Single Launch Research Articles

Exploring the Feasibility of Kuiper Belt Missions Supported by Centrifugal Nuclear Thermal Propulsion

Missions to the Kuiper belt have previously been carried out only as flybys and with very small payloads. Investigating launch windows for Kuiper belt missions supported by centrifugal nuclear thermal propulsion (CNTP) contributes to defining its operational use case. Results indicate that CNTP enables rendezvous missions to the Kuiper belt, both with direct transfer trajectories or planetary gravity assist trajectories, although there are many challenges to making these mission architectures feasible. The direct trajectories have transfer times of roughly 14 to 16 years while combining CNTP with gravity assists from Jupiter could lower transfer time to as low as 10 to 12 years to Kuiper belt objects such as Pluto and Quaoar. These missions are then shown to inform the architecture of the CNTP injection stage vehicle, which can be supported by heavy and super-heavy commercial launch vehicles with a single launch. Last, drawbacks of the mission and vehicle architectures are given that impose limits on the use case for CNTP on these missions.

Analytic Model of Capture Probability for Salvo of Two Electromagnetic Launched Anti Torpedo Torpedoes

To pre-estimate the capture probability of electromagnetic coil salvo two anti-torpedo torpedoes (ATTs), an analytic model is proposed. Based on the analysis of the influencing factors, including the target position dispersion of the incoming torpedo, the entry point’s dispersion of the electromagnetic launched ATT, the underwater tracking speed and heading error of the ATT, and the differences between the salvo and single launch mode, a single integral analytic model of the capture probability is established. The course errors of the ATTs are comprehensively calculated utilizing search model, and the search boundaries of the two ATTs are calculated using the geometric method, taking the optimal searching course of the virtual single launch ATT as the desired searching course of the parallel salvo of the two ATTs. The calculation results of the proposed analytic model are consistent with the statistical results of the Monte Carlo approach through simulation comparison and analysis. The proposed analytic model’s consequences for decision-making and effect assessment are discussed.

Cost estimation for launch vehicle families considering uncertain market scenarios

When optimizing the staging of a launch vehicle, usually a reference mission with a fixed payload mass is used to determine the specific launch cost (or a surrogate thereof, e.g. total mass or dry mass). While this approach neglects suboptimal loading, it delivers sufficiently good results for the evaluation of an individual launcher if the design mission is well known. However, this approach does not capture the full range of payloads that a launch vehicle is expected to launch during its operational lifetime. In order to estimate the cost and subsequently optimize the staging of such launch vehicle families, a full launch scenario has to be considered. The cost of an individual launch vehicle loses importance and the deciding factor is the total cost associated with launching the specified scenario. However, estimates of the future market inevitably include large uncertainties. These uncertainties and their impact are ideally considered and quantified when assessing the specific recurring cost of a single launch vehicle. The consideration of a whole launch market scenario is especially relevant when evaluating a family of launch vehicles, an option currently being discussed for future European launchers. In this case, the distribution of the payloads onto the different members of the launch vehicle family has to be optimized in order to reach the lowest possible cost. Within this manuscript, a method that allows the determination of the recurring cost of such a launcher family based on a launch market scenario including uncertainties is proposed. Based on the launch numbers derived from the combinatorial optimization, parametric cost models for recurring and non-recurring cost are adapted and applied to three possible future launch vehicle families.

A Single-Launch Deployment Strategy for Lunar Constellations

Satellite constellations can provide communication and navigation services to support future lunar missions, and are attracting growing interest from both the scientific community and industry. The deployment of satellites in orbital planes that can have significantly different inclinations and right ascension of the ascending node requires dedicated launches and represents a non-trivial issue for lunar constellations, due to the complexity and low accessibility of launches to the Moon. In this work, a strategy to deploy multiple satellites in different orbital planes around the Moon in a single launch is examined. The launch vehicle moves along a conventional lunar escape trajectory, with parameters selected to take advantage of gravity-braking upon encountering the Moon. A maneuver at the periselenium allows the transfer of the spacecraft along a trajectory converging to the equilibrium region about the Earth–Moon libration point L1, where the satellites are deployed. Providing a small ΔV, each satellite is transferred into a low-energy trajectory with the desired inclination, right ascension of the ascending node, and periselenium radius. A final maneuver, if required, allows the adjustment of the semimajor axis and the eccentricity. The method is verified using numerical integration using high-fidelity orbit propagators. The results indicate that the deployment could be accomplished within one sidereal month with a modest ΔV budget.

Launch parameters of a lunar ice payload traveling via electromagnetic mass accelerator

This paper describes a solution for procuring propellent needed for future space exploration missions based on the Lunar Electromagnetic Mass Accelerator (LEMMA) concept analyzed by McNab and McGlasson. This study examined the feasibility of using an electromagnetic launcher (EML) to transport raw materials used in propellent production from the lunar south pole to the National Aeronautics and Space Administration's (NASA) Lunar Gateway. The proposed Gateway space station, to be located in a lunar near-rectilinear halo orbit (NRHO), is a critical part of NASA's Artemis program. Cheaply and efficiently sourcing lunar hydrogen from surface ice to the Gateway would benefit the program's success and future exploration of the solar system. Our study of the launch requirements for a lunar EML payload are described here. AGI Inc's Systems Tool Kit (STK) was used to calculate the required launch azimuth, elevation, magnitude, epoch, and trip duration for a payload launched from the Moon to rendezvous with Gateway. The study shows that it is feasible to conduct a single launch from the lunar south pole that can intercept any point along the Gateway's orbit with variable launch conditions.

Research on Loss and Temperature Rise of Pulse Inductor for Electromagnetic Rail Launch

The copper foil pulse inductor used for electromagnetic rail launch works in instantaneous super-high current conditions, and the copper foil needs to bear high current density, which also affects the eddy current distribution in the metal shell. The loss and temperature rise generated in the copper foil and the shell affect the efficiency and operation stability of the pulse inductor, and are important performance indicators to test whether the pulse inductor can achieve rapid and continuous pulse discharge. The magneto-thermal coupling simulation model was built by the finite element method (FEM) and boundary element method (BEM), and the numerical calculation results of the loss and temperature rise of the copper foil and shell of the pulse inductor were obtained when the electromagnetic rail launched. Aiming at the problem of large shell loss, a structural optimization scheme is proposed to increase the gap between the copper foil and the shell and reduce the wall thickness of the shell. The simulation shows that after the structure is optimized, the copper foil loss slightly increases by 1.29% during a single launch, but the shell loss decreases by 21.33%, at the same time, the temperature rise of copper foil and the strength of the shell meet the safety requirements of <inline-formula xmlns:mml="http://www.w3.org/1998/Math/MathML" xmlns:xlink="http://www.w3.org/1999/xlink"> <tex-math notation="LaTeX">$6\times $ </tex-math></inline-formula> consecutive launches. Finally, the accuracy of the simulation calculation model and the scheme to reduce eddy current loss of shell are verified by experimental measurement.

Effect of active cooling on the thermal effect of electromagnetic rail under single launch mode

In order to investigate the influence of active cooling on the rail thermal effect of electromagnetic gun under single launch mode, characteristics of the peak temperature, cross-section current density and temperature distribution, axial temperature distribution of rail with and without active cooling are simulated. In addition, the time-varying characteristics of the rail axial temperature distribution, the temperature time-varying characteristic in different temperature parts on the rail and the influence of the cooling water flow velocity on the temperature in different temperature parts are experimentally studied. The results show that during the launch, there is almost no difference between the rail peak temperature and its location with and without active cooling, and the peak temperature is basically located at the initial position of the armature. Without active cooling, the rail axial temperature distribution shows a trend of higher at the gun tail and lower at the gun muzzle. The current density and temperature distribution of rail cross-section show a trend of higher on the rail inner side and lower on the rail outer side, and the diffusion of heat lags behind the diffusion of current. Under the action of active cooling, the rail peak temperature will move to the muzzle along the armature moving direction, and there is a “thermal inversion” phenomenon. The rail cross-section temperature drops faster, and the lowest temperature region appears near the middle cooling channel. Active cooling has different effects on rail temperature in different temperature parts. In the high-temperature part, the rail temperature decreases monotonically with the increase of cooling water flow velocity, increasing the water velocity can accelerate the decline of rail temperature. In the middle-temperature parts, the variation trend of rail temperature is different with different water flow velocities. At low water velocity, the rail temperature increases gradually. When the water flow velocity is greater than 0.75 m/s, the phenomenon of “thermal inversion” will occur. i.e., the rail temperature increases first and then decreases. In the low-temperature part, “thermal inversion” will occur at different flow velocities, that is, the rail temperature increases first and then decreases. When the water flow velocity is greater than 1.0 m/s, the improvement of rail cooling effect will no longer be obvious. It is suggested that the appropriate water flow velocity is 0.75–1.0 m/s under the present experimental conditions.

Manned Mars Mission Analysis Using Mission Architecture Matrix Method

With the development of deep space exploration technology, manned missions to Mars are expected to be realized in the near future. However, the journey to Mars requires more supplies and fuel than near-Earth missions, including moon landings. Using the traditional Apollo-style mission method will result in the spacecraft reaching the LEO mass of 2000 tons, which is not easy to achieve. The use of the modular design method and multiple launches will significantly reduce the mass of a single launch. In addition, the use of nuclear power engines (Nuclear Thermal Propulsion, NTP, and Nuclear Electric Propulsion, NEP) can greatly improve propulsion efficiency, reducing the mass of the propellant. This paper uses the Mission Architecture Matrix (MAM) method to concisely and precisely analyze a series of mission architectures in the case of impulse maneuver transfer and low-thrust transfer. The results show that when only chemical propulsion (specific impulse is 440 s) is used, the maximum LEO launch mass in the optimal mission architecture is 325 tons, and the total LEO mass of the system is 1142 tons. With the usage of NTP (specific impulse is 900 s) and NEP (the specific impulse is 6000 s) technology, the maximum LEO launch mass in the optimal mission architecture is only 85 tons, and the total LEO mass of the system is only about 400 tons. Considering the current rocket technology, the total cost is about USD 1149 million US.

Design of 3-D trajectory sequences for multiple asteroid flyby missions

Prospecting is a necessary pre-requisite of future asteroid mining ventures. It is generally assumed that inspection for the purpose of identifying the asteroid composition can be effectively accomplished from a distance through remote sensing. To be carried out in a timely and economically viable way, prospecting is best performed by devising trajectories such that a single spacecraft manages to fly by as many asteroid as possible, yet seeking to minimize a cost function that we assume to be coincident with propellant consumption. In this paper, we present a method to identify trajectory sequences to multiple asteroids. We restrict our analysis to Near-Earth Asteroids (NEAs), i.e.,, those with perihelion at less than 1.3 AU from the Sun, focusing on Apollo class NEAs only. The volume of space where encounter seeking takes place is a torus-shaped 3-D region in the proximity of the ecliptic. Under the assumption of using impulsive maneuvers to connect ballistic coast arcs, we show that a deterministic building blocks approach is successful in finding a significant number of multi-flyby mission profiles with the desired characteristics. Using this scheme, it is possible to envisage realistic asteroid prospecting missions using a single launch to deploy a number of small spacecraft, with tens—or possibly hundreds—of asteroids visited in a few years.

Investigation on the Spatial-Temporal Distribution of Electromagnetic Gun Rail Temperature in Single and Continuous Launch Modes

In order to study the spatial-temporal distribution characteristics of electromagnetic rail gun temperature, based on moving grid processing method and mod function method, a 3-D geometric model of electromagnetic rail gun under single launch mode, a 2-D geometric model of electromagnetic rail gun under continuous-fire mode, and the corresponding transient electromagnetic field-temperature field coupling model are established by using the finite element analysis software COMSOL Multiphysics, and the temperature spatial-temporal distribution characteristics in the rail-armature contact area, cross section of rail, and along the rail axis is simulated under single launch and continuous launch. The results show that, under the conditions of single launch and natural cooling, the high temperature area in the rail-armature contact surface is located at the rear of the armature due to the skin effect. The temperature distribution along the axial inner side of rail tends to be higher near the tail and lower near the muzzle, it increases with time, but the peak temperature does not exceed the melting point of copper. Under the continuous launch and natural cooling, after five launches, the peak temperature of the rail increases to 709 °C, which is close to the melting point of copper. Continued launching will cause rapid accumulation of heat in the rail and a rapid increase of peak temperature, resulting in shortened life of the launcher and even launch failure. Therefore, without active cooling measures, the electromagnetic rail gun cannot be used in actual combat. The results of the study are an important guide for the design and thermal management of the cooling system of electromagnetic rail gun.

Optimal staging of reusable launch vehicles for minimum life cycle cost

This paper proposes an optimal staging procedure of a reusable launch vehicle considering its life cycle operations.Instead of minimizing the lift-off mass of a single launch adopted by traditional optimal sizing problems, the proposed procedure minimizes the total life cycle cost of reusable stages using an integrated optimization module and a cost model. The recurring/nonrecurring cost models are established to compute the life cycle cost of a reusable launch vehicle throughout the operational period of its reusable stage. The integrated optimization module obtains a vehicle configuration and ascent/descent trajectories to put the desired payload into the target orbit while minimizing the expendable stage mass for a given reusable stage. The validity and effectiveness of the proposed procedure are demonstrated through an optimal staging case study for a real reusable launch vehicle, Falcon 9.

Effect of Active Cooling on the Thermal Effect of Electromagnetic Rail Under Single Launch Mode

Effect of Active Cooling on the Thermal Effect of Electromagnetic Rail Under Single Launch Mode

Developing novel multi-plane satellite constellation deployment methods using the concept of nodal precession

• Multi-plane satellite constellations can be deployed using J 2 perturbation effects. • Two methods are proposed to deploy a satellite constellation with a single launch. • Deployment methods have been optimized multi-objectively in terms of time and required velocity change. Deployment of multi-plane satellite constellations has become a fresh challenge in the space program missions since all the direct solutions including plane changing maneuvers and separate launches are energy and time-consuming. A proposed solution to this challenge is launching several satellites simultaneously and deploying them using gravitational perturbations of the Earth. Utilizing this solutions, only in-plane satellite maneuvers are sufficient to deploy a constellation of satellites in several orbital planes. The aim of this study is to extend and develop the present methods that use the concept of nodal precession to deploy a multi-plane satellite constellation with only one launch in order to cover a wider range of mission specifications, provide mathematical expressions for the total deployment time and the fuel expenditure of these methods, optimize the methods multi-objectively and provide a thorough analysis and comparison between them. Two general methods are proposed in this study; one that places all the satellites in a parking orbit and injects them into their final orbits sequentially, and one that first places them in different drifting orbits and after a certain amount of time, maneuvers them to their final orbits. The findings of the optimization process suggest that the method which involves all the satellites in the process of deployment simultaneously and by this means exploits the full nodal precessing potential of all of them provides better results in comparison with the method with a more passive approach, which considers the deployment a sequential process. Finally, since this mathematical modeling is performed for the first time in the literature, the mathematical expressions for both methods are analyzed to provide beneficial insights for the designers.

Statistical Characterization of Temperature and Pressure Vertical Profiles for the Analysis of Laser Heterodyne Radiometry Data.

The statistical analysis of historic pressure and temperature profiles from radiosonde launches for use in the fitting of molecular oxygen line shapes is presented. As the O2 mixing ratio is nearly constant throughout the lower atmosphere, only variations in pressure and temperature profiles will affect the fit of observed O2 features in Laser Heterodyne Radiometry (LHR) spectra. Radiosonde temperature and pressure data are extracted from the Integrated Global Radiosonde Archive (IGRA) for a given station, date, and launch time. Data may be extracted for a single launch, for the same date over several years, and/or within a window centered on a target date. The temperature and pressure profiles are further characterized by the statistical variation in coefficients of polynomial fits in altitude. The properties of the probability distributions for each coefficient are used to constrain fits of O2 line shapes through Nelder–Mead optimization. The refined temperature and pressure profiles are then used in the retrieval of vertically resolved mixing ratios for greenhouse gases (GHGs) measured in the same instrument. In continuous collections, each vertical profile determination may be treated as a Bayesian prior to inform subsequent measurements and provide an estimate of uncertainties.

A Hummingbird in Space

Two distinct trends are apparent in the design and planning of satellite missions. Until the late 1990s, multibillion-dollar space programs centered on large satellites, such as Envisat [1], promised to provide a common platform to support a variety of co-located sensing equipment. A reduction in cost was expected, as several instruments shared a single bus and a single launch. These benefits did not materialize due to the rise of a plethora of engineering and scheduling problems: electromagnetic incompatibilities between diverse technologies; instruments inducing vibrations on the platform that affect other equipment; and deployment-ready instruments waiting for other equipment in earlier development stages. As a reaction to these issues, the second trend where programs based on single-instrument satellites of much smaller sizes and mass began to emerge, eventually leading to the deployment of space devices that nowadays we call small satellites [11].

From the Martian Surface to Its Low Orbit in a Reusable Single-Stage Vehicle—Charon

With Mars colonisation becoming a reality for the near future, it is of importance to analyse how crew and cargo can be transported between Earth and a colony on Mars. This article is a feasibility and design study of a launch vehicle whose mission is to shuttle crew and cargo from Low Mars Orbit to a colony on the Martian surface. A single-stage reusable rocket has been selected to fulfil this mission, code-named Charon. The mission profile of such a vehicle was created, leading to a Maximum Growth Allowance (MGA) Delta-V budget of 6.2 km/s. With the mission profile in mind, each subsystem underwent a preliminary design. With reliability and maintainability in mind, subsystems were designed for redundancy and modularity, and an abort system was included for an added level of safety. The iterative design process resulted in a vehicle with a MGA mass of 198.7 tons, capable of transporting 1200 kg of cargo and a crew of 6 people to a 500 km orbit and back. The preliminary design of the vehicle is deemed safe. Following a fault tree analysis, the Single Launch Loss of Mission, Vehicle and Crew (SL-LOM, SL-LOV, SL-LOC) probabilities are computed to be of 0.975%, 0.12%, and 0.079%. Finally, from the vehicle’s constraints on the base, the feasibility of the project has been reflected upon. It is deemed that such a concept is of high interest only when the base is already operational, due to the launch and maintenance infrastructure that it requires, as well as the power it requires from the Martian base.

On-Line Interpretation and Real-Time Diagnosis of Rocket’s Single Equipment

The work state of a launch vehicle is generally interpreted automatically on software. However, the sheer number of target parameters makes it difficult to realize real-time interpretation, and abnormal interpretation result does not necessarily mean that the vehicle is in abnormal state. This paper introduces the edge computing to achieve on-line interpretation and real-time diagnosis of a single launch vehicle. Firstly, the parameters to be interpreted were subjected to thresholding, leaving only those with high interpretation value. Next, the interpretation server layer of the real-time diagnosis model was built based on the attribute and value reduction algorithm of variable precision rough set (VPRS). Moreover, the higher-grade criteria were written in criterion modeling language (CML) and used to interpret the various higher-grade interpretation data pushed by the edge layer in real time. On this basis, the outputs of the edge layer and interpretation server layer were integrated to achieve the real-time diagnosis of single vehicle faults. Finally, the proposed model was proved feasible through the application in a launch vehicle.

Venus Flagship Mission Concept: A Decadal Survey Study.

More than any other known planet, Venus is essential to our understanding of the evolution and habitability of Earth-size planets throughout the galaxy. We address two critical questions for planetary science: 1) How, if at all, did Venus evolve through a habitable phase? 2) What circumstances affect how volatiles shape habitable worlds? Volatile elements have a strong influence on the evolutionary paths of rocky bodies and are critical to understanding solar system evolution. It is clear that Venus experienced a different volatile element history from the Earth and provides the only accessible example of one end-state of habitable Earth-size planets. Venus will allow us to identify the mechanisms that operate together to produce and maintain habitable worlds like our own. The (VFM) concept architecture relies on five collaborative platforms: an Orbiter, Lander, variable-altitude Aerobot and two Small Satellites (SmallSats) delivered via a single launch on a Falcon 9 heavy expendable. The platforms would use multiple instruments to measure the exosphere, atmosphere and surface at multiple scales with high precision and over time. VFM would provide the first measurements of mineralogy and geochemistry of tessera terrain to examine rocks considered to be among the most likely to have formed in a habitable climate regime. Landed, descent, aerial and orbital platforms would work synergistically to measure the chemical composition of the atmosphere including the Aerobot operating for 60 days in the Venus clouds. Loss mechanisms would be constrained by the SmallSats in two key orbits. The baseline payload for VFM includes instruments to make the first measurements of seismicity and remanent magnetism, the first long-lived (60 day) surface platform and the first life detection instrument at Venus to interrogate what could be an inhabited world. The VFM concept directly addresses each of the three Venus Exploration Analysis Group (VEXAG) goals as well as several of the strategic objectives of the 2020 NASA Science Plan, Planetary Science Division, Heliophysics and Astrophysics. The simultaneous, synergistic measurements of the solid body, surface, atmosphere and space environment provided by the VFM would allow us to target the most accessible Earth-size planet in our galaxy, and gain a profound new understanding of the evolution of our solar system and habitable worlds.

Unsteady overflow behavior of polydisperse granular flows against closed type barrier

The overflow behavior of polydisperse granular flows against closed type barrier is a crucial yet poorly understood aspect of multiple-barrier design. We use a robust numerical model based on the three-dimensional discrete element method (DEM) to simulate polydisperse granular flows and identify fundamental mechanisms. The morphology of the dead zone behind the upstream barrier significantly influences the overflow behavior of granular flows. This implies that aside from the case of a gentler barrier (<70°), the launch direction is controlled by the inclination surface of the dead zone. We identified a three-stage evolution process of the overflow velocity. The velocity component in the z-direction controls the first stage, whereas the latter two stages are controlled by the velocity component in the x-direction. The launch angle is usually larger than 45° during the initial overflow stage. Such a high launch angle is maintained for about 0.2–0.3 s. Boulders are transported downstream by the overflow process owing to particle-size segregation and a more concentrated load is received by the downstream barrier, which can lead to localized barrier damage. A different overflow mechanism would result in a different launch distance. Steeper (135°) and gentler (70°) barrier configurations are dangerous because the normalized launch distance is up to 20% larger than that of a barrier with a medium inclination. And a point-mass based assumption is used to estimate the maximum launch distance with the use of maximum overflow velocity and a single launch angle. In addition, we discuss some limitations of numerical modelling, including the effect of fluid phase on overflow, as well as the direct application of our results in engineering design.

ARPIR: automatic RNA-Seq pipelines with interactive report

BackgroundRNA-Seq is an increasing used methodology to study either coding and non-coding RNA expression. There are many software tools available for each phase of the RNA-Seq analysis and each of them uses different algorithms. Furthermore, the analysis consists of several steps regarding alignment (primary-analysis), quantification, differential analysis (secondary-analysis) and any tertiary-analysis and can therefore be time-consuming to deal with each step separately, in addition to requiring a computer knowledge. For this reason, the development of an automated pipeline that allows the entire analysis to be managed through a single initial command and that is easy to use even for those without computer skills can be useful. Faced with the vast availability of RNA-Seq analysis tools, it is first of all necessary to select a limited number of pipelines to include. For this purpose, we compared eight pipelines obtained by combining the most used tools and for each one we evaluated peak of RAM, time, sensitivity and specificity.ResultsThe pipeline with shorter times, lower consumption of RAM and higher sensitivity is the one consisting in HISAT2 for alignment, featureCounts for quantification and edgeR for differential analysis. Here, we developed ARPIR, an automated pipeline that recurs by default to the cited pipeline, but it also allows to choose, between different tools, those of the pipelines having the best performances.ConclusionsARPIR allows the analysis of RNA-Seq data from groups undergoing different treatment allowing multiple comparisons in a single launch and can be used either for paired-end or single-end analysis. All the required prerequisites can be installed via a configuration script and the analysis can be launched via a graphical interface or by a template script. In addition, ARPIR makes a final tertiary-analysis that includes a Gene Ontology and Pathway analysis. The results can be viewed in an interactive Shiny App and exported in a report (pdf, word or html formats). ARPIR is an efficient and easy-to-use tool for RNA-Seq analysis from quality control to Pathway analysis that allows you to choose between different pipelines.