Today most wind tunnels operated for aerodynamic testing are not able to reproduce flight Reynolds numbers. To overcome this problem, a shock tube flow is used to generate a high-Reynolds-number transonic flow. With this technique, Reynolds numbers up to 40 × 10 6 based on model chord length are attainable in the Mach number range between 0.6 and 0.9. Time-resolved surface pressure measurements, as well as flow visualizations, are performed for the airfoil BAC3-11/RES/30/21. Two models of different chord lengths, 80 and 100 mm, have been used. The maximum Reynolds number achieved to date is 38 x 10 6 for a freestream Mach number of 0.8. The results show that by calibrating the test-section flow carefully, this wind-tunnel technique can generate high-Reynolds-number flows, allowing aerodynamic testing at low cost. The technique avoids the complex cryogenic wind tunnel and model technique, at the expense of relatively short running times and relatively small test-section cross size. The testing time is on the order of 10-15 ms, which is sufficiently long for flow establishment and development of a useful measuring phase with stationary or at least quasi-stationary flow

Airfoil Flow Visualization and Pressure Measurements in High-Reynolds-Number Transonic Flow

The effect of the specific heat ratio gamma of the incoming ideal gas on the flow properties, especially on pressure distributions along the base and sting surfaces and on reattachment distance, was investigated. The specific heat ratios considered were gamma = 1.2, 1.4, and 1.667. Also, effects of other major parameters, such as eddy-viscosity coefficient (or effective Reynolds number) and Mach number, on the afterbody pressure and reattachment distance were studied and are discussed. Evolution of shock induced flow and stabilization time were examined and are discussed for a transient problem. The important influence of the flow-field geometry, pressure distributions, and reattachment distance on the aerodynamics radiative heat transfer for an atmosphere entry probe in high speed flight are briefly described.

Steady and Nonsteady Supersonic Turbulent Afterbody Flow

Today most of the wind tunnels used for aerodynamic testing are not able to reproduce flight Reynolds numbers. To overcome this, a shock tube flow is used to generate a high Reynolds number transonic flow. With this technique Reynolds numbers up to 40 Mio. based on model chord length are attainable in the Mach number range between 0.6 and 0.9. Time resolved surface pressure measurements as well as flow visualizations are performed for the airfoil BAC3-11/RES/30/21. Two models of different chord lengths, 80 mm and 100 mm, respectively, have been used. The maximum Reynolds number achieved so far amounts to 21 Mio. for a free stream Mach number of 0.7. The results show that, calibrating the test section flow carefully, this wind tunnel technique is suitable to generate high Reynolds number flows allowing aerodynamic testing at low costs. It avoids the complex cryogenic wind tunnel and model technique, on the expense of relatively short running times and relatively small test section cross size. The testing time is in the order of 10 to 15 milliseconds which is sufficiently long for flow establishment and an useful measuring phase with stationary or at least quasi-stationary flow.

Flow visualization and pressure measurements on an airfoil in high Reynolds number transonic flow

Two methods are investigated for alleviating wall interference effects in a shock tube test section intended for testing two-dimensional transonic airfoils. The first method involves contouring the test section walls to match approximate streamlines in the flow. Contours are matched to each airfoil tested to produce results close to those obtained in a conventional wind tunnel. Data from a previous study and the present study for two different airfoils demonstrate that useful results are obtained in a shock tube using a test section with contoured walls. The second method involves use of a fixed-geometry slotted-wall test section to provide automatic flow compensation for various airfoils. The slotted-wall test section developed exhibited the desired performance characteristics in the approximate Mach number range 0.82 to 0.89, as evidenced by good agreement obtained between shock tube and wind tunnel results for several airfoil flows.

/pdf/a-study-of-test-section-configuration-for-shock-tube-testing-3vf49313p1.pdf

A study of test section configuration for shock tube testing of transonic airfoils

Flow Visualization and Pressure Measurements on an Airfoil in High Reynolds Number

Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers of the order of 100 million can be produced, with limitations being imposed by the structural integrity of the facility or the model. A study of flow development over a simple circular arc airfoil at zero angle of attack was carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls.

https://ntrs.nasa.gov/api/citations/19790003842/downloads/19790003842.pdf

M. J. Chaney

Papers

Application of shock tubes to transonic airfoil testing at high Reynolds numbers