• The KSFM Journal of Fluid Machinery
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• v.4 no.1 s.10
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• pp.7-13
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• 2001

In developing a multistage compressor, the stage matching is one of the critical design issues. The mismatching can be often observed even if each stage has been proven good and then used as part of a compression system. A good matching among the stages can be achieved by changing various design parameters (i.e., passage cross sectional areas, blades angles, stagger angles, curvature, solidity, etc.). Therefore, designers need to find out what parameters must be changed and how much. In this study, a method to search the design parameters for optimum stage matching has been used based on an 1-D mathematical model of a compressor, which uses the data obtained from the preliminary test to identify the design parameters. This methodology is applied with a two-stage axial compressor, which was originally designed for a helicopter gas turbine engine. After identifying design parameters using preliminary test data, an optimization process has been employed to achieve the best matching between the stages (i.e., maximum efficiency of the compressor at its operation modes within a given range of the rotor speed under given restrictions for required stall margins and mass flow). 3-D flow calculations have been performed to confirm the usefulness of the corrections based on the 1-D mathematical model. Calculational results agree well with the experimental data in view of the performance characteristics. Some promising results were produced through the methodology proposed in this paper in conjunction with flow calculations.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.5
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• pp.118-126
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• 2002

The Altitude Engine Test Facility(AETF) was built at the Korea Aerospace Research Institute and has been being operated for the gas turbine engines in the class of 3,000 lbf thrust. To enhance the confidence level of AETF to the international level, a series of studies and facility modification have been conducted to improve the measurement uncertainty and reliability. In this paper, some part of the facility evaluation tests performed with a single spool turbojet engine are introduced. Tests were performed simulating the flight conditions as steady state, sea level for various flight speeds (i.e., Mn=0.3, 0.5, 0.7, 0.9). The obtained test results are compared with the predicted values of the engine DECK. The measurement uncertainties of airflow, net thrust, fuel flow and SFC showed 0.791~0.914%, 0.851~1.706%, 1.372~7.348% and 1.642~5.205%, respectively. Thus, from this research, the improvement methods of uncertainties on AETF has been confirmed.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.1
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• pp.104-109
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• 2005

The aircraft gas turbine engines with the Electronic Engine Controls(EEC) had been developed to save fuel and enhance their performance in the early days, and had employed the health monitoring function in the Full Authority Digital Engine Controls(FADEC) to improve their reliability. This has led to an increasing demand for the certification technology of these controls. The design and certification issues of power supply, aircraft supplied data, failure modes, software verification/validation, and lightning requirements need to be addressed. This paper presents the design considerations and the certification techniques applied to the electronic engine controls. And it is believed that this paper will be basis to establish a requirement in Korean Airworthiness Standard.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.250-257
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• 2007

The development of aircraft engine requires a lot of time and cost to estimate system attributes such as performance, reliability, stability and life. A virtual engine test based on "Numerical test cell" can extremely reduce the time and cost for the development of a hardware by coupling multidisciplinary analyses. This paper presents the development activities of full engine simulation programs in U.S.A. and Europe. NASA Glenn research center of U.S.A. leads the development efforts of NPSS(Numerical Propulsion System Simulation) by assembling the existing codes and improving their functions. VIVACE (Value Improvement through a Virtual Aeronautical Collaborative Enterprise), a consortium of universities, research centers and companies in Europe is developing the PROOSIS(PRopulsion Object Oriented SImulation Software) by integrating the various programs of the institutes. The capability for the domestic development is also estimated by surveying the current status.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.75-78
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• 2004

Various jet engines (Turbine engine family and RAM Jet engine) have been developed for high speed aircrafts. but their application to hypersonic flight is restricted by principle problems such as increase of total pressure loss and thermal stress. Therefore, the development of next generation propulsion system for hypersonic aircraft is a very important subject in the aerospace engineering field, SCRAM Jet engine based on a key technology, Supersonic Combustion. is supposed as the best choice for the hypersonic flight. Since Supersonic Combustion requires both rapid ignition and stable flame holding within supersonic air stream, much attention have to be given on the mixing state between air stream and fuel flow. However. the wider diffusion of fuel is expected with less total pressure loss in the supersonic air stream. So. in this study the direction of fuel injection is inclined 30 degree to downstream and the total pressure of jet is controlled for lower penetration height than thickness of boundary layer. Under these flow configuration both streams, fuel and supersonic air stream, would not mix enough. To spread fuel wider into supersonic air an aerodynamic force, baroclinic torque, is adopted. Baroclinic torque is generated by a spatial misalignment between pressure gradient (shock wave plane) and density gradient (mixing layer). A wedge is installed in downstream of injector orifice to induce an oblique shock. The schlieren optical visualization from side transparent wall and the total pressure measurement at exit cross section of combustor estimate how mixing is enhanced by the incidence of shock wave into supersonic boundary layer composed by fuel and air. In this study non-combustionable helium gas is injected with total pressure 0.66㎫ instead of flammable fuel to clarify mixing process. Mach number 1.8. total pressure O.5㎫, total temperature 288K are set up for supersonic air stream.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.889-891
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• 2017

Usually piston or rotary engines are installed at UAV's under 100 kg payload class. Those engine are less expensive and easy to get, but they require higher operating and maintenance costs due to shorter life and unique fuel usage. They are also too noisy to operate in urban area and have too strong vibration to carry sophisticated payloads. On the contrary, a gas turbine engine has drawbacks like higher specific fuel consumption and weight to power ratio, even it has many operating and maintenance benefits. This study aims to design a small turboshaft engine with a recuperator to overcome those demerits. A tilt rotor UAV(TR-60) developed by KARI was chosen as an imaginary target aircraft, and engine power and size were derived from it. This paper describes engine requirements, design process, and initial reference point cycle design.

• Journal of Aerospace System Engineering
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• v.13 no.1
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• pp.38-46
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• 2019

There is a growing interest in engine diagnostic technology for gas turbine engines. An engine simulation program, precisely simulating the engine performance, is required in order to apply it to the engine diagnosis technology for engine health monitoring. In particular, the simulation program can predict not only design point performance but also off-design point and partial load performance in accurate. So the engine simulation program for the 2-spool separate flow type turbofan engine was developed and the JT9D-7R4G engine of PW(Pratt & Whitney) was analyzed. The steady-sate performance analysis is conducted at both design and off-design points in flight path and the differences between analysis results of takeoff and cruise conditions are compared. The effect of Reynold's correction method was analyzed as a scaling method of the engine component performance. The simulation results was compared with NPSS.

• Tribology and Lubricants
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• v.40 no.1
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• pp.8-16
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• 2024

This study conducts numerical modeling and eigen-analysis of a rod-fastened rotor, which is mainly used in aircraft gas turbine engines in which multiple disks are in contact through curvic coupling. Nayak's theory is adopted to calculate surface parameters measured from the tooth profile of the curvic coupling gear. Surface parameters are important design parameters for predicting the stiffness between contact surfaces. Based on the calculated surface parameters, elastoplastic contact analysis is performed according to the interference between two surfaces based on the Greenwood-Williamson model. The equivalent bending stiffness is predicted based on the shape and elastoplastic contact stiffness of the curvic coupling. An equation of motion of the rod-fastened rotor, including the bending stiffness of the curvic coupling, is developed. Methods for applying the bending stiffness of a curvic coupling to the equation of motion and for modeling the equation of motion of a rotor that includes both inner and outer rotors are introduced. Rotordynamic analysis is performed through one-dimensional finite element analysis, and each element is modeled based on Timoshenko beam theory. Changes in bending stiffness and the resultant critical speed change in accordance with the rod fastening force are predicted, and the corresponding mode shapes are analyzed.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.4
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• pp.104-111
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• 2019

Staged combustion cycle engines exhibit higher combustion performance compared with open cycle engines with a gas generator. An advanced research of the staged combustion cycle engine is going on for the next program following the KSLV-II program. Various experiments have been carried out for the technology demonstration model, TDM0A and TDM0B. The experiments on the combustion performance are aimed to understand the engine start condition and combustion characteristics. They also aim to develop the oxidizer-rich pre-burner and the combustor of the staged combustion cycle engine. The engine-shaped model, TDM1A is fabricated based on the experimental data. The combustion experiment of the TDM1A shows that the combustion pressure of the combustor is approximately 91 bar and the turbine rotation is approximately 28,00 rpm. The result is stable and satisfies the development requirements. The present paper reports on the development process and characteristics of engine models from TDM0A to TDM1A.

• Journal of the Korean Ceramic Society
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• v.50 no.6
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• pp.353-358
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• 2013

The hot corrosion behavior of plasma sprayed 4 mol% $Y_2O_3-ZrO_2$ (YSZ) thermal barrier coatings (TBCs) with volcanic ash is investigated. Volcanic ash that deposited on the TBCs in gas-turbine engines can attack the surface of TBCs itself as a form of corrosive melt. YSZ coating specimens with a thickness of 430-440 ${\mu}m$ are prepared using a plasma spray method. These specimens are subjected to hot corrosion environment at $1200^{\circ}C$ with five different duration time, from 10 mins to 100 h in the presence of corrosive melt from volcanic ash. The microstructure, composition, and phase analysis are performed using Field emission scanning electron microscopy, including Energy dispersive spectroscopy and X-ray diffraction. After the heat treatment, hematite ($Fe_2O_3-TiO_2$) and monoclinic YSZ phases are found in TBCs. Furthermore the interface area between the molten volcanic ash layers and YSZ coatings becomes porous with increases in the heat treatment time as the YSZ coatings dissolved into molten volcanic ash. The maximum thickness of this a porous reaction zone is 25 ${\mu}m$ after 100 h of heat treatment.