• 유체기계공업학회:학술대회논문집
• /
• 2004.12a
• /
• pp.125-131
• /
• 2004

1.6 MW class supersonic partial admission impulse turbine has been designed and tested in Korea Aerospace Research Institute for the liquid rocket engine application. The test has been performed using a high pressure air source facility in KARI. For the turbine power absorption, a hydraulic dynamometer is used. Appropriate similarity relations are used for the determination of test condition. Various settings of turbine pressure ratio and rotational speed are tested to investigate global turbine characteristics. From measured data, Parameters related to the turbine design are derived and validated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.105-110
• /
• 2010

Supersonic ejectors are simple mechanical components, which generally perform mixing and recompression of two fluid streams. Ejectors have found many applications in engineering. In aerospace engineering, they are used for high altitude testing (HAT) of a propulsion system by reducing the pressure of a test chamber. It is composed of three major sections: a vacuum test chamber, a propulsive nozzle, and a supersonic exhaust diffuser (SED). This paper aims at the improvement in HAT facility by focusing attention on the vertical firing rocket test stand with shock generators. Shock generators are mounted inside the SED for improving the pressure recovery. The results clearly showed that the performance of the ejector-diffuser system was improved with the addition of shock generators. The improvement comes in the form of reduction of the starting pressure ratio and the vertical height of test stand. It is also shown that shock generators are useful in reducing the total pressure loss in the SED.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.758-759
• /
• 2010

In this study, it was experimentally confirmed that the phenomenon of resonance effect of wave drag in two wavy walls. The channel had saw-tooth type of relief surfaces when supersonic gas flows into this channel. Experiment was carried out on the differential apparatus by conducting the comparative test of two nozzles (round sonic nozzle and two-dimensional nozzle with wavy walls). The two-dimensional nozzle was joined alternately with flat walls which had saw-tooth type of symmetrical and asymmetric reliefs. Two-dimensional nozzle was designed for the M=3 and profiled parabolic contour.

• 한국전산유체공학회:학술대회논문집
• /
• 2006.05a
• /
• pp.229-232
• /
• 2006

• 한국항공운항학회:학술대회논문집
• /
• 2004.11a
• /
• pp.182-187
• /
• 2004

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.185-188
• /
• 2009

The rocket motor of the space launch vehicle offers thrust for satellite to enter into the orbit. Characters of the solid propellant for rocket motors are affected by the space conditions such as vacuum and space radiation. The solid propellant used for such a purpose should not undergo physical, internal ballistic and energetic changes when exposed to vacuum and space radiation. This study describes the development of the solid propellant composition for the rocket motor of the space launch vehicle. Also, experimental study was conducted on supersonic diffuser in order to verify the performance of the solid propellant composition which was applied to standard motor on the ground in the vacuum condition.

• The KSFM Journal of Fluid Machinery
• /
• v.18 no.5
• /
• pp.49-53
• /
• 2015

Thrust vector control is the method which generates the side force and roll moment by controlling exhausted gas directly in a rocket nozzle. TVC is classified by mechanical and fluid dynamic methods. Mechanical methods can change the flow direction by several objects installed in a rocket nozzle exhaust such as tapered ramp tabs and jet vane. Fluid dynamic methods control the flight direction with the injection of secondary gaseous flows into the rocket nozzle. The tapered ramp tabs of mechanical methods are used in this paper. They installed at the rear in the rocket nozzle could be freely moved along axial and radial direction on the mounting ring to provide the mass flow rate which is injected from the rocket nozzle. TVC of the tapered ramp tabs has the potential to produce both large axial thrust and high lateral force. We have conducted the experimental research and flow analysis of ramp tabs to show the performance and the structural integrity of the TVC. The experiments are carried out with the supersonic cold flow system and the schlieren graph. This paper provides to analyze the location of normal shock wave and distribution of surface pressure on the region enclosed by the tapered ramp tabs.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.42 no.6
• /
• pp.488-494
• /
• 2014

Rocket Plumes can cause serious damage to launch vehicles and adjacent structures. This paper describes the impact of an adjacent structure by a rocket plume. Each parameter related with dynamic behavior of a missile is modeled with probabilistic distributions of variables. Flyout analyses of initial behavior of a vertically launched missile are performed using Monte-Carlo simulation and flow-motion analyses were conducted by using CFD. In this way, when a missile is fired by a ship, the impact of an adjacent structure by a rocket plume was analyzed.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.23 no.2
• /
• pp.68-77
• /
• 2019

C/SiC composites were developed by a liquid silicon infiltration(LSI) method for use as heat-resistant parts of solid and liquid rocket propulsion engines. The heat resistance characteristics according to the composition ratio (carbon / silicon / silicon carbide) were evaluated by specimen test through arc plasma, supersonic torch test. An ablation equation for oxidation reactions was presented. Through the combustion test it was verified that various parts such as nozzle insert, exit cone and combustion chamber heat resistant parts for rocket propulsion can be manufactured and proved high ablation performance and thermal structure performance.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.05a
• /
• pp.414-416
• /
• 2009

In order to build a transient simulation program for a high thrust liquid rocket engine(LRE), a static performance simulation program for components were made. The components were the thrust chamber (combustion chamber and supersonic nozzle), centrifugal pump (impeller and volute casing), impulse turbine, and flow control devices (control valve and orifice). Simplified mathematical models based on classical thermodynamic and inviscid theories were used to remove complexity and enhance the utility of the program. We examined the results of each program qualitatively for validate each component modeling.