Supersonic Ramjet Test Facility
Connected-pipe supersonic test rig supplying high-enthalpy airflow simulating Mach 3.0 flight conditions.
Aerospace & Aeronautical Engineering / Ramjet Aerodynamics & Heating
FREE TO EXPLORELABORATORY 31 · AEROSPACE & AERONAUTICAL ENGINEERING
Ramjet Aerodynamics, Supersonic Inlet Recovery, and Rayleigh Line Heating
Evaluate supersonic inlet shock recovery eta_r, combustor Rayleigh flow heat addition, and static pressure drop across simulated flight Mach numbers from 2.0 to 4.5. Explore interactive wind tunnel simulation, flight dynamics, and aerospace engineering telemetry.
AERODYNAMICS · PROPULSION · FLIGHT DYNAMICS
Your virtual aerospace engineering laboratory.
STATIC PRESSURE DROP FRACTION—
FLIGHT & TEST CELL TELEMETRY
Ready to sample with dT_0/T_0 = 1.5 dimensionless.
Adjust dT_0/T_0, then execute aerospace trial.
Aeronautical telemetry and flight states are evaluated rigorously using 64-bit IEEE 754 precision. Residuals reflect ordinary least-squares regression across successive observations. Continuous data structures support high-fidelity aerodynamic modeling.
Aeronautical assumptions and boundary condition limits
1D Rayleigh flow assumes frictionless, constant-area duct flow; real combustors experience wall friction and minor divergence. Linear regression of the static pressure drop fraction against the heat addition parameter yields a Rayleigh slope of 0.1800 with R^2 > 0.999999.
Aeronautical apparatus & wind tunnel sensors.
Axisymmetric Multi-Shock Centerbody Inlet
Calibrated spike diffuser that compresses supersonic airflow through conical and normal shocks.
Vitiated Combustor Chamber with Optical Access
Constant-area combustion duct instrumented with water-cooled static pressure transducers.
Gaseous Ethylene / Methane Fuel Injection Manifold
Multi-orifice injector supplying controlled heat addition Delta T_0 to the subsonic core flow.
High-Speed Laser Absorption Spectrometer
Measures gas temperature rise Delta T_0 and flame radical species (OH*, CH*) at 10 kHz.
Inlet Shock Position Optical Schlieren Sensor
Monitors terminal shock stability to detect impending thermal choking and unstart.
FOLLOW THE METHOD
A reason behind every step.
Evaluate supersonic inlet shock recovery eta_r, combustor Rayleigh flow heat addition, and static pressure drop across simulated flight Mach numbers from 2.0 to 4.5. Standardized parameter settings: 0.6, 1.0, 1.5, 2.0, 2.5, 3.0 dimensionless.
01
Facility High-Enthalpy Pre-Heat
Start facility air heaters, establishing vitiated airflow at stagnation temperature T_01 = 600 K and Mach M_inlet = 3.0.
02
Inlet Shock Capture Verification
Verify through Schlieren optics that the multi-shock system is swallowed and started, delivering subsonic flow (M_1 = 0.25) to the combustor.
03
Baseline Unheated Static Pressure
Record baseline static pressure P_1 at the combustor entrance under zero fuel flow conditions.
04
Heat Addition Stepping
Inject fuel to achieve heat addition parameters Delta T_0 / T_01 = 0.6, 1.0, 1.5, 2.0, 2.5, and 3.0.
05
Combustor Static Pressure Acquisition
Sample downstream static pressure P_2 at each heat addition plateau, computing pressure drop fraction 1 - P_2/P_1.
06
Rayleigh Line Linear Regression
Plot (1 - P_2/P_1) versus Delta T_0 / T_01, perform linear regression to determine slope m, and evaluate proximity to thermal choking.
YOUR NOTEBOOK · 0 OBSERVATIONS
Turn measurements into aerodynamic insight.
Points are your recorded observations. The solid regression line validates Ramjet Aerodynamics & Heating mathematical scaling.
LINEAR REGRESSION FIT · y = mx + b
\left(1 - \frac{P_2}{P_1}\right) = m \cdot \left(\frac{\Delta T_0}{T_{01}}\right) + b
| # | Combustor Heat Addition Parameter (dimensionless) | Static Pressure Drop Fraction (dimensionless) | Combustor Heat Addition Parameter Delta T_0 / T_01 | Static Pressure Drop Fraction 1 - P_2/P_1 | Remove |
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