Normal Shock Waves

This module provides functions for analyzing normal shock waves, which are compression waves that occur perpendicular to the flow direction in supersonic flows.

Overview

Normal shock waves are sudden compression phenomena that occur when supersonic flow encounters an obstruction or adverse pressure gradient. They are characterized by:

• Instantaneous property changes across a thin wave front
• Entropy increase (irreversible process)
• Mach number reduction from supersonic to subsonic
• Pressure, density, and temperature increases
• Stagnation pressure loss

Theory

Normal shock waves are governed by the Rankine-Hugoniot relations, derived from conservation of mass, momentum, and energy:

Downstream Mach number: $M_2 = \sqrt{\frac{1 + \frac{\gamma-1}{2}M_1^2}{\gamma M_1^2 - \frac{\gamma-1}{2}}}$

Pressure ratio: $\frac{p_2}{p_1} = 1 + \frac{2\gamma}{\gamma+1}(M_1^2 - 1)$

Density ratio: $\frac{\rho_2}{\rho_1} = \frac{(\gamma+1)M_1^2}{2 + (\gamma-1)M_1^2}$

Temperature ratio: $\frac{T_2}{T_1} = \frac{p_2}{p_1} \cdot \frac{\rho_1}{\rho_2}$

Stagnation pressure ratio: $\frac{p_{02}}{p_{01}} = \left(\frac{2\gamma M_1^2 - (\gamma-1)}{\gamma+1}\right)^{\frac{1}{1-\gamma}} \left(\frac{(\gamma+1)M_1^2}{2 + (\gamma-1)M_1^2}\right)^{\frac{\gamma}{1-\gamma}}$

Where subscript 1 denotes upstream conditions and subscript 2 denotes downstream conditions.

Functions

mach_after_normal_shock(M::Float64, gamma::Float64=1.4)

density_ratio_normal_shock(M::Float64, gamma::Float64=1.4)

pressure_ratio_normal_shock(M::Float64, gamma::Float64=1.4)

temperature_ratio_normal_shock(M::Float64, gamma::Float64=1.4)

total_pressure_after_normal_shock(M::Float64, gamma::Float64=1.4)

total_pressure_ratio_normal_shock(M::Float64, gamma::Float64=1.4)

solve_normal(M::Float64, gamma::Float64=1.4)

Function Details

machafternormal_shock

mach_after_normal_shock(M, gamma=1.4)

Calculate the downstream Mach number after a normal shock wave.

Arguments:

• M::Real: Upstream Mach number (must be > 1)
• gamma::Real=1.4: Specific heat ratio

Returns:

• Float64: Downstream Mach number (M₂)

Formula: $M_2 = \sqrt{\frac{1 + \frac{\gamma-1}{2}M_1^2}{\gamma M_1^2 - \frac{\gamma-1}{2}}}$

Example:

julia> mach_after_normal_shock(2.0)
0.5773502691896257

julia> mach_after_normal_shock(3.0)
0.4752229748622233

densityrationormal_shock

density_ratio_normal_shock(M, gamma=1.4)

Calculate the density ratio across a normal shock wave.

Arguments:

• M::Real: Upstream Mach number
• gamma::Real=1.4: Specific heat ratio

Returns:

• Float64: Density ratio (ρ₂/ρ₁)

Formula: $\frac{\rho_2}{\rho_1} = \frac{(\gamma+1)M_1^2}{2 + (\gamma-1)M_1^2}$

Example:

julia> density_ratio_normal_shock(2.0)
2.6666666666666665

julia> density_ratio_normal_shock(5.0)
5.714285714285714

pressurerationormal_shock

pressure_ratio_normal_shock(M, gamma=1.4)

Calculate the pressure ratio across a normal shock wave.

Arguments:

• M::Real: Upstream Mach number
• gamma::Real=1.4: Specific heat ratio

Returns:

• Float64: Pressure ratio (p₂/p₁)

Formula: $\frac{p_2}{p_1} = 1 + \frac{2\gamma}{\gamma+1}(M_1^2 - 1)$

Example:

julia> pressure_ratio_normal_shock(2.0)
4.5

julia> pressure_ratio_normal_shock(3.0)
10.333333333333334

temperaturerationormal_shock

temperature_ratio_normal_shock(M, gamma=1.4)

Calculate the temperature ratio across a normal shock wave.

Arguments:

• M::Real: Upstream Mach number
• gamma::Real=1.4: Specific heat ratio

Returns:

• Float64: Temperature ratio (T₂/T₁)

Formula: $\frac{T_2}{T_1} = \frac{p_2}{p_1} \cdot \frac{\rho_1}{\rho_2}$

Example:

julia> temperature_ratio_normal_shock(2.0)
1.6874999999999998

julia> temperature_ratio_normal_shock(3.0)
1.8076923076923077

totalpressureafternormalshock

total_pressure_after_normal_shock(M, gamma=1.4)

Calculate the downstream stagnation pressure after a normal shock wave.

Arguments:

• M::Real: Upstream Mach number
• gamma::Real=1.4: Specific heat ratio

Returns:

• Float64: Downstream stagnation pressure coefficient

Example:

julia> total_pressure_after_normal_shock(2.0)
8.526315789473685

julia> total_pressure_after_normal_shock(3.0)
12.061224489795916

totalpressurerationormalshock

total_pressure_ratio_normal_shock(M, gamma=1.4)

Calculate the stagnation pressure ratio across a normal shock wave.

Arguments:

• M::Real: Upstream Mach number
• gamma::Real=1.4: Specific heat ratio

Returns:

• Float64: Stagnation pressure ratio (p₀₂/p₀₁)

Example:

julia> total_pressure_ratio_normal_shock(2.0)
0.7208738938053097

julia> total_pressure_ratio_normal_shock(3.0)
0.32834049679486917

solve_normal

solve_normal(M, gamma=1.4)

Complete normal shock analysis - calculate all property changes across the shock.

Arguments:

• M::Real: Upstream Mach number
• gamma::Real=1.4: Specific heat ratio

Returns:

• M2::Float64: Downstream Mach number
• rho2::Float64: Density ratio (ρ₂/ρ₁)
• p2::Float64: Pressure ratio (p₂/p₁)
• p0ratio::Float64: Stagnation pressure ratio (p₀₂/p₀₁)

Example:

julia> sol = solve_normal(2.5)
(M2 = 0.5130161854439888, rho2_ratio = 3.2, p2_ratio = 7.125, p0_ratio = 0.49949999999999994)

julia> println("M₂ = $(round(sol.M2, digits=3))")
M₂ = 0.513

julia> println("Pressure loss = $(round((1-sol.p0_ratio)*100, digits=1))%")
Pressure loss = 50.1%

Applications

Supersonic Inlet Analysis

Analyze normal shock for inlet deceleration:

# Flight conditions
M_flight = 2.2
alt = 11000  # m

# Get atmospheric properties (simplified)
p_amb = 22600   # Pa at 11 km
T_amb = 216.7   # K at 11 km

# Shock analysis for inlet
sol = solve_normal(M_flight)

# Recovery factor
eta_inlet = sol.p0_ratio  # Inlet total pressure recovery

println("Supersonic Inlet Analysis:")
println("Flight Mach: $M_flight")
println("Downstream Mach: $(round(sol.M2, digits=3))")
println("Inlet total pressure recovery: $(round(eta_inlet, digits=3))")
println("Pressure recovery: $(round(eta_inlet*100, digits=1))%")

Shock Strength Analysis

Analyze how shock strength varies with upstream Mach number:

println("M₁\tM₂\tp₂/p₁\tρ₂/ρ₁\tT₂/T₁\tp₀₂/p₀₁\tLoss %")
println("---\t----\t-----\t-----\t-----\t------\t------")

for M1 in 1.1:0.2:4.0
    sol = solve_normal(M1)
    T_ratio = temperature_ratio_normal_shock(M1)
    loss_percent = (1 - sol.p0_ratio) * 100
    
    println("$(round(M1, digits=1))\t$(round(sol.M2, digits=2))\t$(round(sol.p2_ratio, digits=2))\t$(round(sol.rho2_ratio, digits=2))\t$(round(T_ratio, digits=2))\t$(round(sol.p0_ratio, digits=3))\t$(round(loss_percent, digits=1))")
end

Pitot Tube Behind Normal Shock

Calculate pitot pressure measurements behind a normal shock:

# Flight conditions
M1 = 2.0
p_static = 25000  # Pa (ambient static pressure)

# Normal shock analysis
sol = solve_normal(M1)

# Pressures
p2_static = p_static * sol.p2_ratio           # Static pressure behind shock
p02_pitot = p2_static * total_to_static_pressure_ratio(sol.M2)    # Pitot pressure behind shock

# Compare with direct pitot measurement (incorrect for supersonic)
p01_direct = p_static * total_to_static_pressure_ratio(M1)    # What pitot would read in supersonic flow

println("Pitot Tube Analysis in Supersonic Flow:")
println("Freestream Mach: $M1")
println("Ambient static pressure: $(p_static/1000) kPa")

println("\nCorrect measurement (behind shock):")
println("  Static pressure behind shock: $(round(p2_static/1000, digits=1)) kPa")
println("  Pitot pressure: $(round(p02_pitot/1000, digits=1)) kPa")

println("\nIncorrect direct measurement:")
println("  Direct pitot reading: $(round(p01_direct/1000, digits=1)) kPa")
println("  Error factor: $(round(p01_direct/p02_pitot, digits=2))")

Shock Tube Calculations

Calculate conditions in a shock tube:

# Initial conditions (driven section)
p4 = 101325   # Pa (atmospheric)
T4 = 300      # K
M_shock = 5.0 # Shock Mach number

# Calculate conditions behind shock (region 2)
sol = solve_normal(M_shock)
T_ratio = temperature_ratio_normal_shock(M_shock)

p2 = p4 * sol.p2_ratio
T2 = T4 * T_ratio

# Contact surface velocity (region 2 and 3 velocity)
a4 = sqrt(1.4 * 287 * T4)  # Sound speed in region 4
u_contact = a4 * M_shock * (1 - 1/sol.rho2_ratio)

println("Shock Tube Analysis:")
println("Shock Mach number: $M_shock")
println("Initial pressure (region 4): $(p4/1000) kPa")
println("Initial temperature (region 4): $T4 K")

println("\nBehind shock (region 2):")
println("  Pressure: $(round(p2/1000, digits=1)) kPa")
println("  Temperature: $(round(T2, digits=1)) K")
println("  Density ratio: $(round(sol.rho2_ratio, digits=2))")

println("\nContact surface velocity: $(round(u_contact, digits=1)) m/s")

Different Gases

Calculate shock properties for different gases:

M1 = 2.5
gases = [
    ("Air", 1.4),
    ("Helium", 1.67),
    ("Argon", 1.67),
    ("CO₂", 1.3)
]

println("Gas\tγ\tM₂\tp₂/p₁\tρ₂/ρ₁\tp₀₂/p₀₁")
println("---\t----\t----\t-----\t-----\t------")

for (gas, gamma) in gases
    M2 = mach_after_normal_shock(M1, gamma)
    p_ratio = pressure_ratio_normal_shock(M1, gamma)
    rho_ratio = density_ratio_normal_shock(M1, gamma)
    p0_ratio = total_pressure_ratio_normal_shock(M1, gamma)
    
    println("$gas\t$gamma\t$(round(M2, digits=3))\t$(round(p_ratio, digits=2))\t$(round(rho_ratio, digits=2))\t$(round(p0_ratio, digits=3))")
end

Validation Examples

Conservation Laws Check

Verify that the Rankine-Hugoniot relations satisfy conservation laws:

M1 = 3.0
gamma = 1.4

# Calculate all ratios
M2 = mach_after_normal_shock(M1, gamma)
p_ratio = pressure_ratio_normal_shock(M1, gamma)
rho_ratio = density_ratio_normal_shock(M1, gamma)
T_ratio = temperature_ratio_normal_shock(M1, gamma)

# Check ideal gas law: p₂/p₁ = (ρ₂/ρ₁)(T₂/T₁)
ideal_gas_check = p_ratio - (rho_ratio * T_ratio)

# Check momentum conservation
# Should satisfy: p₂ + ρ₂u₂² = p₁ + ρ₁u₁²
# Where u = Ma for convenience (normalized by sound speed)

println("Conservation Laws Verification:")
println("Ideal gas law check: $(round(ideal_gas_check, digits=10))")
println("(Should be ≈ 0)")

Limitations and Considerations

1. One-dimensional flow: Assumes uniform properties across the shock front
2. Steady flow: Time-invariant shock position
3. Perfect gas: Constant specific heats and ideal gas behavior
4. No viscous effects: Inviscid flow assumption
5. Thin shock: Shock thickness much smaller than characteristic length scales

For three-dimensional effects, unsteady phenomena, or real gas behavior, additional analysis methods are required.

See Also

• Oblique Shock Waves: For angled shock waves
• Isentropic Relations: For shock-free flow analysis
• Prandtl-Meyer Expansion: For expansion waves
• Nozzle Analysis: For internal flow applications