「どうせ無理」をやめよう――ロケット開発する町工場・植松電機:SOLIDWORKS WORLD JAPAN 2014から(1/2 ページ)

Experiments were conducted for measuring the amount of waste heat coming from a single liquid droplet stream of silicone oil as the working fluid, using a radiant flux sensor RFS. By using this information, it is possible to handle the exoergic heat of the crystallization from the supercooling state by changing the AMP mass and minimum temperature during cooling. To meet a demand for telescopes lighter than the foil optics but with a better angular resolution less than 1 arcmin, we are developing micropore X-ray optics based on micromaching technologies. The three most common internal panel configurations in a 50 kg satellite with body-mounted solar cells and dimensions of 550 x 350 x 550 mm were considered. This study investigates the continuous transition from flame-spreading to stabilized combustion near the blow-off limit in opposed forced flow by using expanding solid fuel duct that makes distribution of oxidizer velocity in the axial direction. One is a hysteresis with respect to increase and decrease of the oxidizer mass flow rate. This paper is the first report verifying these virtues using a laboratory scale motor. Numerical analyses were also performed to separate the radiation from the background of the experimental apparatus included in the radiation measured by the RFS. Evaluation of the thermal onset of graphite nozzle erosion. AMP has the attractive characteristic of storing heat energy in its solid supercooling state, similar to solid-liquid PCMs. Static firing tests of a hybrid rocket motor using liquid nitrous oxide N2O as the oxidizer and high-density polyethylene HPDE as the fuel are analyzed using a novel approach to data reduction that allows histories for fuel mass consumption, nozzle throat erosion, characteristic exhaust velocity c efficiency, and nozzle throat wall temperature to be determined experimentally. This study is an investigation of axial-injection end-burning hybrid rockets aimed at revealing fuel regression characteristics under relatively high-pressure conditions. However, there is no hysteresis characteristics in calculation results. The axial-injection end-burning hybrid rocket proposed twenty years ago by the authors recently recaptured the attention of researchers for its virtues such as no xi oxidizer to fuel mass ratio shift during firing and good throttling characteristics. Moreover, the heat energy that is kept in the supercooling state can be also controlled by crystal nucleus addition or impact. The authors of this paper introduce a new reconstruction technique titled nozzle-throat reconstruction technique to estimate nozzle-throat-erosion history and oxidizer-to-fuel-mass-ratio history in hybrid rockets. In this study, the authors conduced ten firings to investigate a hysteresis characteristics in Axial-Injection End-Burning hybrid rockets under throttling operation. The combustion of aluminum and water is of relevance to many propulsion and energy conversion applications. The fast response is explained by a thermal lag in the solid fuel, whereas the slow response requires further inquiry. Experimental results show that except the fore-end face of the uppermost block and the back-end face of the rearmost block, similarity conditions based on convective heat transfer mechanisms are valid on end faces of fuel blocks. Firing tests are conducted using gaseous oxygen as the oxidizer at chamber pressures and oxidizer port velocities ranging from 0. It was found that uncertainties in LOX travel time, Reynolds number grouping and model assumptions for the first upstream burning surface have the largest impact on the simulator accuracy and are identified as the main focus points for further research. A numerical model was developed based on the conservation of mass in the chamber. There is a transient time in the c efficiency histories of around 2. Therefore, the hysteresis characteristics are not caused by unsteady residence time in the chamber and port merging combustion. Published by SPIE under a Creative Commons Attribution 3. The tool output combinations of optical properties that satisfied the predefined allowable temperature range of structures and components. Because the impinging jet onto the fore-end face of the uppermost block is not high-temperature combustion gas but virtually pure oxygen, a similarity about chemical reaction is necessary in addition to those about convective heat transfer to realize a similarity condition. Thanks to the benefits of hybrid rocket propulsion, assuring and safe, economic launcher dedicated to lightweight satellites can be developed with a reasonable amount of quality assurance and quality control actions being taken in all aspects of development such as raw material, production, transportation, storage, and operation. It is considered that the former hysteresis has the influence of a chamber pressure response time. Each grain consisted of an array of 0. The emissivity of the RFS at low temperature was measured, using a black radiation ball with known emissivity at 26°C. Correlations of the fuel regression rate result in oxidizer port mass flux exponents of 0. Moreover, the emissivity of the liquid droplet was found to be approximately in the range 0. Chamber pressure and oxidizer mass flow rate were measured during each firing. ORBIS is a small X-ray astronomy mission to monitor supermassive blackholes, while GEO-X is a small exploration mission of the Earth's magnetosphere. We plan to demonstrate this type of telescope in these two missions around 2020. Empirical correlations are shown to predict nozzle erosion rates with a coefficient of determination upwards of 0. It was determined that at approximately -180°C, the emissivity of the RFS falls from the catalog value of 0. Experimental results show that a threshold oxidizer velocity of the transition can be determined. These positive attributes make AMP a good candidate to assist in heating a system. In this study, these characteristics were obtained by a combination of experiments and numerical analyses. Therefore, modification of the model is needed for calculating the fuel regression rate of an end-burning hybrid rocket. The oxidizer velocity used in this experiment ranges from 0. The trend in results as calculated using the granular diffusion flame model agrees with that in experimentally observe values. There are several requirements for realizing this type of hybrid rocket: 1 high fuel filling rate for obtaining an optimal xi ; 2 small port intervals for increasing port merging rate; 3 ports arrayed across the entire fuel section. Polymethyl methacrylate PMMA rectangular ducts were used as a fuel, and gaseous oxygen was used as an oxidizer. By applying a multi-objective optimization technique for such a launch system, examples of possible launch systems are obtained for a typical mission scenario for the launch of lightweight satellites. The results were compared using several experimental and numerical data to check the plausibility of the model. We plan to demonstrate this optics in these two missions around 2020, aiming at future other astronomy and exploration missions. Two Japanese missions, ORBIS and GEO-X, will carry this telescope. The results show that oxidizer to fuel ratio remains constant for similar values of oxidizer mass flow rate. It was found that if greater mixing occurs, suction performance is degraded, explaining the actual phenomena of the experiments. AIAA Propulsion and Energy 2020 Forum. This study is an investigation of the response mechanisms in Axial-Injection End-Burning Hybrid Rockets. Two new Japanese missions ORBIS and GEO-X will carry this optics. The results clearly distinguish the initial transient and steady periods of the end-burning mode and prove that no oxidizer-to-fuel ratio shift occurs during firing. This increase in tubular equivalent regression rate is shown to correspond to an increase in performance range from a classic tubular hybrid rocket at low CAMUI Numbers 0. A data reduction method was developed to avoid the difficulty in calculating oxidizer-to-fuel ratio. Therefore, in the throttling tests where oxidizer flow rate was turned-up and returned to the initial condition twice back-to-back, the chamber pressure history was higher in the second iteration than in the first. A preliminary thermal design was proposed by determining all the possible combinations of solar absorptivity and infrared emissivity on the panel surfaces of Earth-pointing satellites deployed from Japan's experimental module small-satellite orbital deployer. All firing tests were conducted at atmospheric pressure. Because the initial transient is a period for the exit end face to attain a steady-state shape, an initial end-face shape being close to the steady-state shape can shorten this period. Therefore, the efficient and quick design method to construct Earth escape trajectory with high flexibility in the boundary condition such as escape velocity, direction and timing is strongly demanded. The regression characteristics of axial-injection end-burning hybrid rocket were experimentally investigated using a laboratory-scale motor. The results show that xi remains almost constant throughout tests of varying oxidizer mass flow rates, and that regression rate in the axial direction is a nearly linear function of chamber pressure with a pressure exponent of 0. The results show that two types of hysteresis characteristics were observed when throttling operation is repeated. The objective of this study is to obtain a rule to define a similarity condition under which subscale tests should be conducted to simulate firings of full-scale CAMUI-type hybrid motors. In addition, the latter hysteresis is not necessarily observed even in the same chamber pressure region. The time-averaged regression rate along the fuel surface was measured by a laser displacement sensor. Seventeen static firing tests are carried out on a small-scale hybrid rocket motor using liquid nitrous oxide as the oxidizer to investigate the chemical erosion characteristics of graphite nozzles. A typical concept of development of such an economic three-stage launcher, in which clustering unit hybrid rocket engines are employed, is described with a development scenario. Second, the chamber pressure history exhibited hysteresis characteristics of oxidizer mass flow rate due to the increasing fuel regression rate. First, a pressure transient was observed when oxidizer mass flow rate was increased turn-up operation. Furthermore, it is observed that the pressure exponent of the fuel regression rate is 1. This study applied this characteristic to methods handling the exoergic heat energy of the crystallization of AMP. Simulation errors compared to test firings are described and followed by an analysis of the potential uncertainties causing this error. The boundary between flame-spreading and stabilized combustion has not been investigated in detail. A modified regression rate formula for the uppermost stage of CAMUI-type hybrid rocket motor is proposed in this study. First, the thermal properties are studied by DSC measurement and thermal cycle tests in different mass conditions. However, two weak points were identified in these throttling firing tests. An experimental setup which simulates the combustion phenomenon involved in the uppermost stage of a CAMUI-type hybrid rocket motor was constructed and the burning tests with various flow velocities and impinging distances were performed. The axial-injection end-burning type fuel grains were made by high-accuracy three-dimensional printing. There is a strong rationale for the usage of the hybrid propulsion for economic space launch as suggested by the assessment conducted here. The CAMUI Number ranges from 0-1: 0 means no CAMUI-type blocks are used, 1 means only CAMUI-type blocks are used. Conductive insulation was applied between the inner and outer structures to decrease the temperature change of inner components. Results of 15 static firings tests show that fuel regression rate increases as the chamber pressure rises, and fuel regression rate decreases as the oxidizer port velocity increases. Proceedings of the Combustion Institute. The combustion characteristics of N2O, which is very useful oxidizer, have not been researched in particular. Results show that fuel mass consumption was nearly perfectly repeatable, whereas the magnitude and timing of nozzle throat erosion was not. Assuming a quasi-steady, one-dimensional, an energy balance against a control volume near the fuel surface is considered. Lastly, nozzle erosion rates exceed 0. Static firing tests with fuel grains of different scaling have estimated the validity of similarity conditions based on convective heat transfer mechanisms. A firing test with fuel having tapered ports is shown to attain a steady-state shape in less than 1s, which is much shorter than the nontapered case of about 6 s. Numerical analysis of nozzle heating and erosion in hybrid rockets and comparison with experiments. However, this does not hold true in tests with varying oxidizer port velocity. The diffusion flame traveled in the opposed-flow field where the oxidizer velocity increases continuously in the upstream direction. The calculation results show that the slow first-order lag can be explained by the time in which ports merge with one another, and the fast first-order lag can be explained by the unsteady residence time in the chamber. Using these combinations, the thermal design for microsatellites in a low Earth and non-sun-synchronous orbit may be shortened. The suction performance ejector-jets has long been studied experimentally and numerically at JAXA, and little success has been achieved in explaining the deterioration of suction performance with high-temperature gas or light gas such as helium. Accordingly, the regression rate formula which can calculate the local regression rate by the quenching distance between the flame and the regression surface is derived. Furthermore, some important technologies that contribute strongly to economic space launch by hybrid propulsion are described. Visualization of fuel regression rate in axial-injection end-burning hybrid rocket. This is done by firing a motor under the same conditions six times, varying only the burn time. Because of difficulty in manufacturing a fuel grain that satisfied requirements such as high volumetric filling rate above 0. The fuel grains used in this study were produced by a high precision light polymerized 3D printer. Nozzle erosion rates reached upwards of 0. Furthermore, through the block-by-block analysis of tubular equivalent regression rate in a fuel grain with a CAMUI Number of 0. Nozzle-throat-erosion histories calculated by the nozzle-throat reconstruction technique agreed well with measured values for initial nozzle-throat radius, and successfully reconstructed the case, in which no measurable amount of nozzle-throat erosion occurred. The analyses were carried out using a simple tool created in MATLAB c. Another is a hysteresis for the cycle. An analytical model is formed based on previous research, and used to develop an informed empirical formula for experimental correlations. © 2019, American Institute of Aeronautics and Astronautics Inc, AIAA. In this study, the threshold velocity was 26. The pressure transient consisted of two distinguishable first order lags, a fast lag followed by a slow lag, which are treated by separate curve fitting functions. The worst hot-and cold case conditions were estimated based on the beta angle of the orbit and the Earth's distance from the sun. In the present study, based on former models, a simple one-dimensional model was introduced incorporating the mixing effects of the primary flow rocket flow and secondary flow induced air flow. To reduce the cost of space transportation, air-breathing engines are considered to be candidates for propulsion. This means PCMs have the ability to sustain heat energy for long periods and select the heat supply timing. In this paper, the families of Moon-to-Moon transfers with sequential lunar swing-by on a hyperbolic orbit are computed and stored in a database. The application of these results can improve the feasibility of LDRs. This paper reports the results of verification firing tests of a novel end-burning-type hybrid rocket made possible for the first time by recent progress in three-dimensional printing technology. The quenching distance during the combustion event was also identified from the observation. Tubular equivalent regression rates are calculated for eight previously reported CAMUI-type hybrid rocket firing tests and compared with extrapolations of previously reported empirical correlations for classic, swirl and vortex hybrid rockets. The authors conducted several firing tests with various oxidizer mass flow rates and chamber pressures, and analysed the results, including xi history, using a new reconstruction technique. Fuel grains of all scales consist of four cylindrical polyethylene blocks with two axial ports. Hybrid Rockets have advantages of low cost and high safety but there are few practical uses at the current state of the art. The stabilized combustion is a diffusion flame that appears in the Axial-Injection End-Burning Hybrid Rocket. Recent advances in high accuracy 3D printing now allow for fuel to be produced that meets these three requirements. A simplified fuel regression model based on the granular diffusion flame model is developed to investigate regression characteristics. Some PCMs store energy when in a non-equilibrium state a supercooling state , and supply energy when released from this state. The use of phase change materials PCMs for heat storage and as a heat source has become an important aspect for energy management. The results show that the supercooling state crystallizes with exoergic heat during the heating process. This paper describes the error and uncertainty analysis of the CAMUI hybrid rocket regression simulator. ORBIS is a small x-ray astronomy mission to monitor supermassive blackholes, while GEO-X is a small exploration mission of the Earth's magnetosphere. For each uncertainty identified, a sensitivity analysis is then performed with the help of a custom-built simulator to evaluate its impact on the simulator accuracy. However, to cover a wide range of flight speeds, the propulsion system has to operate in various modes to be efficient under incoming atmospheric-air conditions. The results of 15 static firing tests show that the fuel regression rate increases as the chamber pressure increases, and regression rates range from approximately 1. Although nozzle erosion was not repeatable, the erosion onset factors were similar between tests, and greater than values in previous research in which oxygen was used as the oxidizer. The Japan Aerospace Exploration Agency is proposing a rocket-based combined cycle engine for operation under various condition, an ejector-jet mode being adopted for the low-speed regime. A granular diffusion flame model only takes into account simple solid propellant regression. Second, the crystallization is investigated by observation of crystal growth. Stabilized combustion of circular fuel duct with liquid oxygen. The comparison between the purely experimental and calculated values showed good agreement, although a large systematic error was expected due to the difficulty in accurately identifying the quenching distance. Nine static-firing tests were carried out on a 2kN-class cascaded multistage impinging-jet-type hybrid-rocket motor under varying oxidizer-flow rates to evaluate the accuracy of reconstructed results. The model explains that the backfiring problem tends to occur in relatively high-pressure conditions, and it leads to the conclusion that increasing the nozzle throat diameter is an effective means of preventing backfiring from occurring. Because there is no end face downstream of the rearmost block, the flow field between fuel blocks with intense turbulence does not exist near the back-end face of the block, resulting in a small convective heat transfer rate. Using sidewalls of micropores through a thin silicon wafer, this type can be the lightest x-ray telescope ever achieved. Because these requirements could not be satisfied by common manufacturing methods, no previous researchers have conducted experiments with this kind of hybrid rocket. Using sidewalls of micropores through a thin silicon wafer, this type can be the lightest X-ray telescope ever achieved. The combustion mode changed when oxidizer velocity at the flame tip exceeded a certain value. In this study, the authors conducted twice experiments to verify the throttling characteristics of axial-injection end-burning hybrid rockets. In the modern space development, small-scale deep space mission should be realized to promote frequent and challenging deep space mission. Fuel regression characteristics of axial-injection end-burning hybrid rocket using nitrous oxide. The authors have previously proposed the concept of end-burning-type hybrid rockets, which would use cylindrical fuel grains consisting of an array of many small ports running in the axial direction, through which oxidizer gas would flow. Lastly, the conditions at the onset of erosion are examined, and used to demonstrate, quantitatively, that chemical erosion can be more easily mitigated when using nitrous oxide as the oxidizer than oxygen. Oxidizer mass flow rate and chamber pressure were throttled by actuating valves in a fluid circuit consisting of two oxidizer supply lines. Oxidizer mass flow rate and chamber pressure were throttled by two methods, actuating valves in a fluid circuit consisting of two oxidizer supply lines and a motor controlling. Since it was not known whether the fuel shape is affected by oxidizer port velocity, several single port firing tests were conducted to confirm the effect of fuel regression shape and ensure the precision of the model. As a result, radiative heat transfer is not negligible on this burning surface and causes an error in the similarity condition. These families are useful to enhance the Earth escape energy and to change escape direction which could lead a spacecraft to further destinations. Over 200 data are obtained by employing an innovative data reduction method to determine time-resolved values for nozzle throat diameter, nozzle throat pressure, equivalence ratio, nozzle throat wall temperature and more. In order to realize a liquid droplet radiator LDR , which is an equipment used for waste heat rejection in large space structures, the exhaust heat characteristics of a single liquid droplet stream in vacuum are required. Firing tests were conducted using gaseous oxygen as the oxidizer at a chamber pressure range of 0. Several firing tests were conducted using a 200N thrust class conventional hybrid rocket motor employing high density polyethylene HDPE as the fuel and liquid nitrous oxidizer as the oxidizer.。

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宇宙環境システム工学研究室

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北大・永田教授、学術会議の圧力に言及 防衛省の制度への応募が禁止に/芸能/デイリースポーツ online

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#148 宇宙を小型化したい研究者・永田晴紀さん(工学研究院 教授)[北大人図鑑 No.8]

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北大名誉教授「学術会議幹部が北大総長室に押しかけ船の抵抗を減らす研究を辞退させた」

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どうせ無理だと思わなければ、宇宙開発だってできる。「植松電機」前編|「colocal コロカル」ローカルを学ぶ・暮らす・旅する

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永田 晴紀(ナガタ ハルノリ)(工学研究院 機械・宇宙航空工学部門 宇宙航空システム)

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堀江貴文氏がそれでも宇宙を目指す本当の理由

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北大名誉教授「学術会議幹部が北大総長室に押しかけ船の抵抗を減らす研究を辞退させた」

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