Computations Supersonic Flow over a Missile Afterbody Containing an Exhaust Jet.

Abstract

A thin-layer Navier-Stokes code, developed for projectile aerodynamics, has been used to compute the supersonic flow over a missile afterbody containing a centered exhaust jet. The thin-layer, compressible, Navier-Stokes equations are solved using a time dependent, implicit numerical algorithm. A unique flow field segmentation procedure is used which preserves the sharp base corner and facilitates the adaption of the grid to the free shear layer in the base region. Solutions have been obtained for an axisymmetric, boattailed afterbody where the free stream Mach number is 2.0 and the jet exit Mach number is 2.5. Computations were made at various jet static pressure to free stream static pressure ratios from 1 through 9. Qualitative features of the base region flow field seen experimentally are very well observed in the computed results. Quantitative comparisons of base pressure with experiment indicate good agreement at high pressure ratios and some disagreement at low pressure ratios.

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Document Details

Document Type
Technical Report
Publication Date
Sep 01, 1986
Accession Number
ADA174295

Entities

People

  • Jubaraj Sahu

Organizations

  • Ballistic Research Laboratory

Tags

Communities of Interest

  • Energy and Power Technologies
  • Weapons Technologies

DTIC Thesaurus Topics

  • Base Pressure
  • Boundary Layer
  • Computational Fluid Dynamics
  • Computations
  • Equations
  • Flow
  • Flow Fields
  • Fluid Dynamics
  • Fluid Flow
  • Free Stream
  • High Pressure
  • Mach Number
  • Navier Stokes Equations
  • Shock Waves
  • Static Pressure
  • Supersonic Flow
  • Turbulent Mixing

Fields of Study

  • Physics

Readers

  • Aerodynamics/Aeronautics.
  • Computational Fluid Dynamics (CFD)
  • Fluid Mechanics and Fluid Dynamics.

Technology Areas

  • Hypersonics
  • Hypersonics - Hypersonic Flow