Abstract

Conjugate heat transfer analysis has been carried out on an 89kN thrust chamber in order to evaluate whether combined discrete film cooling and regenerative cooling in a rocket nozzle is feasible. Several cooling configurations were tested against a baseline design of regenerative cooling only. New designs include combined cooling channels with one row of discrete film cooling holes near the throat of the nozzle, and turbulated cooling channels combined with a row of discrete film cooling holes. Blowing ratio and channel mass flow rate were both varied for each design. The effectiveness of each configuration was measured via the maximum hot gas-side nozzle wall temperature, which can be correlated to number of cycles to failure. A target maximum temperature of 613K was chosen. Combined film and regenerative cooling, when compared to the baseline regenerative cooling, reduced the hot gas side wall temperature from 667K to 638K. After adding turbulators to the cooling channels, combined film and regenerative cooling reduced the temperature to 592K. Analysis shows that combined regenerative and film cooling is feasible with significant consequences, however further improvements are possible with the use of turbulators in the regenerative cooling channels.

Thesis Completion

2016

Semester

Fall

Thesis Chair/Advisor

Kapat, Jayanta

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

Open Access

Length of Campus-only Access

3 years

Release Date

12-1-2019

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