GitHub - Arktor/shockexpansion

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Name		Name	Last commit message	Last commit date
Latest commit History 16 Commits
test		test
README		README
check_empty_string.cpp		check_empty_string.cpp
get_token.cpp		get_token.cpp
main.cpp		main.cpp
makefile		makefile
parse_density.cpp		parse_density.cpp
parse_geometry.cpp		parse_geometry.cpp
parse_geometry_point.cpp		parse_geometry_point.cpp
parse_mach_number.cpp		parse_mach_number.cpp
parse_pressure.cpp		parse_pressure.cpp
parse_problem_def.cpp		parse_problem_def.cpp
parser_out.cpp		parser_out.cpp
pass_empty_strings.cpp		pass_empty_strings.cpp
shexp.hpp		shexp.hpp
shexpexception.cpp		shexpexception.cpp
shexpexception.hpp		shexpexception.hpp
shexpnode.cpp		shexpnode.cpp
shexpnode.hpp		shexpnode.hpp
shexppanel.cpp		shexppanel.cpp
shexppanel.hpp		shexppanel.hpp
solve_expansion.cpp		solve_expansion.cpp
solve_shock.cpp		solve_shock.cpp
solver_out.cpp		solver_out.cpp

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SHOCKEXPANSION

    Calculates drag force applied to a symmetrical airfoil in the supersonic
adiabatic flow at the zero angle-of-attack.
Physical model of the fluid is the ideal nonviscous gas without heat transfer.


METHOD DESCRIPTION

    The method used for drag calculation in this application is called
"shock-expansion". Its basic idea is that supersonic flow around an airfoil
can be represented with a system of shock waves and expansion waves.

    An airfoil is partitioned into a number of segments called "panels".
Flow on each panel is considered to be homogeneous and parallel to its surface.
If parameters of the flow on a certain panel are known, parameters of the flow
on the adjacent panel downstream can be calculated using equations of a shock
wave or expansion wave. Type of equations depends on relative slope of the panel
downstream. If its slope is larger than the one of the panel upstream, shock wave
equations are used, otherwise flow parameters are calculated using expansion wave
equations.


USAGE

Application usage: shockexpansion.exe <problem_definition_file>

Example of a problem definition file:

****************************************
mach_number = 4.0
pressure = 100000.0
density = 1.225


#geometry#
0.     0.
0.0625 0.0106
0.125  0.0202
0.1875 0.0292
0.25   0.0375
0.3125 0.0449
0.375  0.0518
0.4375 0.0577
0.5    0.0628
0.5625 0.0669
0.625  0.0701
0.6875 0.0720
0.75   0.0726
0.8125 0.0716
0.875  0.0688
0.9375 0.0633
1.     0.
****************************************

Application generates an output file ("out.dat") and writes solving process
log and calculation results into it.


LIMITATIONS

1. Airfoil geometry must be scaled to have its first node
   at (0.0, 0.0) and its last node at (1.0, 0.0); 
2. Since only symmetrical airfoil can be calculated, you should provide
   only its upper contour;
3. Too large relative slope values can lead for the flow to become subsonic
   or to separate, which will terminate the calculation.


TO DO

1. Improve application output and exception handling;
2. Optimize solving algorithm: improve accuracy and solving speed.
3. Add the ability to set angle-of-attack;
4. Add flow separation calculation for cases of very high speeds ang large
   slopes.