سال انتشار: ۱۳۸۵

محل انتشار: چهاردهمین کنفرانس سالانه مهندسی مکانیک

تعداد صفحات: ۸

نویسنده(ها):

Hoseinalipour – Assistant Professor College of Mechanical Engineering, Iran University of Science &Technology Tehran, Iran
Ebrahimi – Assistant Professor College of Mechanical Engineering, Iran University of Science &Technology Tehran, Iran
Sajjadi – Research Engineer College of Mechanical Engineering, Iran University of Science &Technology Tehran, Iran

چکیده:

The modeling of Aerodynamic heating of supersonic and hypersonic flights has been under intensive consideration in recent years.
In present work an Engineering fast code for calculating transient temperature distribution inside blunt nose bodies and internal devices in high Mach flows, considering aerodynamicheating effects. The flow field around the blunt body is divided into three rejoins. The code is evaluated by some existing experimental and numerical results in the literature for existing first stage of pgasus stage booster data. Then the code is used for prediction of temperature variation of two other configuration and the results are presented. In proposed method the local slope method is used for inviscid flow and approximation methods applied for boundary layer near the surface. The temperature variation inside the body is also calculated.
Aerodynamicheatingrises the temperature of the surface which is considerably high at certain regions especially for blunted nose missile in high speed trajectories and it can affect the internal part of missile through radiation and convection into internal elements.
In this study unsteady temperature distribution for internal device is determined by solving the Fourier’s heat conduction equation for internal element.
appropriate correlation is used to predict convention heat transfer coefficient in various regions around an axisymetric body moving at some moderate and high mach numbers, at mid-altitude without ablation and sleep and free molecules. The complicated boundary condition around the heating object is divided into tree regions, stagnation region, near stagnation region which is sphere shape, cone region in which shock expansion theory is applied to compute the inviscid flow.
Thermo physical properties of the air are assumed to vary with flight height and temperature.
For internal element two dimensional energy equation in cylindrical coordinates is used which is coupled with Fourier’s equation.
For this reason, in the present study, a finite difference method as a numerical tool is developed to solve the energy equation in an axisymetric body with a given corresponding boundary conditions.
As a result the developed finite difference program can be used to predict unsteady temperature distribution of internal elements in any axisymetric body with different trajectories.