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Product Code : SCL-CELE-14085

This experimental module is dedicated to the study of supersonic aerodynamics and highly compressible flow phenomena, typically conducted within a supersonic wind tunnel. The investigation centers on analyzing pressure curves and associated losses in flow systems where the Mach number (M) exceeds unity, particularly within specialized geometries like the Laval nozzle and supersonic tunnel sections. A key objective is the visual and quantitative analysis of shock waves, which are abrupt discontinuities in flow properties that characterize supersonic flow over drag bodies. The use of advanced optical techniques, such as Schlieren optics, allows for the clear visualization of these shock waves, enabling empirical determination of the Mach number through the measurement of the shock wave angle.

Objectives: Supersonic Flow Measurement and Model Validation

Objective	Key Metric / Phenomenon	Application in Compressible Flow
Pressure Curves in Supersonic Nozzles (Laval Nozzle)	Isentropic Flow, Pressure Distribution	Analyzing expansion and acceleration in the de Laval nozzle to achieve supersonic speed.
Pressure Curves and Losses in Tunnel Flows with Mach > 1	Non-Isentropic Flow, Stagnation Pressure Loss	Quantifying the total pressure losses induced by shock waves and boundary layer effects in supersonic sections.
Observe Shock Waves in Drag Bodies using Schlieren Optics	Wave Visualization, Flow Field Imaging	Utilizing Schlieren techniques for qualitative visualization of density gradients and shock wave structures.
Determining the Mach Number from the Angle of the Shock Waves	Mach Angle mu, Oblique Shock Theory	Empirical calculation of the free-stream Mach number using trigonometric relations derived from the observed wave angle.
Comparison of Theory and Experiment	Model Validation, Predictive Accuracy	Verifying experimental data against established compressible flow theory, such as Prandtl-Meyer expansion and Rankine-Hugoniot relations.