are the third-body (TB) nonspherical gravity (NSG) and solar radiation pressure (SRP) perturbations. In their simplified forms, each of the main perturbations can be expressed as part of an autonomous and conservative dynamical system. Therefore, the ensuing analyses can benefit from a host of tools and insights including integrals of motion and system averaging that assist fast numerical methods and the recovery of equilibria and secular trends. Mission destinations of interest to the space science and space flight communities are discussed according to their respective dynamic environments. Special attention is given to orbit mechanics and its impact on mission and trajectory design techniques in the context of past and planned space missions to these dynamically rich environments. Mission classes and design techniques considered include lowand high-altitude orbiters, grand tour trajectories, patched conics, gravity-assistedflybys, Tisserand graphs, three-body models, periodic orbits, orbit averaging, and stable/unstable manifold design. A main result of this survey is a mission design reference (see Table 1 and the associated Fig. 1) that quantifies the absolute and relative importance of the main perturbations terms in the dynamic environments of many target bodies. It is acknowledged that atmospheric drag can induce extreme spacecraft heating and torques in the case of aerobraking, aerocapture, or entry decent and landing for missions to bodies with significant atmospheres such as Earth, Mars, Venus, or Titan. However, such extreme drag-related events are generally restricted to brief mission phases and can often be treated as nonpropellant propulsion (energy reducing) sources in the overall mission trajectory design. Therefore, atmospheric drag is not included in the scope of the paper. Finally, in the context of asteroids and comets, it is worth noting that small forces such as the Yarkovski and Yarkovsky–O’Keefe–Radzievskii– Paddacky (YORP) effects (arising mainly from imbalanced thermal radiation) and outgassing are important in the consideration of longterm small body orbital and spin rate evolution [1]. However, their indirect perturbations on the short-term dynamics of a nearby spacecraft are too small to warrant consideration in the current scope. For further information on the details and relative importance of dynamical perturbations on spacecraft in both the Earth and interplanetary environments, see [2–4].

Control of close formations in low earth orbit

A number of recent mission concepts have highlighted the need for utilisation of formations of spacecraft. These include high precision scientific missions (using optical interferometry), disaster monitoring and RF interferometry. A significant low Earth orbit application is distributed radar systems, involving linked radar transmitter/receiver systems located on spacecraft flying in close proximity. The resulting configuration is such that these spacecraft undergo a relative motion described by a set of concentric ellipses. This example has been the focus of a study for the development of a station control system. The formation must maintain the relative spacecraft separations to within a tolerance of a few metres. This paper is concerned with continuing developments in orbit control strategies designed to maintain the formation with minimum DeltaV. Demonstration is by development of a closed loop control system. Simultaneous position and attitude control are executed. The method presented involves frequent small manoeuvres over the period of a day. The model includes navigation system simulation. The sensitivity of the translational control to attitude demands is examined. Closed loop simulation shows the feasibility of this station keeping control approach as an automatic control system. Safe spacecraft separation is assured and efficient fuel usage is obtained.

Effects of electrode thickness on power capability of a sintered-type nickel electrode

The performance capability of sintered-type nickel electrodes with thicknesses of up to 4.6 mm was studied in an Ni/H 2 cell. The utilization of active material at high rate discharge diminishes as the thickness of electrode increases. Full utilization at C rate discharge was obtained with electrodes with a thickness of 2.5 mm. A 1 mm thick electrode showed a well-defined discharge curve up to 20 C rate discharge. Electrodes of up to 3.5 mm thickness showed sufficiently high rate capability for a geosynchronous earth orbit (GEO) application, while the maximum acceptable thickness for a low earth orbit (LEO) application is limited to a value below 2.5 mm. It appears that there is a limited amount of usable capacity at a given current density because of electrode polarization. The value of the current density appears to be independent of electrode thickness.

Nickel-hydrogen batteries - An overview

This article on nickel-hydrogen batteries is an overview of the various nickel-hydrogen battery design options, technical accomplishments, validation test results, and trends. There is more than one nickel-hydrogen battery design, each having its advantage for specific applications. The major battery designs are Individual Pressure Vessel (IPV), Common Pressure Vessel (CPV), bipolar, and low-pressure metal hydride. State-of-the-art nickel-hydrogen batteries are replacing nickel-cadmium batteries in almost all geosynchronous Earth orbit applications requiring power above 1 kW. However, for the more severe Low-Earth Orbit (LEO) applications (greater than 30,000 cycles), the current cycle life of 4000-10,000 cycles at 60 - 80 % DOD should be improved. A NASA Lewis Research Center innovative advanced design IPV nickel-hydrogen cell led to a breakthrough in cycle life enabling LEO applications at deep Depths of Discharge (DOD). A trend for some future satellites is to increase the power level to greater than 6 kW. Another trend is to decrease the power to less than 1 kW for small low-cost satellites. Hence, the challenge is to reduce battery mass, volume, and cost. A key is to develop a lightweight nickel electrode and alternate battery designs. A CPV nickel-hydrogen battery is emerging as a viable alternative to the IPV design. It has the advantage of reduced mass, volume, and manufacturing costs. A 10-A-h CPV battery has successfully provided power on the relatively short-lived Clementine spacecraft. A bipolar nickel -hydrogen battery design has been demonstrated (15,000 LEO cycles, 40 % DOD). The advantage is also a significant reduction in volume, a modest reduction in mass, and like most bipolar designs, features a high-pulse power capability. A low-pressure aerospace nickel-metal-hydride battery cell has been developed and is on the market. It is a prismatic design that has the advantage of a significant reduction in volume and a reduction in manufacturing cost.

Mission Design Drivers for Closed Brayton Cycle Space Power Conversion Configuration

Future space power requirements will vary from the subkilowatt range for deep space probes, to the hundreds of kilowatts range for a lunar base, to the multimegawatt range for interplanetary propulsion systems. Closed Brayton cycle (CBC) power conversion has the flexibility to be used in all these power ranges and with a variety of heat source options such as isotope, solar, and nuclear. Each of these types of heat sources has different characteristics that make it more appropriate for particular mission profiles and power output ranges. Heat source characteristics can also be major design drivers in the closed Brayton cycle design optimization process. This paper explores heat source selection and the resulting CBC system designs, and discusses optimization methods as a function of the main design drivers. Such power system requirements as power level, man-rated radiation shielding, fuel costs, eclipse/darkness duration, system mass, radiator area, reliability/mission duration, and insolation level are evaluated through several CBC parametric case studies. These cases include: (1) a 500 We power system for deep space probes; (2) a 50 kWe solar dynamic system for earth orbit and other applications; (3) a 100 kWe man-rated lunar/Mars stationary /rover power system; (4) a 200 to 825 kWe power system for the lunar outpost; and (5) 3300 kWe modules for interplanetary propulsions.

Application of high temperature multilayer insulations

Low temperature multilayer insulation systems are highly efficient in usual satellite applications. Thermal control in the field of Earth reentry systems as well as space power plants benefit from the application of high temperature multilayer insulations. The analytic prediction of high temperature multilayer insulation performance in a changing environment is enabled if the dependence of the heat transfer modes on the various design parameters is known. Systematic and tailored design of the insulation becomes feasible then. This is illustrated by three typical examples like a planetary probe for the exploration of the solar system, an Earth based space transportation system and a solar-dynamic space power system for low Earth orbit application.

Advances in laser-pumped helium magnetometers for space applications (abstract)

Recently tunable lasers have been developed which are suitable for use as pumping sources in optically pumped helium magnetometers. This advance in laser technology is leading to a new generation of improved laser-pumped resonance magnetometers for use in geophysical and space applications. The history of optically pumped resonance magnetometers in space exploration will be reviewed, in order to describe the anticipated benefits of laser pumping. Experiments will be described which have demonstrated signal strengths in laser-pumped helium magnetometers which are 45 times greater than those observed in the same sensor pumped with a conventional rf electrodeless discharge helium lamp. This was a collaborative effort with the Magnetics Group at the Jet Propulsion Laboratory and Professor L. D. Schearer’s group at the University of Missouri at Rolla. The impact of the laser pumping technique on sensitivity, accuracy, and versatility of space magnetometers will be discussed. Proposed laser-pumped magnetometers will be described including the vector helium magnetometer for interplanetary field measurement, the scalar helium magnetometer for earth orbital applications, and the vector/scalar hybrid magnetometer for planetary and cometary encounters. Recent developments in miniaturized turnable laser sources will be presented. This was supported by the Jet Propulsion Laboratory/NASA.

A new concept for high-cycle-life LEO: rechargeable MnO 2—hydrogen batteries

The nickel—hydrogen secondary battery system, developed in the early 1970s, has become the system of choice for geostationary earth orbit (GEO) application. However, for low earth orbit (LEO) satellites with long expected lifetimes the nickel positive limits performance. This requires derating of the cell to achieve very long cycle life. A new system, rechargeable MnO 2—hydrogen, which does not require derating, is described here. For LEO applications, it promises to have longer cycle life, higher rate capability, a higher effective energy density, and much lower self-discharge behavior than those of the nickel—hydrogen system.