Unsteady flow characteristic analysis of turbine based combined cycle(TBCC)inlet mode transition10.1016/j.jppr.2015.07.006

A turbine based combined cycle(TBCC)propulsion system uses a turbine-based engine to accelerate the vehicle from takeoff to the mode transition flight condition,at which point,the propulsion system performs a“mode transition”from the turbine to ramjet engine.Smooth inlet mode transition is accomplished when flow is diverted from one flowpath to the other,without experiencing unstart or buzz.The smooth inlet mode transition is a complex unsteady process and it is one of the enabling technologies for combined cycle engine to become a functional reality.In order to unveil the unsteady process of inlet mode transition,the research of over/under TBCC inlet mode transition was conducted through a numerical simulation.It shows that during the mode transition the terminal shock oscillates in the inlet.During the process of inlet mode transition mass flow rate and Mach number of turbojet flowpath reduce with oscillation.While in ramjet flowpath the flow field is non-uniform at the beginning of inlet mode transition.The speed of mode transition and the operation states of the turbojet and ramjet engines will affect the motion of terminal shock.The result obtained in present paper can help us realize the unsteady flow characteristic during the mode transition and provide some suggestions for TBCC inlet mode transition based on the smooth transition of thrust.

Unsteady supercritical/critical dual flowpath inlet flow and its control methods

The characteristics of unsteady flow in a dual-flowpath inlet, which was designed for a Turbine Based Combined Cycle（TBCC） propulsion system, and the control methods of unsteady flow were investigated experimentally and numerically. It was characterized by large-amplitude pressure oscillations and traveling shock waves. As the inlet operated in supercritical condition,namely the terminal shock located in the throat, the shock oscillated, and the period of oscillation was about 50 ms, while the amplitude was 6 mm. The shock oscillation was caused by separation in the diffuser. This shock oscillation can be controlled by extending the length of diffuser which reduces pressure gradient along the flowpath. As the inlet operated in critical condition, namely the terminal shock located at the shoulder of the third compression ramp, the shock oscillated,and the period of oscillation was about 7.5 ms, while the amplitude was 12 mm. At this condition,the shock oscillation was caused by an incompatible backpressure in the bleed region. It can be controlled by increasing the backpressure of the bleed region.

Evolution of turbulent boundary layer over a three-dimensional bump

A bump is typically used in the inlet system of an aircraft engine to compress the incoming airflow and to reduce boundary layer thickness developed over fuselage.In this work,the turbulent flow over a three-dimensional bump is experimentally studied.The bump model is mounted in a closed return wind tunnel operated at the nominal velocity 10 m/s,corresponding to a friction Reynolds number of 2300.The flow field upstream the bump,along the bump centerline and at two different spanwise planes is measured with Particle Image Velocimetry(PIV).It is observed that a favorable pressure gradient develops until the suction peak of the bump,and that the average turbulence intensity within boundary layer is attenuated due to this favorable pressure gradient.The boundary layer thickness identified by examining profiles of streamwise velocity decreases significantly along the bump.The wall-normal position of the Turbulent/Non-Turbulent Interface(TNTI)identified with the vorticity criterion is also observed to decrease along the bump.When studying the behavior of the boundary layer thickness at different spanwise positions,we found that it tends to be larger in planes away from the centreline,which suggests that the bump diverts the flow.