Understanding start–unstart behavior of intakes in hypersonic Mach numbers is essential for seamless operation of scramjet engines. We consider a high compression ratio intake (CR = 40) at a Mach number of M = 6 in this work. Start–unstart characteristics are studied in a hypersonic wind tunnel at a flight realistic Reynolds number (Re = 8.7 × 106/m, M = 6). A flap provided at the rear end of the isolator simulates the effect of backpressure for throttling ratios in the range of 0–0.69. Experiments are conducted in two modes: (a) with the flap fixed at a particular throttling ratio and (b) the flap moved to a particular throttling ratio after the started flow has been established. Unsteady pressure measurements and time-resolved Schlieren visualization are undertaken. Modal analysis of pressure (using fast Fourier transform) and Schlieren images (using dynamic mode decomposition) are carried out. The intake shows started behavior for throttling ratios up to 0.31 and a dual behavior, where it remains started in dynamic flap runs but unstarted in fixed flap runs for throttling ratios of 0.35 and 0.42. The intake exhibits a staged evolution to a large amplitude oscillatory unstart for throttling ratios of 0.55 and 0.69, with frequencies of 950 Hz and 1100 Hz, respectively. For the first time, a staged evolution (5 stages) to a subsonic spillage oscillatory unstart of a hypersonic intake is detailed using corroborative evidence from both time-resolved Schlieren and pressure measurements. A precursor to the final large amplitude oscillatory unstart is identified, and the flow mechanism for sustained oscillations is explained.

In the present work, we report on an experimental study of bending/flap mode flutter of a blade within a linear cascade at transonic conditions. Driven by the motivation to understand the contribution of shock location/dynamics to flutter characteristics, we have performed simultaneous measurements of shock dynamics using high-speed shadowgraphy combined with unsteady load measurements on an oscillating blade within the cascade. The flutter characteristics in terms of energy transfer from the fluid to the blade and shock dynamics have been mapped out over a range of blade oscillation frequencies and static pressure ratios (SPRs) across the cascade, the latter being important as they decide the mean location of the passage shocks. SPR values studied include both conditions where the shock is within the passage (started cascade) and where the passage shock is pushed ahead of the leading edge of the blades (unstarted cascade). These measurements show characteristically different flutter behavior for an unstarted cascade compared with a started cascade, the former having received very little attention in the literature. While both these cases show small excitation levels at low reduced frequencies, the unstarted cascade case exhibits an additional relative narrow region of excitation at higher reduced frequencies with approximately an order of magnitude higher excitation energies. Comparison of the shock dynamics between the two excitation regimes shows significant differences in the phase of the leading edge shock in addition to changes in the suction side shock phase indicating that the two excitation regimes are of different origin.

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Abstract

Vibration related issues such as flutter have always been a cause of concern for aircraft engine designers. They not only incur unwarranted cost and time overruns, but also significantly compromise performance and can cause structural damage. This phenomenon has become more relevant for the modern aircraft engines, which employ relatively thin, long blade rows to satisfy ever growing demand for a powerful yet compact engine. The tip sections of such blade rows operate with supersonic relative velocity, where prediction of flutter can get challenging due to unsteady flow features like oscillating shocks and their interaction with the blade motion. Linear cascades that represent a specific radial location of the rotor have proven to be a reliable tool for flutter studies. To facilitate flutter experiments at flow Mach numbers realistic to the aircraft engine components, a transonic cascade facility operating at a Mach Number (M) of 1.3 with the ability to oscillate the central blade in the cascade has been developed. The cascade consists of 5 blades and two false blades of which the central blade is oscillated in heave, which represents the bending mode of the rotor. The typical reduced frequencies associated with this kind of flutter in practice (k ∼ 0.1) correspond to a high dimensional frequency of 200 Hz for the present case. A barrel cam mechanism is used to provide such high frequency oscillations. The parameters varied in the present study include the reduced frequency (k) and the static pressure ratio (SPR) across the cascade, which is varied with the help of tailboard and flap arrangement located at the back end of the cascade. Three SPR cases of 1.05, 1.25, and 1.35 are considered and at each of these pressure ratio cases, the reduced frequency is varied. The unsteady loads are measured on the oscillating central blade during the oscillation cycle to quantify the energy transfer from flow to blade and shadowgraphy is used to visualize the shocks. The results from these experiments indicate flutter at lower k values for all the SPR cases tested, while the higher k values are damped. The magnitude of excitation or damping at any particular frequency is also observed to increase with increasing SPR.