Ph.D. (Engg): Experimental Study of Isolator Shock Trains in Confined Co-Flowing Supersonic Streams

Futuristic high Mach number flight systems using advanced air-breathing propulsion technologies typically have multiple flow paths with supersonic flows that merge before exiting the vehicle. The supersonic-supersonic co-flow configuration is a canonical model used to study the fundamental aerodynamics of these interactions. A pseudo-shock is a composite gas-dynamic feature produced in viscous-dominated internal flows due to shock-boundary layer interaction. It consists of a series of shocks (shock train) and a mixing region. The terms pseudo-shock and shock train are often used interchangeably. The isolator is a finite-length, constant-area duct that contains the shock train across a wide range of operating conditions. Understanding and predicting the length, adverse pressure handling capacity, and instability of the shock train in the isolator is crucial for designing weight-critical aerospace systems. Most research on shock trains in isolators involves configurations without a co-flowing supersonic stream, where the adverse pressure ratio is imposed mechanically. Fluidic throttling, however, establishes the isolator shock train in a supersonic-supersonic co-flow configuration, which differs fundamentally from mechanical throttling, which necessitates separate investigations. The limited literature on shock trains in supersonic-supersonic co-flow configurations shows the shock train in a narrow operating regime, either in the overexpanded regime or with combustion in the mixed stream producing back pressure. These studies, conducted in opaque tubular ducts, relied on pressure measurements to infer shock train characteristics. Empirical relations of the shock train pressure distribution and length were not in consensus. This thesis aims to understand the shock train in a supersonic-supersonic co-flow configuration using an optically accessible test section that provides simultaneous time-resolved schlieren imaging and static pressure measurement. A wide range of operating conditions is achieved by converting an existing blowdown supersonic jet facility to a pressure-vacuum-driven system. A new modular supersonic-supersonic co-flow test section is established with independent control over Mach number, isolator length, and stagnation conditions of the separate streams, offering a larger parameter space than previous studies. The flow topology and morphology of 158 shock train cases are studied experimentally, leading to several key insights. Novel image analysis techniques and static pressure profile analysis enabled the extraction of the last shock in the shock train, correctly identifying the number of shocks and separating the mixing region. The maximum number of shocks for the supersonic-supersonic co-flow configuration ranges from 6 to 8, and the maximum length of the shock train in the pseudo-shock occupies an average of 6 to 6.5 times the isolator duct height. A major outcome is the revelation of a secondary shock at the isolator duct exit due to local entrainment effects of the supersonic co-flow. This secondary shock can significantly contribute to about 20% to 25% of the overall adverse pressure ratio of the isolator. Consequently, the addition of the secondary shock increases the overall adverse pressure-handling capacity of the isolator to 85% to 90% of the normal shock pressure ratio corresponding to the isolator entrance Mach number. Four transition points are identified based on significant changes in shock train topology. Across various operating conditions and geometries, the normalized adverse pressure ratio (normalized with respect to the normal shock pressure ratio for the isolator entrance Mach number) ranges between 0.4 and 0.85. The flow topology in cases where the core flow is overexpanded is notably different due to the absence of the secondary shock in the shock train and the core flow’s contribution to the overall adverse pressure ratio. A comparative study between fluidic and mechanical throttling is conducted by implementing a mechanical flap module in the same setup. In the mechanically throttled case, the shock train system has a lower adverse pressure ratio than the fluidically throttled case and a higher number of shocks, with a maximum of about 10 to 11. The large dataset produced in this study allows a critical evaluation of well-known empirical correlations for shock trains, leading to a new prediction algorithm to address gaps in their predictive ability. A regression-based correlation is developed to estimate the imposed adverse pressure ratio for the given Mach number and stagnation pressure combinations of both flows. An adaptive pressure increase factor for estimating the shock train leading edge is obtained using a linear regression model for cases with available wall static pressure data. The ratio of the imposed adverse pressure ratio to the incipient pressure ratio for a turbulent boundary layer is used to estimate the initiation of large amplitude oscillations of the shock train leading edge, with an average factor of 2. Spectral analysis of the STLE oscillations using wall static pressure fluctuations and data-driven analysis of schlieren image datasets showed a broad-band spectrum without distinguishable tones, with a spread of less than 200 Hz.

Speaker: A Balaji Himakar

Research Supervisor: Srisha Rao M V