• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.39 no.10
• /
• pp.966-978
• /
• 2011

Transpiration cooling is the most effective cooling technique for the high-performance liquid rockets and air-breathing engines operating in aggressive environments with higher pressures and temperatures. When applying transpiration cooling, combustor liners and turbine blades/vanes are cooled by the coolant(air or fuel) passing through their porous walls and also the exit coolant acting as an insulating film. Practical implementation of the cooling technique has been hampered by the limitations of available porous materials. But advances in metal-joining techniques have led to the development of multi-laminate porous structures such as Lamilloy$^{(R)}$ fabricated from several diffusion-bonded, etched metal thin sheets. And also with the availability of lightweight, ceramic matrix composites(CMC), transpiration cooling now seems to be a promising technique for high-performance engine cooling. This paper reviews recent research activities of transpiration cooling and its applications to gas turbines, liquid rockets, and the engines for hypersonic vehicles.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.05a
• /
• pp.130-134
• /
• 2006

The injection uniformity of the fuel manifold in a liquid rocket engine has been analyzed with dividing plates to improve the cooling performance at the face plate. Three dimensional computational fluid dynamics analysis has been performed to compare the injection uniformity for 5 candidate designs and has been verified to compare with the measured data for the optimal manifold design. For the case I and II, the coolant mass flux increases as the whole working fluid is enforced to flow under the dividing plate. The injection uniformity decreases due to the variation of mass flux at the end of dividing plate and the concentration of mass flow rate at the center of manifold. However case III and IV have uniform injection performance due to reduced mass flux concentration as the coolant can flow along both upper passage and lower passage of the dividing plate. Among the candidate designs, case IV is thought to be the optimal dividing plate with regard to cooling performance and injection uniformity.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2005.11a
• /
• pp.353-357
• /
• 2005

LOx post damage occurs from the development process of the full-scale gas generator which is necessary to 30 tonf class engine development was described. The cause and analysis for damage was described. The combustion test result of 4 injector, the full-scale gas generator and redesigned injector was described. Combustion instability, purge, the low momentum of LOx spray, small recess number, the low flow of LOx, and the high spray angle is main reason the possibility of knowing. The redesign for the injector in the direction of increase of recess number, increase of LOx and fuel spray angle, decrease of gap interval between the LOx post outer wall and fuel screen and increase of LOx post wall thick became accomplished.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.10 no.4
• /
• pp.40-46
• /
• 2006

A 30-tonf-class fullscale thrust chamber for the application to a Low-Earth-Orbit Space Launch Vehicle has been combustion tested over the wide ranges of a mixture ratio and a chamber pressure. The thrust chamber designed for a pump-fed open cycle engine was tested with an ablative chamber instead of a regenerative one for the initial evaluation of its performance and function. The test results revealed stable combustion characteristics. The hardware survived the harsh environment and showed very sound functional characteristics. The measured combustion efficiency turned out to be 95% and a specific impulse at sea level was estimated as 254sec, which are comparable to or above the predetermined design values.

• Aerospace Engineering and Technology
• /
• v.1 no.2
• /
• pp.105-112
• /
• 2002

In this paper, a servo filter design for the gimbal engine actuator system of KSR-III(Korea Sounding Rocket-III) is considered. A reasonable filter structure is determined based on the actuator analytic models. The servo filter consists of a 2-nd order lowpass filter and a 1-st order compensator. The lowpass filter is required to protect the actuator from high frequency vibration, and the compensator to enhance the resulting stability. A Butterworth type servo filter is considered as the simplest one. The final servo filter type is determined by evaluating simultaneously both high frequency gain reduction performance and the corresponding KSR-III stability margin. Consequently it is revealed that a notch type servo filter located on the error between command signal and feedback signal in the control loop is very effective. Later, based on the proposed servo filter type, an onboard servo filter hardware of KSR-III will be designed and tested.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.622-627
• /
• 2017

In this study, the numerical analysis of propellant supply manifold of a sub-scale model combustor with the same combustion and acoustic resonance conditions as the real combustor was carried out. The analysis of the results is based on the shape of the manifold and the number and type of inflow channels. The manifold form of the main propellant was rounded so that the recirculation area was small and easy to manufacture. The liquid oxygen mainfold included a distributor to uniformly supply the flow, and since the kerosene manifold was directly in contact with the flame side, the inflow channel was formed at a radial $360^{\circ}$ so as to minimize the recirculation region. The liquid nitrogen manifold was configured as a radial inflow channel to reduce the velocity difference near the injector.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.24 no.1
• /
• pp.64-77
• /
• 2020

An electric-pump cycle, which is a propellant supply system for driving pumps of a liquid rocket engine using an electric motor, has the advantages of simple system configuration and easy control of supply flow rate and pressure. This paper investigates and analyzes the overseas research trends of the electric-pump cycle. In addition, the research and development country, performing organization, application, engine thrust, pump pressure increase, motor power, and rotation speed are summarized. Among them, the design variables of the overseas research that applied the upper-stage propulsion system with the thrust range of 0.445~2.2 kN could be used in the study of a similar electric-pump cycle in Korea.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.25 no.3
• /
• pp.62-80
• /
• 2021

The requirements for the next-generation propulsion system and for a good propellant have been summarized. The characteristics and effectiveness of kerosene, hydrogen, and methane, which are the fuels that are mainly attracting attention in Korea and abroad, were compared with each other. As a result of the comparison, methane was evaluated to be more advantageous than other fuels in reliability, cost, reusability, maintenance, eco-friendliness, safety, lifespan, technical difficulties, engine cycle selection, application of common bulkhead, and non-disassembly/reassembly delivery. And in terms of performance, the specific impulse of methane is higher than that of kerosene, so the efficiency of the launch vehicle can be increased. Methane's properties incluidng eco-friendliness, low-temperature combustion, long life, and maintenability make it beneficial for reuse and for the development of multi-purpose engines.

• Aerospace Engineering and Technology
• /
• v.4 no.1
• /
• pp.143-149
• /
• 2005

To evaluate the hydraulic and cavitation performance of a fuel pump for a liquid rocket engine, performance tests of the pump, which consist of hydraulic tests and cavitation tests, were conducted in water environment with various condition. In the hydraulic tests, the head, efficiency and volute pressure distribution of the pump are found to follow the conventional similarity rule, whereas the secondary flow pressure shows a small deviation from the similarity rule. As the floating gap is constricted, the efficiency of the pump improves and the secondary flow pressure decreases. However, the inner diameter of bypass line orifice does not show clear relationship with the pump efficiency. In the cavitation tests, measurements of the head and the NPSH indicate that the pump shows better cavitation performance as the rotational speed of the pump increases.

• Aerospace Engineering and Technology
• /
• v.4 no.1
• /
• pp.171-178
• /
• 2005

An acoustic analysis of a fuel-rich gas generator for the drive of a turbopump in a liquid rocket engine has been performed and the length of a simulating duct has been determined by comparing the resonant frequency of unstable acoustic modes to simulate an actual flight model gas generator. To simulate more realistically, a realistic short-length simulating duct has been determined by considering 1 or 2 wavelength of the unstable modes. Duct-length adjustment to turbopump can be a method to suppress a combustion instability problem by decoupling of acoustic mode and combustion characteristics. This method has been set up and validated with acoustic analysis and hot firing tests.