• Advances in aircraft and spacecraft science
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• v.5 no.4
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• pp.431-444
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• 2018

Numerical simulations have been conducted to study the unstart/restart characteristics of an over-under turbine-based combined-cycle propulsion system (TBCC) inlet during the inlet transition phase. A dual-solution area exists according to the Kantrowitz theory, in which the inlet states may be different even with the same input parameters. The entire transition process was divided into five stages and the unstart/restart hysteresis loop for each stage was also obtained. These loops construct a hysteresis surface which separates the operating space of the engine into three parts: in which a) inlet can maintain a started state; b) inlet keeps an unstarted state; c) inlet state depends on its initial state. During the transition, the operation of the engine follows a certain order with different backpressures and splitter angles, namely control route, which may result in disparate inlet states. Nine control routes with different backpressures and transition stages were designed to illuminate the route-dependent behavior of the inlet. The control routes operating towards the unstart boundary can make the inlet transit from a started state into an unstarted one. But operating backward the same route cannot make the inlet restart, additional effort should be made.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.334-342
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• 2005

A numerical analysis is performed using two dimensional axisymmetric RANS (Reynolds Averaged Navier-Stokes) equations system on the transition sequence of the Integrated Rocket Ramjet and the unsteady reacting flow-field in a ramjet combustor during unstable combustion. The mode transition of an axisymmetric ramjet is numerically simulated starting from the initial condition of the boost end phase of the entire ramjet. The unsteady reacting flow-field within combustor is computed for varying injection area. In calculation results of the transition, the terminal normal shock is occurred at the downstream of diffuser throat section and no notable combustor pressure oscillation is observed after certain time of the inlet port cover open. For the case of a small injection area at the same equivalence ratio, periodic pressure oscillation in the combustor leads to the terminal shock expulsion from the inlet and hence the buzz instability occurred.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.513-520
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• 2004

A steady-state/transient performance simulation model was newly developed for the propulsion system of the CRW (Canard Rotor Wing) type UAV (Unmanned Aerial Vehicle) during flight mode transition. The CRW type UAV has a new concept RPV (Remotely Piloted Vehicle) which can fly at two flight modes such as the take-off/landing and low speed forward flight mode using the rotary wing driven by engine bypass exhaust gas and the high speed forward flight mode using the stopped wing and main engine thrust. The propulsion system of the CRW type UAV consists of the main engine system and the duct system. The flight vehicle may generally select a proper type and specific engine with acceptable thrust level to meet the flight mission in the propulsion system design phase. In this study, a turbojet engine with one spool was selected by decision of the vehicle system designer, and the duct system is composed of main duct, rotor duct, master valve, rotor tip-jet nozzles, and variable area main nozzle. In order to establish the safe flight mode transition region of the propulsion system, steady-state and transient performance simulation should be needed. Using this simulation model, the optimal fuel flow schedules were obtained to keep the proper surge margin and the turbine inlet temperature limitation through steady-state and transient performance estimation. Furthermore, these analysis results will be used to the control optimization of the propulsion system, later. In the transient performance model, ICV (Inter-Component Volume) model was used. The performance analysis using the developed models was performed at various flight conditions and fuel flow schedules, and these results could set the safe flight mode transition region to satisfy the turbine inlet temperature overshoot limitation as well as the compressor surge margin. Because the engine performance simulation results without the duct system were well agreed with the engine manufacturer's data and the analysis results using a commercial program, it was confirmed that the validity of the proposed performance model was verified. However, the propulsion system performance model including the duct system will be compared with experimental measuring data, later.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.563-564
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• 2010

A Numerical analysis of preventing pressure oscillation for the supersonic inlet of IRR at the boosting mode was conducted in this study. To reduce the fluctuation of pressure, the supersonic inlet was equipped with annular slit. By decreasing the entrained air mass, the amplitude of fluctuation and maximum pressure at the inlet were decreased while the frequency was increased. In case, the optimal trade-off value is obtained between frequency and amplitude, it is expected that the pressure oscillation declining method using the slit will be helpful in various boosting and transition mode problem for IRR.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.6
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• pp.21-31
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• 2017

In the present study, theoretical and numerical analyses have been carried out to investigate the detailed flow characteristics during the mode transition. The theoretical analysis rearranged the knowledge of gasdynamics and the previous studies, and the numerical analysis has conducted to solve the 2D unsteady compressible Navier-Stokes equations with a fully implicit finite volume scheme. To validate the numerical analysis, the experiment was compared with it. The total temperature at the inlet of isolator and the hydrogen fuel equivalent ratio were changed to investigate their effects on the mode transition phenomenon. As the results, the numerical analysis reproduced well the experiment qualitatively, the increment in the hydrogen fuel equivalent ratio induced the scram-mode to ram-mode transition which is discontinuous with a non-allowable region, and the variation in the total temperature changed the boundary of the mode transition.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.1
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• pp.45-51
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• 2018

The generation mechanism of NAR is not yet understood. In the present study, an in-depth analysis of the computational results previously obtained by the authors is conducted to investigate the flow mechanism responsible for NAR. A theoretical analysis has also been performed to understand the gas dynamic features during transition from scramjet to ramjet mode. It is known that there exists a critical value of the fuel equivalence ratio at which the flow states at the inlet of isolator remain unchanged. An increase in the equivalence ratio over the critical value leads to a sudden change in the static pressure and the Mach number at the inlet of the isolator, which is responsible for the generation of NAR.

• Journal of ILASS-Korea
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• v.27 no.1
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• pp.18-25
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• 2022

Hot-firing tests were performed to experimentally confirm the effect of the eigenmode in the fuel-air mixing section on combustion instability by changing mixing section length, inlet mean velocity, equivalence ratio, and swirler geometry. A premixed gas composed of air and ethylene was supplied to the combustion chamber through an mixing section and an axial swirler. As the mixing section length increased, the inlet velocity perturbation decreased, but the combustion instability increased more. It was found that the resonance frequency of the first longitudinal mode in the mixing section shifted to the third longitudinal mode as the length of the mixing section increased. The results implied that the transition of the resonace frquency by changing the length of the mixing section might cause combustion instability.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.499-505
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• 2004

A Propulsion System of the CRW(Canard Rotor Wing) type UAV(Unmanned Aerial Vehicle) was composed of the turbojet engine to generate the propulsive exhaust gas, and the duct system including straight bent ducts, tip-jet nozzles, a master valve and a variable main nozzle for three flight modes such as lift/landing mode, low speed transition flight mode and high speed forward flight mode. In this study, in order to operate safely the propulsion system, the dynamic Performance behavior of the system was modeled and simulated using the SIMULIN $K^{ }$, which is the user-friendly GUI type dynamic analysis tool provided by MATLA $B^{ }$. In the transient performance model, the inter-component volume model was used. The performance analysis using the developed models was performed at various flight condition, valve angle positions and fuel flow schedules, and these results could set the safe flight mode transition region to satisfy the inlet temperature overshoot limitation as well as the compressor surge margin. Performance analysis results using the SIMULIN $K^{ }$ performance program were compared with them using the commercial program GSP.m GSP.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.4
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• pp.76-83
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• 2004

A Propulsion System of the CRW(Canard Rotor Wing) type UAV(Unmanned Aerial Vehicle) was composed of the turbojet engine to generate the propulsive exhaust gas, and the duct system including main and rotary ducts, the nozzle subsystem including main and tip jet nozzle for three flight modes such as lift/landing mode, low speed transition flight mode and high speed forward flight mode. Transient simulation performance utilized the ICV (Inter-component volume) method and simulated using the SIMULINK. Transient performance analysis was performed on 3 cases. Fuel flow schedules to accelerate from Idle to maximum rotational speed were divided into the step increase of the most severe case and ramp increase cases to avoid the overshoot of turbine inlet temperature, and variations of thrust and the turbine inlet temperature were investigated in some transient analysis cases.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.5
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• pp.426-437
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• 2007

off-design conditions, supersonic air inlets often encounter the problem of aerodynamic instability, called inlet buzz, which causes the significant degradation of the engine performance. An experimental and numerical study was conducted to investigate the phenomenon of supersonic inlet buzz on a generic, axisymmetric, external-compression inlet with a single-surface center-body. It is understood the mechanism of buzz onset as proving that the origin of buzz is the flow choking induced by separation at the intake throat. Also it is observed the intermittent and continuous buzz mode as area ratio varies and understood the transition process through this study. The buzz frequency become to be higher as decreasing the area ratio, but for each area ratio, the frequency of pressure oscillation is the same at all points of intake.