Experimental investigations of the Experimental investigations of the flow during the stage separation of a flow during the stage separation of a

space transportation systemspace 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

Pressure Measurement

• 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 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

-0.3

-0.2

-0.1

0

0.1

0.2

0.3

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1

x/l(EOS)

cp

Model bei Mach 2,0

Platte bei Mach 2,0

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 Experimental investigations of the flow during the stage separation of a flow during the stage separation of a

space transportation systemspace 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-0.3

-0.2

-0.1

0

0.1

0.2

0.3

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1 1.1

x/l(EOS)

cp

Model bei Mach 2,0

Platte bei Mach 2,0

Ma = 2.0 = +5° h = 40mm