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The most influential factor for self-ignition of high-pressure hydrogen is known to be the strength of the shock. Thus, the self-ignition can be suppressed by weakening the shock strength, which is possible by reducing the area where the hydrogen is ejected in this study. To confirm the possibility of this method, experiments were done by controlling the burst pressure of up to 302 bar and the ratio of the opening area. The experimental results showed that the minimum burst pressure of self-ignition is increased exponentially as the opening area is reduced. This confirmed that reducing the opening area under the same burst pressure conditions has an effect on the suppression of self-ignition. However, it was also found that the minimum shock speed that causes self-ignition gradually decreases as the opening area becomes smaller, which results from an increasing in mixing. The CFD simulation results showed that the volume of the flammable region in the tube was increased and the hydrogen-air mixing efficiency also increased when the opening area became smaller. The results suggest that reduction of the opening area can suppress a self-ignition by weakening the shock strength, but it should be noted that an increase in mixing effect also occurs.

The propellant supply system of a liquid rocket engine using an electric pump has high reliability because of the relatively small number of components. The system also has the merit of a quick response and ease of control owing to its simple configuration. Recently, the rocket lab developed the Rutherford engine, which has an electric pump cycle, because of the improved technology in the electric motor and battery. This paper examined the development of the electric-pump cycle and compared the performance with other cycles for a small-sized low-thrust rocket engine. Performance analysis and mass estimation were conducted using the developed analysis program, in which reliability in mass estimation was improved based on the designed configuration or real performance data from commercial products. In addition, the modeling method and analysis procedure were described in detail. The results showed that it is possible to develop a small-sized engine with an electric-pump cycle when the present technologies are applied. The electric-pump cycle had a smaller dry mass than the gas-generator cycle, even at a low thrust level of 500 N, and showed higher performance in specific impulse and speed increments.

The correlation between solid propellant grain configuration and burning surface area profile is a complicated nonlinear problem. Nonlinear optimization has been adopted to design grain configurations that satisfied the objective area profiles. However, as conventional design methods are impractical, with limited performance, it is necessary to investigate alternatives. Useful information for grain design can be obtained by analyzing the aforementioned correlation. However, this aspect has not been studied owing to the requirement of large amounts of data and analysis techniques. In this study, machine learning was used to develop a new design method. The objective of machine learning was to train a model to classify classes of data. The database stores various sets of configuration variables and their classes. The proposed Gaussian kernel-based support vector machine model predicts the class of newly designed grains. The results verified that the model accurately predicted the class of the set of configuration variables and can be used to modify the set of configuration variables to satisfy the requirement. Thus, it was confirmed that machine learning is an appropriate approach to grain design, however, further research is needed to analyze its practicality.

A numerical investigation has been conducted for the drag reduction of a counterflow jet with a blunt body at Mach 6. The computational study was carried out by solving two-dimensional axisymmetric Reynolds-averaged Navier–Stokes (RANS) equations. The Spalart–Allmaras one-equation turbulence model was used in this study. The jet flow of exit Mach number 2.86 interacts with the free stream and changes the single shock structure to a multiple shock structure. The purpose of this study was to investigate the influence of the stagnation pressure based on a parameter, the momentum parameter ratio (MPR), which can characterize the jet. The result shows that the flow field can be categorized as being in a long penetration mode (LPM) or short penetration mode (SPM) depending on the penetration length of the free-stream flow. In the LPM, under overexpanded jet conditions, the shock structure fluctuates continuously, so the flow field is unstable. On the other hand, in the SPM, the shock structure is nearly fixed, and the entire flow field is stable. For this reason, even if the penetration length of the flow field in the SPM is less than that in the LPM, the aerodynamic drag can be reduced by up to 40%. Therefore, it becomes clear that jet pressure conditions can significantly improve aerodynamic performance by reducing the drag applied to the blunt body.

고속 고온 유동에서 나타나는 고온 기체 현상을 모사하기 위해서는 마하수뿐 아니라 절대속도도 재현할 수 있어야 한다. 이러한 유동을 초음속 유동과 구분하여 극고속 유동이라 부르며, 충격파 터널과 같은 고엔탈피 시험 장치를 통해 연구가 이루어지고 있다. 그러나 이러한 고엔탈피 시험 장비는 높은 온도와 압력 때문에 노즐에서 열화학적 비평형 현상을 경험하게 되며 기존의 이론적 방법으로 그 실험 조건을 규정하기 어렵다. 이에 본 연구에서는 알려진 비평형 노즐 코드의 단점들을 보완하고 충격파 터널의 운용 조건에서 시험부 유동 특성을 빠르게 예측하기 위하여 열화학적 비평형을 고려한 준 1차원 노즐 해석 코드를 개발하였다. 개발된 코드는 시험 결과 및 2차원 축대칭 해석 결과와 비교를 통하여 충격파 풍동 시험부 유동 조건 예측을 위한 활용성 및 한계를 살펴보았다.

Acquiring density field of a flow is important to understand the flow structure, especially for studying aero-optics. Background-oriented schlieren (BOS) is one of a flow visualization technique which can obtain density field of the flow. Using BOS images density gradient can be acquired quantitatively and density field can be obtained when additional numerical method is applied. In this paper, density field of a hot jet is reconstructed from BOS images by assuming the jet as asymmetric flow and applying adaptive Fourier–Hankel Abel method. Reconstructed density field is evaluated by comparing with wavefront of the optical wave propagating through the hot jet. Optical path difference (OPD) is calculated from the reconstructed density field and compared with Shack–Hartmann (S–H) sensor measured OPD map. OPD obtained from the reconstructed density field and S–H sensor show similar tendency, especially when results are time averaged and turbulence characteristics are removed. Comparing RMS OPD of each time-averaged result show 8.8% error, which is within reasonable range. Thus, it seems that the density field acquisition using BOS technique is reliable for the hot jet flow.

꼬리날개 조종 유도무기의 주날개-꼬리날개 간섭 현상에 대한 연구를 수행하였다. 풍동시험 데이터를 이용하여 주날개-꼬리날개 간섭 정도를 산출하였으며 날개간의 간섭 현상이 전체 공력에 미치는 영향을 분석하였다. 성분 시험 결과를 이용하여 downwash angle을 산출하였으며 날개간의 간섭 영향을 받음각에 대한 비율로 나타내었다. 날개간의 간섭 현상 발생 시 유동 특성을 살펴보기 위해 수치해석을 실시하였으며 받음각에 대한 vorticity 특성을 비교하였다. 실험적, 수치적 연구를 통해 주날개-꼬리날개 간섭현상이 유도무기의 정안정성에 큰 영향을 미침을 확인하였다.

In this study, numerical simulations were conducted to confirm the possibility of improved mixing performance by using a fluidic oscillator as a fuel injector. Three-dimensional URANS non-reacting simulations were conducted to examine air&ndash, fuel mixing in a supersonic flow field of Mach 3.38. The numerical methods were validated through simulations of the oscillating flow generated from the fluidic oscillator. The results show that the mass flow rate and momentum are reduced at the outlet because the total pressure loss increases inside the fluidic oscillator, which means that higher pressure needs to be applied to supply the same mass flow rate. The simulation showed that the flow structure varies over time as the injected flow is swept laterally. With lateral injection, the fuel distribution is long and narrow, and asymmetric vortexes are generated. However, with central injection, the fuel distribution is relatively similar to the case of using a simple injector. Compared to the simple injector, the penetration length, flammable area, and mixing efficiency were improved. However, the total pressure loss in the flow field increases as well. The results showed that the supersonic fluidic oscillator could be fully utilized as a means to enhance the mixing effect, however a method to reduce the total pressure loss is necessary for practical application.

Numerical simulations were conducted to investigate the transitional characteristics of shock reflections in the dual solution region using laser energy deposition. The numerical approach was validated by comparison to experimental results of the deposition laser energy in front of a blunt model. The simulation results show that the energy deposition in the freestream region can induce a transition of the shock reflection system and the transition characteristics can vary with the position and energy of the laser deposition. As the amount of energy increases, the time required for the transition also increases, and the transition cannot occur when the energy that is applied exceeds a certain level. The time required for the transition can be reduced when the position of energy deposition is moved downstream. The results also show that the transition does not occur regardless of the deposited energy when the laser energy is deposited on the symmetry line.

In this study, a minimum amount of required propellant was calculated by analyzing the sequence with various re-entry conditions. This study aims to obtain data related to variation in trajectory and required propellant weight according to various re-entry scenarios. The drag coefficient at various altitudes, velocities, and thrust was calculated through numerical simulations to raise the reliability of the results. The calculation results were compared to the optimal values extracted from the genetic algorithm. It was observed that the duration of the entry-burn phase is dominant to the total required propellant weight. As a general tendency, high entry-burn starting altitude, high ending Mach number, and low landing-burn starting thrust make the required propellant weight low. However, if the entry-burn ending condition is set to the Mach number, it is necessary to select an appropriate re-entry condition. Additionally, from comparisons with the optimized results, it was confirmed that accurate calculation of the drag coefficient is important to succeed a soft landing of RLV.

This study was performed to analyze the performance of a solid propulsion system quickly and accurately. Systems that are currently in operation use a propellant with a complicated configuration to satisfy the performance requirements. Multi-dimensional internal ballistics analysis is essential in this case. The propellant's burning surfaces move in all directions, and the local burning velocity changes, so a 3-D analysis technique is necessary for the moving boundary. Previous research focused on improving the accuracy of the combustion configuration analysis using new methods, but it is difficult to apply them because user convenience was not considered. Therefore, the problems of the analysis methods were analyzed, and a hybrid burn-back analysis method was developed. The method's accuracy was improved by selectively using the minimum distance function (MDF) manipulation method, which has fast calculation speed, as well as a level set method for regions where the configuration of the propellant grains is complicated. The results verified that the accuracy was increased, and the calculation time was reduced.