The science and technology of ultracapacitors are reviewed for a number of electrode materials, including carbon, mixed metal oxides, and conducting polymers. More work has been done using microporous carbons than with the other materials and most of the commercially available devices use carbon electrodes and an organic electrolytes. The energy density of these devices is 3¯5 Wh/kg with a power density of 300¯500 W/kg for high efficiency (90¯95%) charge/discharges. Projections of future developments using carbon indicate that energy densities of 10 Wh/kg or higher are likely with power densities of 1¯2 kW/kg. A key problem in the fabrication of these advanced devices is the bonding of the thin electrodes to a current collector such the contact resistance is less than 0.1 cm2. Special attention is given in the paper to comparing the power density characteristics of ultracapacitors and batteries. The comparisons should be made at the same charge/discharge efficiency.

Ultracapacitors: Why, How, and Where is the Technology

A heat shield called the thermal protection system (TPS) is an important structure in hypersonic vehicles as it prevents hot air from entering vehicles and potential impacts from space debris. With the increase in demand for low-cost reusable launch vehicles as well as for searching and exploration of new planets in both unmanned and manned missions, the need for developing an effective TPS has increased across many countries. The structural design of TPSs has become more prominent in the early stage of hypersonic vehicle development. Sandwich structures that have the advantages of low density and high performance are integrated into the structural design of an effective TPS. This paper provides a comprehensive review of recent research efforts on sandwich structures for TPSs. The topics discussed in this paper include aspects of structural and material design, mechanical and thermomechanical performances, and manufacturing methods. In particular, we review and discuss the structural design as well as the material design of sandwich structures for different TPS types with various configurations, including corrugated cores, lattice cores, multilayer cores, foams, honeycomb cores, bio-inspired cores. The materials used for the sandwich structures, such as various types of laminated composite, ceramic matrix composite, and metals, are included. We also discuss the performance of the TPS sandwich structures in terms of temperature gradients, deformation limits, and mechanical strengths and provide a discussion on the manufacturing methods of TPS sandwich structures for hypersonic vehicles. Finally, further research directions and challenges of sandwich structures for TPSs are presented.

Advanced sandwich structures for thermal protection systems in hypersonic vehicles: A review

Self-discharge is an important performance factor when using supercapacitors. Voltage losses in the range of 5–60% occur over two weeks. Experiments show a dependency of the self-discharge rate on various parameters such as temperature, charge duration and short-term history. In this paper, self-discharge of three commercially available supercapacitors was measured under various conditions. Based on different measurements, the impact of the influence factors is identified. A simple model to explain parts of the voltage decay is presented.

Detailed Analysis of the Self-Discharge of Supercapacitors

Ablative insulators are used in the interior surfaces
of solid rocket motors to prevent the mechanical structure of the
rocket from failing due to intense heating by the high-temperature
solid-propellant combustion products. The complexity of the
ablation process underscores the need for ablative material
response data procured from a realistic solid rocket motor
environment, where all of the potential contributions to material
degradation are present and in their appropriate proportions. For
this purpose, the present study examines ablative material behavior
in a laboratory-scale solid rocket motor. The test apparatus
includes a planar, two-dimensional flow channel in which flat
ablative material samples are installed downstream of an aluminized
solid propellant grain and imaged via real-time X-ray radiography.
In this way, the in-situ transient thermal response of an ablator
to all of the thermal, chemical, and mechanical erosion mechanisms
present in a solid rocket environment can be observed and recorded.
The ablative material is instrumented with multiple
micro-thermocouples, so that in-depth temperature histories are
known. Both total heat flux and thermal radiation flux gauges have
been designed, fabricated, and tested to characterize the thermal
environment to which the ablative material samples are exposed.
These tests not only allow different ablative materials to be
compared in a realistic solid rocket motor environment but also
improve the understanding of the mechanisms that influence the
erosion behavior of a given ablative material.

/pdf/assessment-of-the-performance-of-ablative-insulators-under-24pyqepw7w.pdf

Assessment of the performance of ablative insulators under realistic solid rocket motor operating conditions

New thermal protection system using high-temperature carbon fibre-reinforced plastic sandwich panel

In-orbit feasibility demonstration of supercapacitors for space applications

Thermochemical performance of a Lightweight Charring Carbon Fiber Reinforced Plastic

A computer code for charring ablation and thermal response analysis has been developed for simulation of one-dimensional transient thermal behavior of multi-layer stack of isotropic charring ablation materials used for re-entry capsules. Mathematical model of charring ablation including governing equations is summarized. Analytical solutions of several thermal response problems with or without ablation agreed very well with numerical results that are obtained with the charring ablation analysis code. Arc-heated tests of CFRP ablator models were carried out and measured results including temperature response data were compared with the simulation results of the ablation analysis code. Test results and the simulation results agreed very well. These results show that the analysis code based on the present mathematical model are available for the thermal response prediction of CFRP ablator used in the re-entry capsules such as USERS/REM.

Numerical Analysis of Charring Ablation for Ablative Materials of Re-Entry Capsules.

As small satellites are becoming more widespread for new businesses and applications, the development time, failure rate and cost of the spacecraft must be reduced. One of the systems with the highest cost and the most frequent failure in the satellite is the Electrical Power System (EPS). One approach to achieve rapid development times while reducing the cost and failure rate is using scalable modules. We propose a solar module integrated converter (SMIC) and its verification process as a key component for power generation in EPS. SMIC integrates the solar array, its regulators and the telemetry acquisition unit. This paper details the design and verification process of the SMIC and presents the in-orbit results of 12 SMICs used in Ten-Koh satellite, which was developed in less than 1.5 years. The in-orbit data received since the launch reveal that solar module withstands not only the launching environment of H-IIA rocket but also more than 1500 orbits in LEO. The modular approach allowed the design, implementation and qualification of only one module, followed by manufacturing and integration of 12 subsequent flight units. The approach with the solar module can be followed in other components of the EPS such as battery and power regulators.

https://www.mdpi.com/2226-4310/6/5/61/pdf

Solar Module Integrated Converters as Power Generator in Small Spacecrafts: Design and Verification Approach

There are provided a carbon fiber reinforced prepreg of gas barrier properties comprising a sheet-like carbon fiber reinforcement and a matrix resin including therein a gas barrier layer having a clay mineral with a plate-like crystal structure unidirectionally aligned and densely laminated, and a carbon fiber reinforced plastic produced from the carbon fiber reinforced prepreg. Such prepreg is obtained by disposing a film material of gas barrier properties having a clay mineral with a plate-like crystal structure unidirectionally aligned and densely laminated, in at least one interlayer of a laminate of a carbon fiber reinforced prepreg comprising a sheet form carbon fiber reinforcement and a matrix resin; and then heating and/or pressurizing the laminate. The carbon fiber-reinforced composite material of the invention exhibits high gas barrier properties, and particularly hydrogen gas barrier properties.

Keiichi Okuyama

Papers

In-orbit feasibility demonstration of supercapacitors for space applications

Thermochemical performance of a Lightweight Charring Carbon Fiber Reinforced Plastic

Numerical Analysis of Charring Ablation for Ablative Materials of Re-Entry Capsules.

Solar Module Integrated Converters as Power Generator in Small Spacecrafts: Design and Verification Approach

Carbon fiber reinforced prepreg of gas barrier properties, carbon fiber reinforced plastic and methods of producing the same