In certain applications, different regions of the computational domain experience flow conditions that are so different that it is very difficult for a single solver to produce accurate results at the extremes. In many situations, such problems can be separated and solved using loosely coupled solvers. Each solver is chosen to provide highly accurate solutions for the prevailing flow conditions.
ESI's CFD-FASTRAN, a compressible flow solver, is ideally suited for high speed external aerodynamics problems and the multi-physics solver CFD-ACE+ is ideally suited for heat transfer problems involving conduction, convection (natural and forced) and radiation.
A large amount of aerodynamic heating is generated over hypersonic vehicles during re-entry. Thermal Protection System (TPS) materials are employed to prevent the heating from conducting into the internal cabin, which holds electronic devices, passengers and other vital components. As a time-dependent process, the material making up TPS is at a low temperature and “soaks up” the heat – the conductivity of the material transports the heat (from the vehicle surface) through the thickness. The material will also re-radiate some of the heat back to the flow – the amount depending on the emissivity of the material. A primary concern is to estimate the effects of aeroheating on the internal volume of the capsule and its effect on electronic devices, passengers and cooling systems. In this application, typically the external flow is hypersonic in nature, whereas the flow within the capsule is a very low speed flow dominated by natural convection. In addition, to hypersonic aerodynamic heating, several other physics including heat conduction, natural convection and radiation has to be accurately modeled. CFD-FASTRAN solves for the external hypersonic flow and CFD-ACE+ solves for heat conduction, convection and radiation. Exchange of heat flux/temperature data between FASTRAN and ACE occurs at defined interfaces.
The problem to be simulated is the inviscid, subsonic flow of air past a cylinder. The diameter of the cylinder is 1 m. The flow has a free-stream Mach number,M∞ , of 0.177. The numerical model employs only a semicylinder due to the symmetry of the flow pattern around the cylinder.
Low Mach preconditioning is a way to accelerate convergence towards steady state solution by scaling the disparate eigenvalues of a system to the same order of magnitude for time-marching schemes. The preconditioning matrix applied in FASTRAN is chosen in such a way to provide an efficient solution for both incompressible (through artificial compressibility) and low Mach flows (through pseudo-acoustic speeds). This feature in FASTRAN is demonstrated using steady flow simulation over a 2D cylinder at a freestream Mach number of 0.0004.
Model dependencies can be used to simulate complex rigid body motions. In this example, a pitching airfoil with flap is used to demonstrate this feature.
This unsteady simulation involves a moving body and demonstrates the
useof chimera and 6-DOF modeling features in CFD-FASTRAN. This tutorial will be setup to run in Parallel. The flow has
a free stream Mach number of 2.0 at AOA of 5 deg. The free-stream
temperature and pressure are 101325Pa and 288.16K, respectively. The
simulation includes two separate 6DOF motion models. 6DOF model # 1
governs the motion of the second stage (payload vehicle), and 6DOFmodel
# 2 governs the motion of the first stage (booster vehicle). The
payload vehicle has a rocket nozzle that is modeled with a time
dependent inlet condition simulating rocket ignition. First a
steady-state solution of the combined vehicle flying at 5 deg. angle of
attack is obtained. Then at time t=0, the rocket motor ignites and
pressure builds up between the stages resulting in the separation of
the two vehicles. The thrust integration option is employed to account
for the thrust component at the nozzle chamber.
Turbulent subsonic flow of air past a cylinder is modeled. Diameter of the cylinder is 1m. The flow has a free-stream Mach number,M, of 0.5.The free stream temperature and pressure are 300K and 1 x 105 Pa, respectively.The Reynolds number, Re, of the flow, based on the chord length of the airfoil, is 9 x 106.The computational domain is modeled with Chimera technology using an O-mesh around the cylinder which is overset on a Cartesian background mesh.
The turbulent flow past a NACA 0012
airfoil is modeled. The flow has a free-stream Mach
number,M, of 0.55 at an angle of attack of 8.34 degrees. The Reynolds number, Re, of the
flow, based on the chord length of the airfoil, is 9x 106.
The problem to be simulated is the turbulent flow past a NACA-0012
airfoil. The flow has a free-stream Mach number, M, of 0.55 at an
angle of attack, alpha, of 8.34 degrees. The Reynolds number, Re, of the flow,
based on the chord length of the airfoil, is 9x106. For this case,
the flowfield develops a supersonic bubble near the leading edge of the airfoil
upper surface. Furthermore, the flow is slightly separated at the foot of the shock that terminates the supersonic region. For this
problem, the k-epsilon turbulence model is employed.
The problem to be simulated is inviscid, supersonic flow of air past a blunt body. The numerical model employs only one half of the body due to the symmetry of the flow pattern. The flow has a free-stream Mach number, M∞, of 23.5. Due to high free stream Mach number, the flow develops high temperatures which initiates chemical reactions between the various components of air. These reactions include 1) dissociation of diatomic Oxygen 2) dissociation of diatomic Nitrogen, 3) dissociation of nitrous Oxide 4) reaction of diatomic Nitrogen with oxygen and 5) reaction of Nitrous Oxide with Oxygen.
