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After over 60 years of development, the space propulsion field in China has made tremendous progress in terms of technology system and product pedigree, and has supported major engineering applications including manned spaceflight, applied satellites, and deep space exploration. However, considering the requirements of future major aerospace missions in China, the technical level of current space propulsion systems is still insufficient, and it is urgent to plan directions for the innovative development and breakthroughs of the field. This study reviews the development status of the space propulsion systems from the aspects of electric, chemical, nuclear, and new-concept space propulsion, and analyzes the requirements of future applications. Specifically, high-performance, low-cost space propulsion systems are required to support the networking of low-Earth-orbit (LEO) small satellites; high-thrust, reusable low-temperature chemical propulsion technologies are required for new space-transportation systems; and multi-type, long-life space propulsion technologies are required for deep-space applications. The results demonstrate that the space propulsion field in China still faces multiple challenges. First, the system reliability and lifespan require further improvement. Second, the space propulsion systems of China still lag behind the advanced international level in terms of product maturity of some technologies. Third, this field typically has a high product cost. Fourth, there are few types of optional technologies, and the research progress of key technologies is lagging. Therefore, we propose the following suggestions: (1) optimizing technology options through top-level planning, (2) improving the development mechanism of the field by emphasizing basic research, and (3) clarifying and promoting scientific research and development in key technological directions, thereby achieving in-orbit applications of the propulsion systems for LEO small satellites, Earth–Moon space transfer, and deep space exploration (including manned spaceflight).

On October 11, 2018, in the Ulytau region of the Republic of Kazakhstan, the Soyuz-FG launch vehicle carrying a crewed MS-10 spacecraft failed, creating two areas where rocket propellants spilled and soil condition was monitored in 2018, 2019, 2022 and 2023. This article presents data on the content of pollutants, namely unsymmetrical dimethylhydrazine (UDMH), nitrosodimethylamine (NDMA) and total petroleum hydrocarbons (TPH), total N, organic carbon, exchangeable Ca and Mg, water-soluble NO3–, NO2–, HCO3–, CO32–, SO42–, Cl–, K+, Na+, pH values, cation exchange capacity and electrical conductivity of the water extract in disturbed and background Aridisols (more than 200 samples in total). This data set contains information on interseasonal (autumn 2022 and spring 2023) differences in the content and vertical differentiation of some soil properties in Aridisols in Central Kazakhstan. In autumn, the content of TPH, water-soluble Cl– and SO42– and alkalinity from CO32– is 1.4, 235, 201, and 2 times higher, respectively, and the content of total N and water-soluble NO3– and NO2–, alkalinity from HCO3– is 2.4, 1.4, 6.4 and 1.9 times lower, respectively (p < 0.05). In spring and autumn, the content of exchangeable Ca and Mg, cation exchange capacity did not differ significantly. The presented materials can be used to optimize restoration of disturbed arid ecosystems and future monitoring work at sites of regular landing of the first stages and emergency crash sites of launch vehicles.

This paper presents the status of developments worldwide regarding use of hybrid rocket motors for space transportation. Historical roots are presented and reasons for revisiting hybrid technology after a few decades of limited interest are examined. Modern developments in sounding rockets, reusable suborbital systems and launch vehicles are discussed with particular focus on propellant technology. Various propellant combinations include use of liquid oxygen, hydrogen peroxide, nitrous oxide and nitrous oxide-oxygen mixtures as oxidizers. Different fuels are considered, taking into account performance, as well as inter alia obtainable regression rates. Results of preliminary calculations for vehicles using different propellant combinations are presented and analysed. This is compared with proposed configurations of hybrid rockets worldwide. Unresolved problems and several unknowns are pointed out, including hybrid rocket motor scalability issues, large motor combustion instabilities, combustion efficiency of metalized fuels, propellant volumetric performance and mass of fuel residuals in case of wagon wheel grain geometry. It is discussed whether new-space hybrid launch vehicles, while typically with limited stage reusability, may be cost-competitive in regard to other chemical rocket propulsion system developments. The paper is summarized with a list of potential future advances and technical opportunities. The main purpose of the conducted research is to provide a comparison between different hybrid propulsion technologies available, or currently under development, worldwide.

Scramjet propulsion offers promise for flexible and sustainable space transportation. Hypersonic airflow compression through the intake plays a major role in successful scramjet-powered ascent. It is thus of crucial importance to design high-performance intakes that can work efficiently and robustly throughout the ascent trajectory. The present research is conducted to gain physical insight into multi-point global optimum scramjet intake design and associated requirements by means of multi-point design optimization and analytical investigation in inviscid and viscous regimes. It has been found and verified theoretically that at any operating conditions, compression efficiency, drag, compression ratio, and mean exit temperature can theoretically be determined uniquely in an interrelated manner under calorically perfect and adiabatic flow assumptions, provided one of them is given or known except for compression efficiency. The multi-point optimization study in the inviscid regime has verified that the multi-point global optimum design can exist with respect to compression efficiency and drag. On the other hand, while the theoretical analysis also prescribes that the global optimum design can exist for a single operating condition in viscous flow, trade-off relations have been found to exist between two operating conditions in the viscous regime, where boundary layer is responsible for counteracting behaviors between compression efficiency and drag at different operating conditions. The insights gained contribute to the guiding principle for intake design of scramjet engines for access-to-space. • Scramjet intake design has been optimized for multiple operating conditions. • Existence of global optimum intake designs has been verified in inviscid regime. • Flow viscosity has been found to cause trade-off relations in intake performance. • Analytical approaches to determine various performance parameters have been derived. [ABSTRACT FROM AUTHOR]

This article presents the full operational experimental capabilities of the plasma wind tunnel facilities at the Institute of Space Systems at the University of Stuttgart. The simulation of the aerothermodynamic environment experienced by vehicles entering the atmosphere of Earth is attempted using three different facilities. Utilizing the three different facilities, the recent improvements enable a unique range of flow conditions in relation to other known facilities. Recent performance optimisations are highlighted in this article. Based on the experimental conditions demonstrated a corresponding flight scenario is derived using a ground-to-flight extrapolation approach based on local mass-specific enthalpy, total pressure and boundary layer edge velocity gradient. This shows that the three facilities cover the challenging parts of the aerothermodynamics along the entry trajectory from Low Earth Orbit. Furthermore, the more challenging conditions arising during interplanetary return at altitudes above 70 km are as well covered. [ABSTRACT FROM AUTHOR]

On October 11, 2018, in the Ulytau region of the Republic of Kazakhstan, the Soyuz-FG launch vehicle carrying a crewed MS-10 spacecraft failed, creating two areas where rocket propellants spilled and soil condition was monitored in 2018, 2019, 2022 and 2023. This article presents data on the content of pollutants, namely unsymmetrical dimethylhydrazine (UDMH), nitrosodimethylamine (NDMA) and total petroleum hydrocarbons (TPH), total N, organic carbon, exchangeable Ca and Mg, water-soluble NO 3 -, NO 2 -, HCO 3 -, CO 3 2- , SO 4 2- , Cl - , K + , Na + , pH values, cation exchange capacity and electrical conductivity of the water extract in disturbed and background Aridisols (more than 200 samples in total). This data set contains information on interseasonal (autumn 2022 and spring 2023) differences in the content and vertical differentiation of some soil properties in Aridisols in Central Kazakhstan. In autumn, the content of TPH, water-soluble Cl - and SO 4 2- and alkalinity from CO 3 2- is 1.4, 235, 201, and 2 times higher, respectively, and the content of total N and water-soluble NO 3 - and NO 2 - , alkalinity from HCO 3 - is 2.4, 1.4, 6.4 and 1.9 times lower, respectively ( p < 0.05). In spring and autumn, the content of exchangeable Ca and Mg, cation exchange capacity did not differ significantly. The presented materials can be used to optimize restoration of disturbed arid ecosystems and future monitoring work at sites of regular landing of the first stages and emergency crash sites of launch vehicles., (© 2024 The Author(s).)

The technology surrounding missiles that fly at five or more times the speed of sound is not mature, but several nations have tested hypersonic missiles. The envisioned military value of hypersonic delivery systems lies in their ability to strike quickly over long distances while evading early-warning radars and ballistic missile defenses—but some proponents of the technology also argue that hypersonics could replace nuclear weapons in some roles. Critics, meanwhile, see hypersonic missiles as destabilizing: They could be mistaken for ballistic missiles, could be used in attacks against nuclear and other strategic military assets, and could even be outfitted with nuclear warheads. Amid these concerns, calls have emerged to halt a developing hypersonic arms race by enacting a moratorium on hypersonic testing and eventually establishing a test ban treaty. Here, Mark Gubrud of the United States (2015), Rajaram Nagappa of India, and Tong Zhao of China (2015) debate how serious a threat hypersonic missiles pose to global security, and whether a test ban is the best way to moderate such a threat. [ABSTRACT FROM AUTHOR]

Efficient operation is important to make full use of the capabilities of China's space station. Determining the stochastic impacts of emergencies on the operational scenario of the space station is critical for successful implementation. However, few studies have assessed the uncertainties in the operational processes of the space station. To fill this gap, discrete event simulation (DES) is used to develop an evaluation method for the contingent operational plan of a space station. First, DES is used to develop a model framework of the space station operations, and the launch delay of cargo vehicles is introduced into the integrated simulation procedure. Second, the precision of the results and the computational efficiency are improved using the variance reduction technique. The corresponding effect on the number of simulation trials is confirmed using four constraints and three measurable metrics. Finally, the proposed method is applied to a two-year space station operational plan. The results show that maintaining a short interval between the launch date of the cargo vehicles and the start of the launch windows can decrease prolonged duration after a launch delay. A statistical analysis can be used to determine a safe interval between the dates of the events and vehicle launch. [ABSTRACT FROM AUTHOR]

For the $N$ -impulse transfer between two earth orbits, this paper introduces $N - 1$ intermediate orbits to describe the orbit transfer scheme. The eccentricity vector of the intermediate orbits and the true anomalies corresponding to the impulsive points are chosen as the optimization variables. Based on the patched conic theory, candidate solution can be analytically derived, constraints are removed from the optimization model, and the original problem is converted to a parameter optimization problem. The only difficulty lies in the initialization because the number of optimization variables increases linearly with $N$ , which can be very large. This is settled by a hybrid optimization algorithm that comprises two searching methods. The problem is solved first by an improved particle swarm optimization method and, then, by an adaptive conjugate gradient method. The proposed method is adaptive to problems with any finite $N$ and can calculate the optimal $N$ in any transfer scenarios. The simulation validates the proposed method with some well-known cases and demonstrates its adaptation to $N$. [ABSTRACT FROM AUTHOR]

The German Aerospace Center (DLR) is currently developing the Reusability Flight Experiment (ReFEx). The successor of the already performed hypersonic flight experiments SHEFEX I and II shall be launched on a Brazilian VSB-30 sounding rocket in 2022 and shall achieve a re-entry velocity of more than Mach 5. The main goals of the project are the demonstration of a controlled autonomous re-entry flight from hypersonic velocity down to subsonic range and the testing of the key technologies required for future reusable first stage systems. Utilizing Concurrent Engineering (CE) approach the fundamental feasibility of this sophisticated flight experiment has been investigated by the entire ReFEx team. All required systems, including sensors and actuators as well as their interfaces have been defined and different options were assessed regarding matters such as the scientific output, complexity, risk and cost. The current configuration of ReFEx has a re-entry mass of about 400 kg, a length of 2.7 m and a wingspan of 1.1 m. This paper provides a system overview and addresses the Reusable Launch Vehicles (RLV) technologies Guidance, Navigation and Control as well as Flight Instrumentation. Furthermore the mission design (launch & re-entry) and the main challenges regarding the mission realization are addressed. • Perform a controlled flight in the velocity range hypersonic down to subsonic. • Perform a controlled heading change. • Development of key technologies for winged RLV first stages. • Perform In Flight Data acquisition using advanced sensors for RLV research and development purposes. • Recovery of the experimental vehicle. [ABSTRACT FROM AUTHOR]

Copyright of Journal of Harbin Institute of Technology. Social Sciences Edition / Haerbin Gongye Daxue Xuebao. Shehui Kexue Ban is the property of Harbin Institute of Technology and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Increased demand for smallsat launch capabilities drives on the search for low-cost launch vehicles. Hybrid rocket technology has been pushed forward in the recent years and is now ready to meet the demand of this low-cost rocket development. With intrinsic safety and reliability the development and production cost of hybrid rocket engines can be a small fraction of bi-liquid propulsion systems. At the German Aerospace Center DLR a hybrid rocket engine is developed in cooperation with the German start-up company HyImpulse Technologies GmbH in order to advance launcher technology in Germany. A concept for a mini-launcher has been designed by HyImpulse and the development of a 75 kN (SL) hybrid rocket engine is on-going with a test campaign foreseen in 2019. This engine will be the largest hybrid rocket engine ever tested in Europe. It takes advantage of the long development and research history of paraffin-based hybrid rockets at the DLR Lampoldshausen. Since 2010 paraffin-based fuels have been analyzed and improved. Additives have been found to greatly improve the mechanical properties of the paraffin solid fuel. At the same time the regression rate has been modulated to adapt the fuel for use in launcher applications. The concept of the HyImpulse mini-launcher includes eleven engines of this type, which will result in large scale serial production for the foreseen launch rate of twelve launches per year. This further reduces production costs of the hybrid rocket engine. The advantage of paraffin-based fuel lies in the simple geometry of the grain caused by the high regression rate. A cylindrical fuel grain will have a much lower manufacturing cost compared to classical wagon wheel fuel grains with HTPB. Hybrid launchers are often used within pressure fed propulsion systems, but in order to increase the payload of the launcher, this is replaced with a gas generator and turbo pump driven propulsion system for this launcher concept. • Hybrid rocket propulsion allows safe propulsion systems for launch vehicles. • New fuels compensate drawbacks of classical hybrid rocket fuels. • The new space startup HyImpulse develops a smallsat launch vehicle. [ABSTRACT FROM AUTHOR]

To avoid the severe thermal load within the embeddedly configured Rocket-Based Combined Cycle engine, we arranged the rocket engine location in the present study, that is, the rocket engine was separately mounted for the scramjet flowpath. The rocket engine was taken out from the scramjet flowpath and located on the ramp wall of the scramjet external nozzle. In this case, the ramp wall acted as the additional nozzle in the spike nozzle manner for the rocket internal nozzle. One-dimensional analysis showed that there was an optimal expansion ratio favorable for the scramjet external nozzle. Subsequently, more complex analysis was partially performed using Method-of-Characteristics based two-dimensional wave model with some novel modifications, which can express the pressure mismatching between the exhaust and the ambient flows. The twodimensional analysis showed that thrust production within the scramjet external nozzle could become much lower due to impingement of the cowl lip expansion waves resulting from the pressure mismatching. This pressure change also sizably reduced the thrust performance of the rocket spike nozzle. It was, however, also demonstrated that presently analyzed nozzle system has great potential to drastically improve its thrust performance by means of controlling the impingement of the cowl lip expansion waves. [ABSTRACT FROM AUTHOR]

Abstract This paper presents the conceptual design and performance analysis of a partially reusable space launch vehicle for small payloads. The system employs a multi-stage vehicle powered by rocket engines, with a reusable first stage capable of glided or powered flight, and expendable upper stage(s) to inject 500 kg of payload into low Earth orbits. The space access vehicle is designed to be air-launched from a modified aircraft carrier. The aim of the system design is to develop a commercially viable launch system for near-term operation, thus emphasis is placed on the efficient use of high TRL technologies and on the commercial potential of the technical design. The vehicle design is analysed using a multi-disciplinary design optimisation approach to evaluate the performance, operational capabilities and design trade-offs. Results from two trade-off studies are shown, evaluating the choice wing area and thus aerodynamic characteristics, and the choice of stage masses and engines selection on the mission performance. • Conceptual design for horizontal two-stage spaceplane launch system for small payloads. • Technical design driven by commercial viability. • Trade-off through multi-disciplinary design optimisation of vehicle and trajectory. [ABSTRACT FROM AUTHOR]

In this study, to be able to establish commercially viable space-based solar energy systems in space, the state of the art and their future aspects are presented with regard to space law, economy and other need-to-improve aspects of related technology such as transportation from the Earth to orbit, generation of the energy and its wireless transmission. Moreover, ideas and solutions how to develop a commercial value for the space-based solar energy systems are elaborated. Finally, considering the Turkey’s 2035 vision, a conceptual design of a GEO commercial space based solar power system, named as “Günebakan” that would generate 500 MW electricity, on earth is given.

In this paper, among the technological innovations for sustainable development of future space transportation, the safety aspects of rocket propulsion are discussed to show that the achievement of fail-safety or fail-tolerance is equivalent to the realization of safety resilience in the space transportation system. Further, it is shown that such realization is extremely difficult with the conventional essentially explosive propulsion systems (EEPS) and that essentially non-explosive propulsion systems (ENEPS) are indispensable to obtain the required safety characteristics in space transportation systems. Specifically, the safety characteristics of the EEPS and ENEPS are analyzed under both the current Safety-I and the Safety-II conditions advocated by Hollnagel. As a result, it is shown that the safety characteristics of EEPS are in the category of Safety-I, and it is difficult to shift to Safety-II, while the safety characteristics of ENEPS allow achieving the category of Safety-II. The technical challenges of hybrid rockets, which are representative of ENEPS, are expected to be overcome by devising the Altering-Intensity Swirling-Oxidizer-Flow-Type method or similar ones. Thus, ENEPS will pave the way for future failure tolerance of space transportation. [ABSTRACT FROM AUTHOR]

Copyright of Military Science Review / Hadtudományi Szemle is the property of National University of Public Service and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

In this paper, a direct approach is developed for discovering optimal transfer trajectories of close-range rendezvous of satellites considering disturbances in elliptical orbits. The control vector representing the inputs is parameterized via different interpolation methods, and an Estimation of Distribution Algorithm (EDA) that implements mixtures of probability models is presented. To satisfy the terminal conditions, which are represented as non-linear inequality constraints, several feasibility conserving mechanisms associated with learning and sampling methods of the EDAs are proposed, which guarantee the feasibility of the explored solutions. They include a particular implementation of a clustering algorithm, outlier detection, and several heuristic mapping methods. The combination of the proposed operators guides the optimization process in achieving the optimal solution by surfing the regions of the search domain associated with feasible solutions. Numerical simulations confirm that space transfer trajectories with minimum-fuel consumption for the chaser spacecraft can be obtained with terminal condition satisfaction in rendezvous proximity operation. • Trajectory optimization in close-range space rendezvous mission is considered. • An evolutionary algorithm based on Estimation of Distribution Algorithms is proposed. • Novel feasibility conserving mechanisms are employed in the algorithm. • Results indicate the effectiveness and the efficiency of the proposed approach. [ABSTRACT FROM AUTHOR]

This paper reports on the conceptual design of a three-stage launch vehicle (LV) with a clustered hybrid rocket engine (HRE) through multi-disciplinary design optimization. This LV is a space transportation concept that can deliver micro-satellites to sun-synchronous orbits (SSOs). To design a high-performance LV with HRE, the optimum size of each component, such as an oxidizer tank containing liquid oxidizer, a combustion chamber containing solid fuel, a pressurizing tank, and a nozzle, should be determined. In this study, paraffin (FT-0070) is used as a propellant for the HRE, and three cases are compared: In the first case, HREs are optimized for each stage. In the second case, HREs are optimized together for the first and second stages but separately for the third stage. In the third case, HREs are optimized together for each stage. The optimization results show that the performance of the design case that uses the same HREs in all stages is 40% reduced compared with the design case that uses optimized HREs for each stage.

With the aims of creating and validating innovative fundamental technologies for high-speed atmospheric flights, a small scale supersonic flight experiment vehicle is designed as a flying test bed. An aerodynamic configuration is proposed for the 2nd-generation vehicle with a cranked-arrow main wing and a single Air Turbo Ramjet Gas-generator-cycle (ATRGG) engine. Its longitudinal, lateral, and control surface aerodynamics are characterized through intensive wind-tunnel tests. They are found to be quite moderate except that the directional stability deteriorates severely at large angles of attack and side slip, and that the elevon deflections for roll control cause adverse yaw. These aerodynamic anomalies will result in a tendency of roll reversal at large angles of attack. It can be prevented to some extent by coordinated rudder deflections. In addition, necessity of transonic drag reduction is clarified through thrust margin and flight capability analyses. Probability of 5 to 20 % drag reduction in the transonic regime (Mach 0.8 to 1.2) is demonstrated by configuration modification on the basis of the area rule at Mach 1.1. [ABSTRACT FROM AUTHOR]

The dynamic elastic characteristics of an experimental physical model of an ampoule for the space transportation of radioactive wastes in the self-delivery mode, i.e., when the wastes are the primary energy source, are described. An experimental system is designed to investigate the characteristics of the model of the ampoule, and a procedure for investigating both the cold and heat-stressed state of the ampoule is developed. The amplitude-frequency characteristics of the ampoule model are obtained for different temperature situations. [ABSTRACT FROM AUTHOR]

In this study, to be able to establish commercially viable space-based solar energy systems in space, the state of the art and their future aspects are presented with regard to space law, economy and other need-to-improve aspects of related technology such as transportation from the Earth to orbit, generation of the energy and its wireless transmission. Moreover, ideas and solutions how to develop a commercial value for the space-based solar energy systems are elaborated. Finally, considering the Turkey's 2035 vision, a conceptual design of a GEO commercial space based solar power system, named as "Günebakan" that would generate 500 MW electricity, on earth is given. [ABSTRACT FROM AUTHOR]

The existence of a platform at the stratosphere is expected to be the role to gain a foothold in space. When the platform is employed airship in, it is transported by advection due to its light-weight body. Thereupon, a mooring of platform is reasonable in order to incarnate fixed-point platform. Conception of a captive platform system is similar to the Rockoon one which rendered remarkable services for the space observation of the early days. Therefore, the application of captive platform system to the launch site is highly expected in order to realize the space transportation system with a large quantity and high frequency. In this study, the motion of a captive platform system constructed from an airship and a tether is analyzed in order to ensure that the motion of a captive platform system is practically incarnated. Moreover, the sensitivity of aerodynamic performance of the airship is observed. Consequently, the practically incarnated motion of a captive platform system has been identified under the condition of the storm environment. Furthermore, the reduction of the advection distance driven away by the storm has been quantitatively indicated under the different condition of aerodynamic performance of the airship. [ABSTRACT FROM AUTHOR]

Abstract: This paper introduces new concepts of on-orbit propellant depots for human space exploration based on contingency propellant. The proposed architecture is useful in that it does not require separate depot filling missions, whereas conventional depot architectures require large “prior investment” type missions for depot filling before gaining the returns. Two concepts for this type of depots are shown: “steady-state” architecture and “stockpiling” architecture. In the “steady-state” mode, the depot always keeps the contingency propellant in orbit as well as the reused habitat module. In each mission, the vehicles collect the habitat and the contingency propellant from the depot in orbit on its way to the destination, perform the maintenance for the habitat, and leave the habitat and the unused contingency propellant in orbit on its way back. In the “stockpiling” mode, on the other hand, the habitat module is reused in the same way, but the depot accumulates propellant so that a later mega-mission can carry larger payload. Numerical results show the usefulness of the proposed architectures. [Copyright &y& Elsevier]

Innovation in technologies for high-speed atmospheric flights is essential for establishment of both supersonic/hypersonic and reusable space transportations. It is quite effective to verify such technologies through small-scale flight tests in practical high-speed environments, prior to installation to large-scale vehicles. Thus we are developing a small-scale supersonic flight experiment vehicle as a flying test bed. Several aerodynamic configurations are proposed and analyzed by wind tunnel tests. A twin-engine configuration with a cranked-arrow main wing is selected as the baseline of the first generation vehicle. Its flight capability is predicted by point mass analysis on the basis of aerodynamic characterization and propulsion performance estimation. In addition, a prototype vehicle with an almost equivalent configuration and dimension is designed and fabricated for verification of the subsonic flight characteristics of the experiment vehicle. Its first flight test is carried out and good flight capability is demonstrated. Furthermore a revised aerodynamic configuration with an air-turbo ramjet gas-generator cycle (ATR-GG) engine is being designed for the second generation design with improvement in flight capability at higher Mach numbers. Development of the engine, airframe structure, and autonomous guidance/control system is underway. This prospective flight experiment vehicle will be applied to flight verification of innovative fundamental technologies for high-speed atmospheric flights such as turbo-ramjet propulsion with endothermic or biomass fuels, MEMS and morphing techniques for aerodynamic control, aero-servo-elastic technologies, etc. [ABSTRACT FROM AUTHOR]

Considering the cost reduction of the space transportation system, suborbital spaceplane, which flights ballistically to the altitude of 100 km, is considered one of a promising approach. Because of the avoidance of the orbital revolution, which leads to the reduction of the initial loading fuel, going to and returning from the space at low cost becomes possible. Due to the reduced re-entry speed, the ablator material is unnecessary and the reusability of the spaceplane is increased. SpaceShipOne, which is one of the suborbital spaceplane, has already reached to the step of the commercial use. The suborbital spaceplane seems possible to enter the orbital revolution by building a Maglev Guide Rail in the form which encircles the earth in Low Earth Orbit, and the Maglev Vehicles, which are driven along to the Guide Rail, berth with the suborbital spaceplane. It considered possible to maintain the Maglev Vehicle stationary in the space by adapting the configuration of the Orbital Ring. In this paper, the orbit of the Guide Rail, which supports not only the Maglev Vehicle but also the suborbital spaceplane, is calculated by astrodynamics. The results demonstrate that, sustainable movement around the earth is possible for Guide Rail in spite of the increase in the weight of Maglev Vehicle due to the docking with the suborbital spaceplane. [ABSTRACT FROM AUTHOR]