Supersonic Wind Tunnel Test Section
Continuously variable supersonic duct operating from Mach M = 1.2 to 2.5.
Aerospace & Aeronautical Engineering / Normal Shock Wave Relations
FREE TO EXPLORELABORATORY 12 · AEROSPACE & AERONAUTICAL ENGINEERING
Normal Shock Wave Pressure and Total Pressure Loss in Supersonic Ducts
Measure static pressure jump P_2 / P_1 and total pressure ratio P_02 / P_01 across a stationary normal shock wave in a supersonic duct as upstream Mach M_1 varies. Explore interactive wind tunnel simulation, flight dynamics, and aerospace engineering telemetry.
AERODYNAMICS · PROPULSION · FLIGHT DYNAMICS
Your virtual aerospace engineering laboratory.
PRESSURE JUMP RATIO—
FLIGHT & TEST CELL TELEMETRY
Ready to sample with M_1^2 - 1 = 1.5 dimensionless.
Adjust M_1^2 - 1, 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
Shock-wave boundary layer interaction (SWBLI) creates a lambda-foot near duct walls that slightly smears the idealized 1D jump. Linear regression of the pressure jump ratio against (M_1^2 - 1) yields a slope of 1.1667 matching the Rankine-Hugoniot relation with R^2 > 0.999999.
Aeronautical apparatus & wind tunnel sensors.
Variable Back-Pressure Exit Throttle Valve
Regulates downstream back-pressure to position a stable normal shock wave in the test section.
Multi-Channel Fast-Response Pressure Scanner
Samples wall static pressures upstream and downstream of the shock wave at 1 kHz.
Supersonic Pitot-Static Total Pressure Rake
Measures total pressure upstream P_01 and downstream P_02 across the shock discontinuity.
High-Speed Schlieren Imaging Camera
Captures optical shadowgraph and Schlieren images of the normal shock wave thickness and location.
Digital Environmental DAQ Computer
Logs stagnation conditions and computes experimental Rankine-Hugoniot pressure ratios.
FOLLOW THE METHOD
A reason behind every step.
Measure static pressure jump P_2 / P_1 and total pressure ratio P_02 / P_01 across a stationary normal shock wave in a supersonic duct as upstream Mach M_1 varies. Standardized parameter settings: 0.8, 1.25, 1.89, 2.61, 3.41, 4.29 dimensionless.
01
Supersonic Flow Stabilization
Start the supersonic wind tunnel, establishing stable supersonic flow at nominal Mach M_1 = 1.34 in the test section.
02
Normal Shock Wave Positioning
Gradually close the downstream throttle valve until Schlieren optics confirm a stationary normal shock centered over tap station 12.
03
Upstream and Downstream Pressure Acquisition
Record upstream static pressure P_1, downstream static pressure P_2, and stagnation pressures P_01 and P_02.
04
Systematic Mach Number Stepping
Adjust the supersonic nozzle block to step upstream Mach M_1 through values corresponding to (M_1^2 - 1) = 0.80, 1.25, 1.89, 2.61, 3.41, and 4.29.
05
Rankine-Hugoniot Jump Calculation
Compute static pressure jump ratio (P_2/P_1 - 1) and total pressure ratio P_02 / P_01 at each Mach operating point.
06
Rankine-Hugoniot Linear Regression
Plot (P_2/P_1 - 1) versus (M_1^2 - 1), perform linear regression to determine slope 2*gamma/(gamma+1), and verify gamma = 1.4 for air.
YOUR NOTEBOOK · 0 OBSERVATIONS
Turn measurements into aerodynamic insight.
Points are your recorded observations. The solid regression line validates Normal Shock Wave Relations mathematical scaling.
LINEAR REGRESSION FIT · y = mx + b
\left(\frac{P_2}{P_1} - 1\right) = \frac{2\gamma}{\gamma + 1} (M_1^2 - 1) \implies \Delta P_{\text{ratio}} = m \cdot (M_1^2 - 1) + b
| # | Upstream Mach Parameter (dimensionless) | Pressure Jump Ratio (dimensionless) | Mach Factor (M_1^2 - 1) | Static Pressure Jump Ratio (P_2/P_1 - 1) | Remove |
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