• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.649-656
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• 2004

Hypersonic wind tunnel test of the rectangular variable geometry intake is performed. For realization of a Precooled turbojet engine, development of a hypersonic ramjet engine is planned. To investigate performance of the intake of the hypersonic ramjet engine, wind tunnel test is done with freestream Mach number of 5.1. The total pressure recovery was 18 % with 12.9 % of ramp bleed. Several reasons for low total pressure recovery are shown. Supersonic internal compression is not enough. Then, the throat Mach number is high (M2.61) and total pressure losses at the terminal shock is large. Supersonic flow at the throat and position of the terminal shock is sensitive to a difference of the second ramp's throat height and the third ramp's throat height. Flow separations at the second ramp's trailing edge and the third ramp's leading edge are seen those could result in the trigger of unstart. The seal mechanism between the ramps and the sidewalls is important.

• International Journal of Aeronautical and Space Sciences
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• v.11 no.3
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• pp.193-200
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• 2010

Through experimental investigations utilizing hypersonic shock tunnel-coaxial thermocouples as well as blow down hypersonic wind tunnel-temperature sensitive paints, the heat flux and the temperature over a protuberance were measured and analyzed. The experimental data were subsequently compared to heat flux data that was obtained by using blow down hypersonic wind tunnel and heat flux gauges. According to the comparison, both sets of data illustrated correlation with one another. The measured heat flux was large when the height of the protuberance was large. Experimental results show that heat flux measurements taken at higher locations were greater than those taken at lower locations. For high protuberances, a severe jump in the heat flux was observed, ranging in values within 0.6-0.7 of the height of the protuberances. However, when the protuberance was sufficiently short, a rise in the heat flux was rarely observed as the protuberance was totally submerged under the separation region.

• 한국전산유체공학회:학술대회논문집
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• 2004.03a
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• pp.17-22
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• 2004

A space launch vehicle departs the ground in a low speed, soon reaches a transonic and a supersonic speed, and then flies in a hypersonic speed into the space. Therefore, the design of a launch vehicle should include the prediction of aerodynamic characteristics for all speed regimes, ranging from subsonic to hypersonic speed. Generally, Empirical and analytical methods and wind tunnel tests are used for the prediction of aerodynamic characteristics. This research presents considerable factors for aerodynamic analysis of a launch vehicle using CFD. This investigation was conducted to determine effects of wake over the base section on the aerodynamic characteristics of a launch vehicle and also performed to determine effects of the sting which exist to support wind tunnel test model.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.10
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• pp.967-974
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• 2009

Experiments were conducted to measure the heat flux and temperature over a protuberance, using an impulse hypersonic shock tunnel-coaxial thermocouples and a blowdown hypersonic wind tunnel-temperature sensitive paints(TSP). Experimental data were compared with the heat flux data using a blowdown hypersonic wind tunnel-heat flux gauges and it was confirmed data sets agreed well. The measured heat flux is large when the height of the protuberance is large. Also, the heat flux measurements at the upper positions are larger than at the lower positions. For high protuberances, a severe jump in the heat flux is observed, from about 0.6~0.7 of the height of the protuberances. However, when the protuberance is sufficiently short, a rise in the heat flux is rarely observed as the protuberance is submerged totally under the separation region.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.463-466
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• 2010

A scramjet engine intake model was tested with a storage air heater type hypersonic wind tunnel. In test results, there is no large performance change with the variation of the sidewall configurations. In the isolator performance analysis, pressure distribution of oblique shock train and normal shock train was observed. Unstart limit of the model was also confirmed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.717-722
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• 2010

Supersonic/Hypersonic wind tunnel is a facility which is intended to test and to observe the physical phenomena around a model by creating supersonic flow in the test section. In designing an airplane, the wind tunnel test is demanded to analyzing aerodynamic characteristics of the model without making a prototype. In this research, the model aerodynamic facility(MAF) is modified for the purpose of increasing running time and its functionality. New pneumatic valves for remote control was installed for safety requirement, and new air tanks was installed on MAF as well. A pipe system is also modified to use those new valves and tanks, and the ceiling and side glasses of the test section are switched to ones with the larger surface area. After the MAF modification, a test is performed at Mach 2, 3 and 4. In this test, shadow graph technique, one of the flow visualization methods, is used to visualize supersonic flow field. The pressure in the settling chamber and working section at Mach 2, 3 and 4 was measured in each case. As a result, the possible model size and running time are obtained.

• 한국전산유체공학회:학술대회논문집
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• 1995.10a
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• pp.235-240
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• 1995

The paper describes hypersonic rarefied flow of helium and nitrogen over a flat plate by the direct simulation Monte Carlo (DSMC) method. The effect of incomplete accommodation and plate thickness are analyzed and the computational results are compared with wind tunnel test data. Also computational aspects of the method are outlined.

• 한국연소학회:학술대회논문집
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• 2012.11a
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• pp.297-299
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• 2012

Korea Aerospace Research Institute has been doing researches on the hypersonic propulsion system and hypersonic wind-tunnel since 2000 and started scramjet engine researches from 2005. Total 5 kinds of scramjet engine were designed and tested and two of them were hydrocarbon-fueled scramjet engine. For verifying the own characteristics of each components like the intake and combustor, several component tests were done at the KSPC of JAXA and KARI. In this paper, current scramjet engine research activities of KARI will be described.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.268-271
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• 2010

Hypersonic engine test facility with ejector system was tested. Ejector system was designed by revised EJSIMP code. The performance of the ejector system was predicted by numerical analysis. As a result, ejector system satisfied the facility design requirement. Based on the pressure level, the facility was successfully started at Mach 3.5 and 20km altitude condition.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.3
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• pp.69-78
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• 2010

Korea Aerospace Research Institute started on design and development of a hypersonic air-breathing engine test facility from 2000 and completed the test facility installation in July 2009. This facility, designated as Scramjet engine test facility(SeTF), is a blow-down type high enthalpy wind tunnel which has a pressurized air supply system, air heater system, free-jet test chamber, fuel supply system, facility control/measurement system and exhaust system. In this paper, details of the specifications, and configuration of the SeTF are described. For verifying characteristics of the SeTF, wind tunnel tests are now on progress and some of the data are also described.