Experimental investigations of the flow during the stage separation of a space transportation system
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Experimental investigations of the flow during the stage separation of a space transportation system. Andrew Hay Aerospace Engineering with German. Project Brief. The ELAC 1 and EOS configuration is a two-stage-to-orbit space transportation system

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Andrew Hay Aerospace Engineering with German

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Experimental investigations of the flow during the stage separation of a space transportation system

Andrew Hay

Aerospace Engineering with German


Project Brief

  • The ELAC 1 and EOS configuration is a two-stage-to-orbit space transportation system

  • Stage separation occurs at Mach number Ma = 6.8 and at an altitude of 31 km

  • Flow visualisation - Oil flow pattern and colour Schlieren photography

  • Static wall pressure measurement

  • Identify aerodynamic interaction effects


Experimental Set-Up

  • 40cm x 40cm “Trisonic” Wind Tunnel

  • 1:150 scale EOS upper stage model and flat plate to simulate ELAC 1 lower stage

  • Test Parameters:

  • Freestream Mach number (Ma = 2.0 to 2.2)

  • Relative angle of attack (Δα = -5° to +10 °)


Test Geometry

  • Relative separation distance also planned but not possible


Flow Visualisation

  • Oil flow pattern - to visualise the near surface flow.Emulsion of oil and pigments move along wall shear stress flow lines.

  • Colour Schlieren photography - to visualise the shock system. Density gradients are made visible, because refraction index changes with density.

Pressure Measurement

  • Pressure coefficient Cp calculated from difference between static wall pressure p and ambient pressure p0.


Oil Flow Pattern

  • EOS bow shock impingement line on flat plate is visible

  • No shock induced boundary layer separation is visible

  • Reflected shock impingement line is not visible on EOS model


Colour Schlieren

  • Observed shock system very weak

  • Shock geometry used with shock theory to calculate flow conditions

  • Disturbances from flat plate very visible


Pressure Measurement

  • Shock impingement points visible (pressure increase)

  • Overall trend is a decrease in pressure downstream

  • Reason - 3D effects of the closed wind tunnel test section


Results Discussion

  • No boundary layer separation observed - confirmed by Schlieren and comparison with experimental data.

  • Shock systems very weak - shock intensities very close to 1

  • 3D effects of test section have a stronger influence on the pressure results than the shock system

  • Comparison of testing methods:All test methods consistent in providing location of shock impingement points. Schlieren is best for visualising system.


Conclusions

  • Shock systems visible, but very weak at tested Mach numbers

  • No shock induced boundary layer separation observed

  • 3D effects of the closed test section had a significant influence on the results

  • Improved test set-up is required to enable testing at more parameter variables


Experimental investigations of the flow during the stage separation of a space transportation system

Andrew Hay

Aerospace Engineering with German


Shock Theory


Shock induced BL Separation


Shock Reflection


Colour Schlierem Photo

Ma = 2.0  = +5° h = 40mm


Static Wall Pressure Measurement

Ma = 2.0  = +5° h = 40mm


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