HEAT TRANSFER TO THE THROAT REGION OF A SOLID PROPELLANT ROCKET NOZZLE
Abstract
A combined experimental and analytical method for obtaining the surface heat-transfer rate in a rocket nozzle was developed at the Naval Ordnance Laboratory. This method is particularly appli cable to high energy rocket nozzle flow where instrumentation directly on the flow surface is impractical. The method employs data of the temperature-versus-time history of two points within the nozzle wall with one of them near the surface of the nozzle. The temperature distri bution between the two points and the temperature of the nearby nozzle surface are computed on the IBM 7090 using the implicit numerical solution to the one-dimensional transient heat conduction equation. The heat-transfer rate at the nozzle surface is then calculated from the computed temperature gradient at the surface. Application of this method to determine the heat-transfer rate at the throat of a molybdenum insert in a conical solid propellant rocket nozzle is presented. The nozzle was operated at nominal chamber conditions of 1150 psia and 2500 K. The experimental data are compared with theoretical predictions and other available experimental results. Good agreement is obtained with turbulent heat-transfer rates computed from the numerical integration of the boundary-layer momentum equation. (Author)
Document Details
- Document Type
- Technical Report
- Publication Date
- Feb 26, 1963
- Accession Number
- AD0406455
Entities
People
- Roland E. Lee
Organizations
- Naval Ordnance Laboratory