✈️ Turbine Blade Film Cooling Effectiveness

Evaluate gas turbine film cooling effectiveness (eta), adiabatic wall temperature (Taw), blowing ratio (M), momentum flux ratio (I), and protected wall heat flux reduction.

⚡ Fortran 90 Engine Double Precision (IEEE 754) ✓ ISO / ASME Validated

Turbine Blade Film Cooling Effectiveness

Convection

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⚡ Solves 37

💾 Downloads 490 📦 Fortran Code 10.8 KB

📅 Released Jun 2026

⏱️ Latency < 1 ms

✈️ Turbine Airfoil Blade Wall & Protective Coolant Film Layer

Real-time visual simulation: Angled cooling hole injection discharging protective thermal barrier film

📝 Configuration & Presets

🔥 HPT First-Stage Vane 🌀 Rotor Blade Showerhead 🏭 Combustor Effusion Tile 🚀 Scramjet Hypersonic Film

📐 Cooling Hole Geometry & Distance

Hole Diameter D [mm]:

Spanwise Pitch P [mm]:

Hole Angle α [deg]:

Evaluation Distance x [mm]:

🔥 Mainstream Hot Gas Flow

Gas Temperature T_∞ [°C]:

Gas Velocity U_∞ [m/s]:

Gas Density ρ_∞ [kg/m³]:

Baseline HTC h₀ [W/(m²·K)]:

❄️ Coolant Bleed Air & Blade Metal

Coolant Temp T_c [°C]:

Coolant Jet Velocity U_c [m/s]:

Coolant Density ρ_c [kg/m³]:

Blade Metal Temp T_w [°C]:

Goldstein & Baldauf Formulation:
• Blowing Ratio: M = (ρ_c U_c) / (ρ_∞ U_∞)
• Momentum Flux Ratio: I = (ρ_c U_c²) / (ρ_∞ U_∞²)
• Film Effectiveness: η = (T_aw − T_∞) / (T_c − T_∞)
• Wall Heat Flux: q″ = h₀ · (T_aw − T_w) [kW/m²]

📊 Film Cooling Results

📊 Output Summary
💾 Fortran Source

Film Effectiveness (η)

η = 0.078 (Centerline η₀ = 0.173)

Adiabatic Wall Temp: 1510.4 °C | Heat Flux Reduction: 10.5 %

JET LIFT-OFF REGIME

Blowing Ratio (M) 1.94 Momentum ratio I = 1.45

Protected Wall Heat Flux 456.2 kW/m² Uncooled = 510.0 kW/m²

Coolant Density Ratio (DR) 2.58 ρ_coolant / ρ_gas

Normalized Distance (x/D) 20.0 Hole pitch P/D = 4.0

📈 Film Effectiveness η vs Downstream Distance x (mm)

📉 Protected Heat Flux q″ (kW/m²) vs Blowing Ratio M

=================================================================
 THERMOFLUIDCALC — TURBINE BLADE FILM COOLING REPORT
=================================================================
Case Title                 : Aero-Engine Combustor Liner Effusion Cooling Tile
Hole Geometry              : Diameter D = 2.00 mm, Pitch P = 8.00 mm (P/D = 4.0), Angle = 25.0 deg
Evaluation Location        : x = 40.0 mm (x/D = 20.0)
Mainstream Gas Flow        : Tinf = 1600.0 C, Uinf = 120.0 m/s, rho = 1.20 kg/m3
Coolant Bleed Flow         : Tc = 450.0 C, Uc = 90.0 m/s, rho = 3.10 kg/m3 (DR = 2.58)
-----------------------------------------------------------------
BLOWING RATIO (M)          : 1.938
MOMENTUM RATIO (I)         : 1.453 [Jet Lift-off Risk]
FILM EFFECTIVENESS (eta)   : 0.0780 (Centerline eta_0 = 0.1732)
ADIABATIC WALL TEMP (Taw)  : 1510.4 deg C
Protected Wall Heat Flux   : 456.21 kW/m2
Uncooled Wall Heat Flux    : 510.00 kW/m2
Heat Flux Reduction        : 10.55 %
=================================================================

📘 Calculation Methodology & Film Cooling Standards

Goldstein & Baldauf Correlations

Computes the adiabatic wall temperature $T_{aw}$ resulting from the thermal mixing between the injected cold wall jet and mainstream boundary layer:

                    η = (Taw − T∞) / (Tc − T∞) = f(x/(MD), I, P/D)
                

Jet Attachment vs Lift-Off

When momentum flux ratio $I > 0.8$, the coolant jet lifts off the surface, entraining hot gas beneath before reattaching, degrading near-hole protection.

Key Engineering Assumptions

Rows of discrete circular holes inclined at $\alpha \approx 30^\circ - 35^\circ$.
Turbulent boundary layer mixing over flat or mildly curved airfoil walls.
Applicable to gas turbine vanes, blades, combustors, and scramjet injectors.

ThermoFluidCalc Engineering Report