Spacecraft surface charging can lead to arcing and a loss of electricity generation capability in solar panels or even loss of a satellite. The charging problem may be further aggravated by atomic oxygen (AO) exposure in Low Earth orbits, which modifies the surface of materials like polyimide, Teflon, anti-reflective coatings, cover glass etc, used on satellite surfaces, affecting materials properties, such as resistivity, secondary electron emissivity and photo emission, which govern the charging behavior. These properties are crucial input parameters for spacecraft charging analysis. To study the AO exposure effect on charging governing properties, an atomic oxygen exposure facility based on laser detonation of oxygen was built. The facility produces AO with a peak velocity value around 10-12km/s and a higher flux than that existing in orbit. After exposing the polyimide test material to the equivalent of 10 years of AO fluence at an altitude of 700-800 km, surface charging properties like surface resistivity and volume resistivity were measured. The measurement was performed in a vacuum using the charge storage decay method at room temperature, which is considered the most appropriate for measuring resistivity for space applications. The results show that the surface resistivity increases and the volume resistivity remains almost the same for the AO exposure fluence of 5.4×10¹⁸ atoms cm^-2.

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We examined the compatibility of radicals contained in plasma jets to fuels through ignition and combustion tests for dimethyl ether (DME)/air and methane (CH₄)/air mixtures with oxygen (O₂) and nitrogen (N₂) as the plasma torch feedstocks. The experiment showed that the DME/air mixture was ignited/combusted with less plasma jet (P.J.) power than the CH₄/air mixture and that the O₂ P.J. is more effective than the N₂ P.J., with a more distinct difference in effectiveness for the CH₄/air mixture in contrast to the DME/air mixture. Plasma jets with fewer feedstock flow rates were more effective, presumably due to the greater amount of radical production under the conditions tested. Numerical estimation on the amount of radicals and ignition delay time demonstrates that the superiority of the O₂ P.J. is not necessarily only due to the effectiveness of the O radicals, but also due to the fact they were produced easier and with less power, and that the effect and behavior according to amount is different for fuels. This is most likely because they depend on the reaction mechanism of each mixture, all of which match well with the experimental results.

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Visual capability is very important for space-based activities, for which small, low-cost space cameras are desired. Although cameras for terrestrial applications are continually being improved, little progress has been made on cameras used in space, which must be extremely robust to withstand harsh environments. This study focuses on commercial off-the-shelf (COTS) CMOS digital cameras because they are very small and are based on an established mass-market technology. Radiation and ultrahigh-vacuum tests were conducted on a small COTS camera that weighs less than 100 mg (including optics). This paper presents the results of the qualification tests for COTS cameras and for a small, low-cost COTS-based space camera.

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A new type of low-density wind tunnel called the “Mars Wind Tunnel (MWT)” has been developed to simulate airfoil flow in the Mars atmosphere. This paper presents the design features of the MWT and its construction. The results of operational performance tests including the driving performance of a multiple supersonic ejector and calibration tests for evaluating the flow characteristics in the test section are shown. These tests were conducted using air as the working gas. It was found that the ejector is an effective means to induce high subsonic flow even under the condition of a low Reynolds number. The pressure ratio achievable with the ejector is inversely proportional to the total pressure. The MWT can simulate flow in the Reynolds number range from 103 to 105 and Mach number up to 0.74. It is confirmed that the MWT has a flow quality and operational characteristics that are suitable for aerodynamic testing of Mars airplanes.

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Korea Space Launch Vehicle-I (KSLV-I) carrying the STSAT-2A satellite, made its maiden flight on August 25, 2009. Liftoff was from the Korea Aerospace Research Institute (KARI) launch site at the Naro Space Center in a southern coastal province of the Korean peninsula. A video telemetry system provided visual monitoring of flight critical events, as well as dynamic moving images, through two cameras equipped on the upper stage. This paper describes the development of the onboard video system consisting of two video cameras, a video compression unit and an RF transmitter meeting the requirements of the KSLV-I launch vehicle. During the flight, the ETTARS, an especially reconstructed small telemetry ground station for the KSLV-I, received the video telemetry data and simultaneously measured automatic gain control (AGC) signal levels from the receiver. The onboard video system achieved a 15fps video rate, flew with a 2Mbps data rate and transmitted data using NRZ-L PCM/FM in the S-band. Errors on real telemetry channels occurred as predicted by link budget equations based on thermal noise and multipath interference. According to error analysis, the BER versus SNR performance degradation from the laboratory reference was about 1dB, which is close to expected limit.

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A system for evaluating the near-field pressure distribution around a supersonic flight model by combining experimental and numerical diagnostics has been developed. Experimental measurement is conducted using a ballistic range with four kinds of axi-symmetric flight models. Schlieren flow visualization is recorded using a high-speed framing camera and near-field pressure histories are measured using piezoelectric pressure transducers flush-mounted on the surface of flat plates in the test section. The numerical diagnostics is done using FaSTAR, a numerical simulation tool developed by the Japan Aerospace Exploration Agency (JAXA). The experimental and numerical data are compared to each other, and the numerical results well validated. Based on the numerical results, it becomes possible to estimate the accuracy of experimental conditions including the flight path and angle of attack, which cannot readily be determined only from experimental data, and to discuss the relationship between peak overpressure and aerodynamic performance. Satisfactory agreement between the experimental and numerical results at a flight Mach number of 1.66±0.02 and important insights related to rear boom strength are obtained.

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Although the pulsed plasma thruster (PPT) was first utilized on a space mission in 1964, after more than four decades, it is still a space-rated technology which has performed various propulsion tasks, from station-keeping to three-axis attitude control for a variety of former missions. With respect to the rapid growth in the small satellite community and the growing interest for smaller satellites in recent years, the PPT is one of the promising electric propulsion devices for small satellites (e.g., CubeSats) due to the following advantages: simplicity, lightweight, robustness, low power consumption, low production cost and small dimensions. Therefore, a laboratory benchmark rectangular breech-fed pulsed plasma thruster using a self-inductor as a coupling element was designed, developed and successfully tested in a bell-type vacuum chamber at 10^-4 Pa for the first time in west Asia (Iran). The PPT has been tested using a 35 μF, 2.5 kV oil-filled capacitor, producing an impulse bit varying from 300 μN-s to 1.3 mN-s at a maximum specific impulse of 1100 s. As a result a research program in Iran was initiated for working on PPTs and the miniaturization of PPTs while increasing the performance parameters. The present paper briefly reviews the PPT design and development.

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Aerodynamic characteristics of a wing with a smart flap under the ground effect are studied through the integration of computational fluid dynamics. A parametric bending profile of a smart flap is designed considering different types of beams. Here, a cantilever beam with uniformly varying load with roller support at the free end is considered. The shape of the smart flap is fixed and its advantage comes from its smooth connection to the main wing. In this research, a pressure-based implicit procedure is used to solve Navier-Stokes equations. A non-orthogonal mesh with collocated finite volume formulation is utilized to simulate flow around the wing under the ground effect. First, the method is validated against experimental data. Then, the algorithm is applied for turbulent aerodynamic flows around a wing with smart and conventional flaps for different flap angles and ground clearance. The results of the two wings are compared. It is found that the pressure coefficient distribution for a wing with smart flaps is smoother than that of a wing with conventional flaps, and tip vortexes of the flap and wing diminish for low ground clearance. Finally, the maximum lift-to-drag ratio (L/D) is obtained for a smart wing when the angle of flap (AOF)=7.5° and h/c=0.3.

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