Compressible Flow Simulators

How to Use

1. Select flow regime: subsonic (M < 0.3), transonic (0.3 < M < 1.3), or supersonic (M > 1.3) based on your Mach number calculation
2. Input upstream conditions: stagnation temperature (T0), stagnation pressure (P0), and working fluid (air γ=1.4, helium γ=1.67)
3. Define throat or reference geometry area, then solve for critical conditions, pressure ratios, and velocity using isentropic relations and normal shock tables
4. Export shock-expansion results or oblique shock angles for inlet/nozzle design verification

Worked Example

Air flow through a converging-diverging nozzle: T0=300K, P0=101.3kPa, throat area At=50mm². At Mach 0.5 in subsonic section, static pressure P=87.4kPa and velocity V=170m/s. When flow accelerates to M=2.0 in supersonic section, P drops to 7.6kPa and V=660m/s. Mass flow remains constant at 0.31kg/s. Using shock tables, a normal shock at M=2.0 reduces downstream Mach to 0.577 with pressure rise to 51.8kPa.

Practical Notes

1. Verify Mach number assumptions before selecting compressibility corrections; incompressible models fail beyond M=0.3 in density-sensitive calculations
2. Account for real gas effects (Z-factor) when helium or hydrogen pressures exceed 50 bar or temperatures drop below 100K
3. Check throat conditions for choked flow; mass flow peaks when At/A* ratio equals unity
4. Use oblique shock diagrams for intake designs to prevent shock-boundary layer separation above M=3.5