Transcript Experimental investigations of the flow during the stage separation

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
-0.3
Model bei Mach 2,0
Platte bei Mach 2,0
-0.2
-0.1
cp
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
1.1
0
x/l(EOS)
0.1
0.2
0.3
• 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
-0.3
Model bei Mach 2,0
Platte bei Mach 2,0
-0.2
-0.1
cp
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
1.1
0
x/l(EOS)
0.1
0.2
0.3
Ma = 2.0
 = +5°
h = 40mm