Abstract

The propagation of a shock and flame from a detonation wave injected orthogonally into a combustible, supersonic flow was observed. The detonation wave was generated through the use of a miniaturized detonation tube, henceforth referred to as a predetonator. Conditions within the test section, including stagnation pressure and equivalence ratio, were varied between cases. Through the use of high-speed schlieren, shadowgraph, and broadband OH chemiluminescence imaging, the leading shock and reaction were recorded as they moved through the test section. Variation of stagnation pressure affected the propagation of the leading shock. Higher stagnation pressures caused greater deflection of the shock wave and jet issued by the predetonator. It was seen that at sufficiently high equivalence ratios, the shock and reaction were able to travel upstream from the test section into the diverging section of the converging-diverging nozzle. Shortly after the shock entered the nozzle, evidence of the initiation of shock induced combustion was observed. Stagnation pressure variation in the range tested had little effect on the ability to initiate a reaction. Multiple behaviors of the shock-induced-combustion were observed, dependent upon the equivalence ratio of the flow through the test section. Behaviors include sustained reaction on the edges of the flow, sustained reaction in the core of the flow, and periodic, non-sustained reaction.

Thesis Completion

2018

Semester

Fall

Thesis Chair

Ahmed, Kareem

Degree

Bachelor Science in Aerospace Engineering (B.S.A.E.)

College

College of Engineering and Computer Science

Department

Mechanical and Aerospace Engineering

Degree Program

Aerospace Engineering

Location

Orlando (Main) Campus

Language

English

Access Status

Campus Access

Length of Campus-only Access

5 years

Release Date

6-1-2024

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