ABSTRACT. To realize space elevator, there are lot of technical issues to overcome, and they must be solved and verified step by step. Among them, some technologies, such as fiber formation of carbon nanotube, wireless transmission of energy to climber, etc., are still immature, but some of them can be realize at present technology and they must be verified in space. In our university, we try to verify two basic technologies of space elevator by using Microsatellites and obtain data for future design, one is the tether (cable) deployment technology, and other is the climber operation along tether in space. A tether deployment is performed by a CubeSat called ‘STARS-C (Space Tethered Autonomous Robotic Satellite - Cube)’ which will be carried by HTV and released from the Japanese experimental module Kibo on ISS in 2016. STARS-C consists of Mother Satellite (MS), Daughter Satellite (DS) and 100m tether between them, and its mission is focused on deploying tether to study the tether dynamics during the tether deployment for designing smooth tether deployment possible in future tether mission including space elevator. MS and DS have the common subsystems as satellite such as power system, communication system, and also have the tether unit with spool and reel mechanism as mission system to control tether tension. STARS-C is in the FM phase at present (autumn in 2015). The detail design and plan of STARS-C experiment will be presented at the conference. We also have been designed the next-step Microsatellite called ‘STARS-E’ (Space Tethered Autonomous Robotic Satellite - Elevator) under the Grant-in-Aid for Scientific Research. STARS-E is the satellite of 500 mm size to verify the climber operation in space, and consists of MS, DS, 1000 m tether between them, and a climber moving along tether. The outline design of STARS-E also will be presented at the conference.

ABSTRACT. The space elevator system which utilizes an extended cable from a geostationary orbit station towards ground, was suggested by Artsutanov and Isaacs, over half a century ago. To understand the feasibility of a space elevator, researchers have studied whether an extremely long structure, 50,000 km to 100,000 km long, stretching from the Earth’s surface into space can be stably maintained in Earth’s gravitational field which changes according to distance from the Earth. Some preliminary studies have already been carried out. A further area of study is whether it is possible to develop an extremely long and stable cable. This has already been discussed by both domestic and foreign researchers, and there appears to be a small difference of opinion concerning some areas. There is also the issue of what influence the elevator car, called the “climber,” would have on cable stability. After much delay, studies regarding this issue have finally started. It is no exaggeration to say that fundamental studies to consider the appropriate mechanism for use as a climber on a space elevator have only just started. However, the significance for the development of a space elevator is found simply in its superiority to that of current space rocket technology. Based on this, development of climber models have started with the application of existing technology, and preliminary experiments are now being conducted at ground level here on Earth. The development of a climber is significant because of the advantageous economical benefits and transportation efficiency when compared to existing rockets. Therefore, the design requirements for the climber include a great importance on these two aspects, and the feasibility was examined in a report by the International Academy of Astronautics and Edward’s research. However, the proof experiments and quality assessments of the climber’s drive mechanisms shown by those conceptual designs are now able to be tested. In this study, we will discuss the fundamental design requirements, including both functions and performance that a partial space elevator climber should have, as well as present prototype models and design proposals. In addition, a basic experiment about the friction drive mechanism for a tether of the partial space elevator was conducted. Also, a self-driving type climber was prototyped and its performance was evaluated.

ABSTRACT. Understanding of cable dynamics at the construction of space elevator is one of the most important issues to design space elevator. There are some studies about the cable dynamics at the construction of space elevator in the past, but, they assumed the cable deployment with rising the space station to higher orbit more than GEO keeping the center of the orbit in GEO. Such method has disadvantage that large propellant is necessary to rise the space station to higher orbit. We propose a new method to construct the space elevator in which cables are simultaneously deployed upward and downward from the main space station with keeping it in GEO and balancing upward and downward cables. Such a method will have a possibility to reduce their total propellant mass at the space-elevator construction. The analysis is performed by the originally developed cable dynamics model in our group. The model is a two-dimension lamped mass model, in which tether is modeled by mass points, and these mass pints are connected each other by springs and damper and are pull out per assumed length. In this study, the cable dynamics and the cable stress are analyzed for assumed deployment conditions of the cables at first, then the total impulse during cable deployment, that is necessary to keep the main station in GEO and to control the cable deployment speed to does not over the maximum stress of the cable, is obtained. The results are also compared with the preceding method for the same conditions. The initial result shows that the total impulse that is necessary to control cable dynamics with keeping GEO station in GEO in our cable-deployment method can be decreased about one-third compared with that is necessary to rise the main station to higher orbit and to control cable dynamics in preceding cable-deployment method. The detail analytical results will be presented at the conference.

ABSTRACT. One proposed technique to enhance current collection from the ambient space environment is charged particle emission from a hollow cathode plasma plume. Simulations suggest this technique is possible given sufficient hollow cathode discharge current, plasma plume expansion, and spacecraft potential. Applications for hollow cathode based particle emission include electrodynamic tether current enhancement, reduction of potential transient magnitude for rapid charging events, and remediation of high power electric propulsion system generated spacecraft charging effects. Experiments were performed in the Large Vacuum Test Facility (LVTF) at the University of Michigan’s Plasmadynamics and Electric Propulsion Laboratory (PEPL) to determine the efficacy and key parameters of this charge emission technique. For these experiments, an electrically isolated “spacecraft” was simulated. “Spacecraft” potential was measured for transient and steady state conditions using various representative “collected” currents while operating the hollow cathode. Hollow cathode parameters are varied to determine their effect on charge emission and collection efficiency. Discharge current and ionization fraction are found to be the key cathode parameters as suggested via theory/simulation. Current equilibrium isn’t reached until a few seconds after current collection is initiated, far exceeding the characteristic reaction time of the plasma plume. Possible system-based explanations for these long transients are presented. The presence of microdischarges within the chamber is examined for high “spacecraft” and plasma potentials.

ABSTRACT. In 1924, Irvin Langmuir and H. M. Mott-Smith published a theoretical model for the complex plasma sheath phenomenon in which they identified some very special cases which greatly simplified the sheath and allowed a closed solution to the problem. The most widely used application is for an electrostatic, or “Langmuir,” probe in laboratory plasma. Although the Langmuir probe is physically simple (a biased wire) the theory describing its functional behavior and its current-voltage characteristic is extremely complex and, accordingly, a number of assumptions and approximations are used in the LMS model. These simplifications, correspondingly, place limits on the model’s range of application. Adapting the LMS model to real-life conditions is the subject of numerous papers and dissertations. The Orbit-Motion Limited (OML) model that is widely used today is one of these adaptions that is a convenient means of calculating sheath effects. The OML equation for electron current collection by a positively biased body is simply: I≅A_ j_eo 2/√π (Φ)^(1/2) where A is the area of the body and Φ is the electric potential on the body with respect to the plasma. Since the Langmuir probe is a simple biased wire immersed in plasma, it is particularly tempting to use the OML equation in calculating the characteristics of the long, highly biased wires of an Electric Sail in the solar wind plasma. However, in order to arrive at the OML equation, a number of additional simplifying assumptions and approximations (beyond those made by Langmuir—Mott-Smith) are necessary. The OML equation is a good approximation when all conditions are met, but it would appear that the Electric Sail problem lies outside of the limits of applicability.

ABSTRACT. There is no validated and accepted theory for a Langmuir probe in the regime of electron collection from a magnetized drifting plasma. The Space Shuttle experiment, TSS-1R, cleared the field of many speculative models and provided data for future model development. Still, there have been few or no subsequent experiments, theories or simulations to independently validate the surviving models. We have chosen to renew investigation into this problem for reasons of both probe theory and its relevance to spacecraft charging/plasma-interactions, and to investigate the electromagnetic momentum coupling between the probe and background plasma. Although the Low earth Orbit probe interaction is expected to be highly electrostatic, the electromagnetic beginnings of research into this problem [Drell et.al., 1965] have led to widely accepted models of momentum coupling that can be challenged or validated as part of a new and combined investigation. Towards these ends, we present the results of electromagnetic particle-in-cell simulations of an electron collecting probe in magnetized ExB drifting plasma. These simulations were run over a broad range of plasma parameters, from under-dense strongly magnetized plasma (ω_ce > ω_pe, r_probe > r_ion) to over-dense plasma (ω_pe > ω_ce, r_probe > r_e). Our results detail the current-voltage curve over a range of plasma parameters and drift velocity, the dynamics of non-neutral current wings, and plasma heating. We also describe the development of a new EXB drifting magnetized plasma chamber.

ABSTRACT. For thrust to be produced using an electrodynamic tether, the current through the tether must be maintained using an appropriate configuration of electron and ion emission or collection at each end. The current is governed through plasma-spacecraft interactions which varies depending on the ambient space-environment. To improve our understanding of some of these interactions, initial experiments were performed at the University of Michigan’s Plasmadynamics and Electric Propulsion Laboratory (PEPL). A plasma was emitted from a simulated spacecraft while the system was charged using a power supply to study the plasma’s transient and steady state characteristics. In the steady state, plasma characteristics such as plasma potential and electron temperature are dependent on the spacecraft bias. Ion and electron currents were collected using Langmuir probes during transient conditions. The collected ion currents increased while the electron currents decreased suggesting the emission of ions from the plume and collection of electrons by the spacecraft.

ABSTRACT. *** coming soon ***

ABSTRACT. The miniaturization of spacecraft, facilitated by an explosion of micro-electronics in the past two decades, reduces useable volume and surface area as well as increases mass constraints. Additionally, power systems such as solar panels or batteries must scale accordingly, reducing the available power or energy storage capabilities. These limitations impose the challenge of scaling existing technology or developing new technology that will operate within the constraints.

Small-scale electrodynamic tether (EDT) systems (on the order of 10s of kilograms or less total system mass) require appropriately-scaled plasma-interface technologies (i.e., plasms contactors) that facilitate sufficient levels of current for electrodynamic operations such as deorbiting, propulsion, or energy harvesting. We present a survey of existing technology with their constraints, relevant characterization parameters, resource consumption, and projected future capabilities. We then identify configurations suitable for small spacecraft and identify size limitations for scaling EDT systems that use state-of-the-art contactors. This research indicates how future advances in technology would impact the future minimum size of EDT systems.