• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.50 no.1
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• pp.21-30
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• 2022

The restart of scramjet engine is almost impossible in case its unstart happens during engine operation. Therefore, it is essential to prognosticate the scramjet engine unstart phenomena. A numerical simulation of a scramjet engine is conducted to investigate the unstart process of the scramjet engine by adjusting the backpressure at the isolator outlet to the engine analysis environment. The start and unstart of the engine are identified by applying a support vector machine (SVM) through the data measured by wall pressure so that the locations of the pressure sensors most suitable for the unstart detection are selected. And the operation conditions in which the engine is avoid to be unstarted are quantified to perceive the safety margin.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.388-391
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• 2012

Intake start characteristics of a Mach 5 scramjet engine were investigated experimentally. Intake unstart was observed. The model-facility interaction or excessive internal contraction ratio were attributed to this unstart. Intake start was achieved by changing the model position to remove the model-facility interaction, and by modifying the cowl to give smaller contraction ratio.

• 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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• 2004.03a
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• pp.26-33
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• 2004

In this study, a boundary-layer bleeding and a two-staged fuel injection were applied to a scramjet engine for suppressing unstart transition and improving the thrust performance under Mach 6 flight conditions. With the boundary-layer bleeding, the engine could operate without unstart transition around at the fuel equivalence ratio of unity ($\Phi$ = 1). The thrust increment from the no fuel condition (dF) increased to 2460 N, which was about 1.4 times as large as that of the case without the bleeding and maximum in our Mach 6 tests. It was confirmed that the boundary-layer bleeding suppressed the separation during the engine operation. The two-staged fuel injection was less effective for improving the thrust performance com-pared with the single-staged one with the bleeding at Mach 6.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.26 no.6
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• pp.765-771
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• 2002

This Paper presents numerical solutions of two-dimensional Euler equations for supersonic steady and unsteady flows with heat addition in a convergent-divergent duct, The Van Leer FVS (flux vector splitting) method in generalized coordinates is employed in order to calculate the inviscid strong shock waves caused by thermal choking. We discuss on transient characteristics, start and unstart phenomena caused by thermal choking, limit of equivalence ratio to avoid thermal choking and fluctuation of specific thrust caused by thermal choking. We prove that thermal choking is a serious problem in view of engine performance.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.29 no.1 s.232
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• pp.80-86
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• 2005

This study presents numerical solutions of the two-dimensional Navier-Stokes equations for supersonic unsteady flow in a convergent-divergent nozzle with heat addition. The TVD scheme in generalized coordinates is employed in order to calculate the moving shock waves caused by thermal choking. We discuss on transient characteristics, start and unstart phenomena, fluctuations of specific thrust caused by thermal choking and viscous effects. We prove that the control of separation of boundary layer is the most important key problem to prevent the thermal choking.

• 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.

• 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.

• Advances in aircraft and spacecraft science
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• v.5 no.3
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• pp.349-362
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• 2018

A non-linear numerical simulation technique for predicting the unsteady performances of an airbreathing engine is developed. The study focuses on the simulation of integrated propulsion systems, where a closer coupling is needed between the airframe and the engine dynamics. In fact, the solution of the fully unsteady flow governing equations, rather than a lumped volume gas dynamics discretization, is essential for modeling the coupling between aero-servoelastic modes and engine dynamics in highly integrated propulsion systems. This consideration holds for any propulsion system when a full separation between the fluid dynamic time-scale and engine transient cannot be appreciated, as in the case of flow instabilities (e.g., rotating stall, surge, inlet unstart), or in case of sudden external perturbations (e.g., gas ingestion). Simulations of the coupling between external and internal flow are performed. The flow around the nacelle and inside the engine ducts (i.e., air intakes, nozzles) is solved by CFD computations, whereas the flow evolution through compressor and turbine bladings is simulated by actuator disks. Shaft work balance and rotor dynamics are deduced from the estimated torque on each turbine/compressor blade row.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.31 no.2 s.257
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• pp.109-115
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• 2007

Airframe-integrated Scramjet engines of NASA Langley type consist of a compressor, a combustion chamber and a nozzle. When some disturbances occur in one module of the engine, its influences are propagated to other modules. In this study, it is investigated numerically how shock waves were caused by thermal choking in one module propagate upstream and how they influence adjacent modules. The calculations are carried out in 2-dimensional supersonic viscous flow model using explicit TVD scheme in generalized coordinates. The adverse pressure gradient caused by heat addition brings about separation of the wall boundary layers and formation of the oblique shock wave that proceed to upstream. This moving shock wave formed one module blocks the flow coming into the adjacent modules, which makes the modules unstarted.