✈️ 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

📊 Solver Telemetry ● ACTIVE

👁️ Consultations 44

⚡ Calculs faits 37

💾 Téléchargements 354 📦 Code Fortran 10.8 KB

📅 Mise en service Jun 2026

⏱️ Latence < 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 = 0.132 (Centerline η₀ = 0.219)

Adiabatic Wall Temp: 1950.1 °C | Heat Flux Reduction: 19.2 %

ATTACHED FILM JET

Blowing Ratio (M) 1.22 Momentum ratio I = 0.41

Protected Wall Heat Flux 2,310.2 kW/m² Uncooled = 2,860.0 kW/m²

Coolant Density Ratio (DR) 3.67 ρ_coolant / ρ_gas

Normalized Distance (x/D) 24.0 Hole pitch P/D = 3.0

📈 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                 : Hypersonic Scramjet Combustor Strut Film Protection
Hole Geometry              : Diameter D = 2.50 mm, Pitch P = 7.50 mm (P/D = 3.0), Angle = 20.0 deg
Evaluation Location        : x = 60.0 mm (x/D = 24.0)
Mainstream Gas Flow        : Tinf = 2200.0 C, Uinf = 1200.0 m/s, rho = 0.60 kg/m3
Coolant Bleed Flow         : Tc = 300.0 C, Uc = 400.0 m/s, rho = 2.20 kg/m3 (DR = 3.67)
-----------------------------------------------------------------
BLOWING RATIO (M)          : 1.222
MOMENTUM RATIO (I)         : 0.407 [Attached Film]
FILM EFFECTIVENESS (eta)   : 0.1315 (Centerline eta_0 = 0.2192)
ADIABATIC WALL TEMP (Taw)  : 1950.1 deg C
Protected Wall Heat Flux   : 2,310.19 kW/m2
Uncooled Wall Heat Flux    : 2,860.00 kW/m2
Heat Flux Reduction        : 19.22 %
=================================================================

📘 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