This study investigates the performance of a high altitude test facility with respect to the geometry of the test engine nozzles. Three types of nozzles, conical, bell and pintle, were evaluated under identical throat and outlet diameter conditions. Both numerical simulations and cold-flow tests were conducted to assess the supersonic diffuser's performance depending on nozzle shape and the presence of internal inserts. Numerical results indicated that all nozzle types were capable of achieving the target altitude, with performance ranked in the order of bell, pintle, and conical. The differences were attributed to variations in the ratio of axial to radial velocity components at the nozzle exit. The bell nozzle achieved the lowest chamber pressure and consequently the highest simulated altitude because its greater ratio of axial to radial velocity vomponents significantly amplifgied the entrainment effect. The cold-flow test results showed consistent trends with the numerical simulations.

In this study, the effect of secondary flow velocity on the performance of an ejector used in a high altitude test facility is analyzed. As the secondary flow velocity increases, flow non-uniformity within the vacuum chamber worsens, leading to the formation of recirculation zones and a decrease in suction performance. Numerical analysis and cold-flow tests were conducted, and it was found that increasing the secondary flow velocity from 80 m/s to 100 m/s and 120 m/s caused the vacuum chamber pressure to rise by approximately 2% and 7%, respectively. These results suggest that additional margin in the primary flow rate should be considered during the design and operation of a high altitude test facility to mitigate performance degradation.

Recent trends in launch vehicle design are shifting toward minimizing both weight and cost. Among the design variables, there is increasing comparison between helium, a high-cost but operationally predictable pressurant, and autogenous pressurization, a method that uses evaporated cryogenic propellant gas as the pressurant. In autogenous pressurization, a portion of the cryogenic propellant is vaporized and routed back into the propellant tanks to maintain tank pressure without the need for external gases. Specifically, the weight required for the pressurization system to reach a target altitude is broken down into propellant mass, propellant tank mass, and pressurant tank mass for comparison. The comparison results showed that no clear variable could be derived to definitively favor either helium or autogenous pressurization systems in terms of weight. Therefore, the selection of the pressurization system should be made based on advantages from design and operational perspectives.

Central ejectors have typically been used in high-altitude test facility. When using a central ejector, a heat exchanger is required to protect the steam generator from the test engine combustion gases, which increases the size of the facility. Conversely, an annular ejector does not require a heat exchanger, enabling a reduction in the equipment size. In a high-altitude test facility, unlike that of a central ejector, the secondary flow for an annular ejector is composed of supersonic combustion gases. In this study, we investigated the impact of the annular ejector shape and operational variables on the ejector’s entrainment ratio performance when the secondary flow is supersonic. The effects of the area ratio between the primary and secondary flows and the Mach number of the primary flow were examined using theoretical equations. Numerical analysis was employed to determine the impact of the divergence angle of the primary flow nozzle. Based on the study results, the divergence angle of the primary nozzle is recommended to be set between 9° and 19°. The findings of this study are expected to contribute to the design of annular ejector for high-altitude test facility.

During performance tests of heat exchangers that use fuel-rich exhaust gas from gas generators operated with kerosene and liquid oxygen, environmental problems may arise owing to gas emissions from incomplete combustion. Moreover, helium is a common yet expensive pressurant gas, and its scarcity may cause supply disruptions. We accurately simulate the operating conditions of a heat exchanger considering existing environmental and economic difficulties. To address environmental concerns, ethanol is used as the fuel in the gas generator instead of kerosene, whereas nitrogen is used as the pressurant gas instead of helium to reduce costs. Theoretical calculations show that the gas temperature after heat exchange has a maximum difference of 2.9% with respect to the experimental values, demonstrating good prediction of the experimental results through simulations. In addition, we obtain the mass flow ratio at which nitrogen and helium, used as pressurant gases, exhibit the same heat transfer performance. The simulation results suggest that performance testing of a heat exchanger can be carried out considering the suggested low-cost and ecofriendly alternatives.

This study was intended to verify the thermal effects inside the injector according to the change in the shape variables of the triple injector applied to the oxidizer rich preburner of staged combustion cycle through an experimental method. 'Oxidizer tangential hole diameter' and 'oxidizer exit diameter' were selected as the main shape variables of the study, which can secure the thermal stability inside the injector. A total of three injector heads with changed shape variables were manufactured. As a result of the combustion test, the injector in which both the oxidizer tangential hole diameter and the oxidizer exit diameter were increased from the standard injector was the most thermally stable despite having the most cumulative burning time

In this study, the methodology of determination of base bleed motor data for base bleed projectile based on the NATO standard trajectory model, especially STANAG 4355 Method 2 were presented. Ground combustion experiments and aerodynamic performance firing tests were conducted to determine the drag reduction motor data of the base bleed projectile and this data was described based on the NATO standard ballistic model. The derived drag reduction motor data were input into the ballistic equations to complete the ballistic model and it was confirmed that the calculated predicted trajectory from the ballistic model matched well with the measured trajectory from the aerodynamic performance firing tests.

When pressurizing cryogenic propellant with room temperature pressurant, condensation of the pressurant into the propellant may occur due to heat transfer and physical disturbances within the propellant tank. This phenomenon leads to an increase in pressurant consumption and dilution of the propellant, potentially degrading the performance of the combustion chamber. In this study, we investigated minimizing pressurant condensation solely through the pressurization speed in the cryogenic propellant tank. The condensation of pressurant was determined by its pressurant consumption rate. Experimental results showed that as the pressurization speed increased, pressurant consumption also increased. Specifically, at a pressurization speed of 1.8bar/s, pressurant consumption increased by approximately six times compared to lower pressurization speed. Additionally, after pressurizing to a run tank pressure of 50bar at about 0.5bar/s, increasing the pressurization speed from 50bar to 100bar did not increase the pressurant consumption.

This study examines the steady-state combustion characteristics of a green hypergolic ignition propellant. The propellant, with an ignition delay time of approximately 7.5 ms as determined by drop tests, utilized a 95

Research was conducted on methods to simulate combustion instability in a liquid rocket engine. Kerosene/liquid oxygen was used as the propellant, and the goal was to simulate the 1T mode at a combustion pressure of 60 bar. In this paper, the design and validation of the non -cylindrical combustion chamber, which has a rectangular cuboid shape and is a key component of the simulation device, were performed. Since it is not a typical cylindrical combustion chamber, structural reviews and a natural frequency analysis were conducted during the design process. Validation of the combustion chamber was carried out through combustion tests at both 10 bar and 60 bar. The test results showed stable ignition and combustion pressure, and there was no structural damage to the combustion chamber. Therefore, it was concluded that the validation of the non -cylindrical combustion chamber was successfully completed.

Carbon fiber-reinforced silicon carbide composites (C/SiC) have high specific strength and excellent oxidation resistance at ultra-high temperatures and are currently being studied for application to hypersonic vehicles. This paper proposes practical modeling techniques and an analysis flow for structural design using 2.5D C/SiC materials. Based on the change in the specimen stiffness, the fiber and matrix characteristics were respectively defined as linear elastic and nonlinear plastic behaviors. Subsequently, by applying the modified rule of mixtures, the materials of the laminated representative volume element (RVE) unit were defined as possessing homogenized physical properties, and effective constitutive equations were established. To reduce calculations, the plasticity behavior of the matrix was approximated using a bilinear function, and the least-squares method was applied to simulate the plasticity region. Subsequently, finite-element analysis (FEA) and test results were compared and analyzed for the bending problem of the most significant interest in general structures. Defining the 2.5D C/SiC as a homogenized physical property confirmed that the bending behavior and stress field were suitable from a physical perspective.

In this study, an acoustic excitation device was developed to simulate the combustion instability of a liquid rocket combustor. A non -cylindrical combustor for 1T mode simulation was composed of an auxiliary nozzle and excitation device, and the excitation device consisted of a servo motor and excitation wheel. The designed excitation device was verified through cold test and combustion test. As a result of the cold test, the acoustic resonance frequency was observed to be close to the theoretical value. Additionally, the phase difference corresponding to each mode could be confirmed according to the position of the dynamic pressure sensor. It was also confirmed that more than 5% of the combustion chamber pressure can be disturbed in the 1T mode. During the combustion test, changes in the sonic velocity of combustion gas led to the simulation of the 1T1L mode instead of the 1T mode. Despite this, the excitation device's disturbance performance was successfully verified. The verification test of the excitation device used for combustion instability simulation was successfully performed.

Among the types of pressure swirl atomizer, the screw-type pressure swirl atomizer has advantage in the atomization performance and high reliability with a simple structure. In this study, the screw type pressure swirl atomizer was designed for fuel and oxidizer so that it could be applied to a bi-propellant liquid rocket. The spray characteristics of the atomizer were identified through the cold flow test. Then, fast and stable combustion was confirmed through the combustion test. It represents that impinging injector head using the screw type pressure swirl atomizer can be applied to a bi-propellant propulsion system.

When measuring the thrust of a liquid propulsion rocket engine, an error occurs due to resistance of the propellant and purge piping, so measures are required to compensate it. In this study, it was developed that a vertical thrust measurement system for liquid rocket engines that increases structural stability using three-point support structures and improves the convenience of measuring system resistance using a self-correction system composed of a pneumatic cylinder and two load cells. In addition, in this study, a performance verification device for TMS was also developed, too. It is a device that simultaneously provides the same simulated thrust to the TMS and a separate load cell so that the two values can be compared. Through this, it was confirmed that the TMS in this study has measurement errors within 0.3%.

The pintle injector is considered to have high reliability because it can control engine thrust by adjusting pro-pellant flow rate via adjusting the orifice area and is strong against combustion instability. Although many studies have been conducted on pintle injectors, their accessibility is limited and studies on performance opti-mizations for propellant type and propellant mixing efficiency according to pintle shape and atomization are incomplete. In this study, a 1.5-tonf class liquid-liquid pintle injector with rectangular two-row orifices that uses kerosene/liquid oxygen as the propellant was designed and manufactured. Combustion tests were performed on the pintle injector to verify performance and stability under supercritical conditions, which are the actual operational conditions of liquid rocket engines. From the combustion tests on the initial prototype, the pintle tip was observed to be damaged by heat, so the pintle injector design was changed, and thermal fluid analysis was performed to analyze the pintle tip cooling and combustion performances. To increase the cooling performance of the pintle tip, we devised a method of changing the shape of the pintle orifice and inserted a cooling device called an insert nozzle into the pintle without material changes or applied coatings, as in previous studies. The thermal fluid analysis results showed that there was a difference of up to 147-400 K in the pintle tip cooling performance depending on the insert nozzle and blockage factor, which was verified through combustion tests. The characteristic velocity efficiency and heat flux showed increasing tendencies with increase in total mo-mentum ratio, and a difference of up to 2.0% was confirmed for the characteristic velocity efficiency.

When constructing a high-altitude test facility using a diffuser and an ejector, a verification test for each part is required. This study proposes a natural intake method to independently verify the ejector. The proposed natural intake method enables a verification process without a secondary flow supply facility. To confirm the validity of the verification via the natural intake method, a sub-scale cold-flow test, and a full-scale combustion test are performed. First, the sub-down cold-flow test shows that the performance of the proposed method is similar to that of the conventional approach using the secondary flow supply facility in terms of vacuum pressure. In addition, through the full-scale combustion test, the ejector performance of the proposed method is about 26% lower than that of the conventional method (i.e., high-altitude test) due to the difference between the L/D of secondary flow supply part.

When developing a combustor, a combustion stability rating test must be performed. In this study, a cryogenic nitrogen supply system was designed as a preliminary study for the development of a non-cylindrical scale-down model combustor, which can determine whether combustion is unstable at the initial stage of combustor development. A cryogenic nitrogen supply system is an essential component for cooling and acoustic frequency regulation in the combustor. To verify the manufactured system, an ignition test at atmospheric pressure and a cryogenic nitrogen supply test were performed. And an operating procedure was established that could reliably ignite when a large amount of cryogenic nitrogen was supplied. After the test, there was no structural damage to the cryogenic nitrogen manifold, and the design mass flow rate was also stably supplied, so it is judged that the verification of the manufactured supply system has been completed.