Wind Tunnels for Supersonic Laminar Flow
ALT and CFI arc not very sensitive to disturbances of the incoming flow or noise radiation (table 3). So. classical wind tunnels should be suitable for investigations. But at the high sweep angles of subsonic leading edges massive suction is required (at high model Reynolds numbers up to 300 Mio). Suction flow cannot be simulated in the wind tunnel, at least not for complete aircraft models: The hole diameter in the suction surfaces is at the limits of manufacture, hole diameter cannot be reduced according to model scale; this violates the model laws. This violation becomes important, when the hole diameter is not small compared to the local boundary layer thickness. For SCT-applications, the hole diameter at the leading edge of the flying aircraft is about the boundary layer thickness. Suction simulation on aircraft models is therefore impossible, at least in the vicinity of the leading edge.
Table 3: Wind Tunnel Simulation
TSI are very sensitive to external disturbances. For supersonic wind tunnels these arc the turbulence of the incoming flow (as for subsonic wind tunnels). Also, strong noise is radiated into the test section. It is – by one part – produced by upstream noise radiated via the reservoir section, e g. valve noise in blow down tunnels But the most severe pan is boundary layer noise radiated by the turbulent boundary layer of the wind tunnel nozzle into the lest section: Each turbulent eddy in the outer boundary layer produces a small shock wave on its back which radiates a strong noise in Mach line direction. This provokes premature transition, so that effectively transition in supersonic wind tunnels scams to be dependent on nozzle Reynolds number instead of the model Reynolds number, the so called unit Reynolds number effect.
Furthermore, in most supersonic wind tunnels the attainable Reynolds numbers are completely insufficient. Often they arc so low, that after provoked transition (tripping) relami – narisation occurs (389). The cruise Reynolds numbers for supersonic transpons arc about Ret = 300 Millions with respect to the aircraft length L!
In the past, surface temperature of supersonic wind tunnel models was not taken into account For investigation of TSI (and HMI). accurate simulation of the temperature profile in
the boundary layer is necessary, i. e. the ratio of model wall temperature to stagnation temperature must be simulated.
To enable supersonic transition measurements, a quiet supersonic wind tunnel was developed at NASA-Langlcy (Figure 111) (390]. It is a small pilot tunnel for Mach 3.5:
In the subsonic part of the nozzle throat the boundary layer is removed to provide a young laminar boundary layer in the wind tunnel nozzle. This laminar boundary layer does not radiate significant noise into the test section. When the nozzle boundary layer becomes turbulent, noise is radiated. But in supersonic flow this noise follows characteristics (Mach lines). So a quiet test zone is provided, beginning with the parallel flow section and ending with the characteristics of the nozzle transition zone. This wind tunnel provided transition measurement results comparable to flight tests.
Figure 111 Quiet Supersonic Wind Tunnel
Another wind tunnel provided test data not showing the unit Re-cfTect: this was the Ludwieg tube in Gdtlingcn (391J with measurements at Mach 5 (Figure 112) (392). The reason for these high quality measurements is not completely understood: The Ludwieg tube provides an incoming flow of extremely low turbulence, but the wind tunnel nozzle has a turbulent boundary layer of high nozzle Reynolds number, i. e. with small boundary layer thickness and – perhaps – not so disturbing noise levels and spectra.
no unit Ri ■ «*Чсі round kx Т» CMг««п. л).*; to* и » Si
Figure 112 Ludwieg Tube
Wilh respect to future supersonic laminar flow experiments, the Ludwieg principle should be considered in Europe, possibly with a quiet nozzle – like in the US. where Ludwieg tubes are designed for SCT-tests. Some advantages arc obvious:
• superb flow quality,
• high Reynolds numbers possible (about 300 Mio),
• quiet nozzle design easier at short testing times,
• model cooling easier for short testing times.
• low costs, affordable in Europe,
• but low productivity (about 0.5 s run time).
For the latter this test facility will not be suited for standard development tests, but rather for high performance quality checks. A suited Ludwieg tube (today Mach numbers 1.75 and 2.5; size more than 1 m diameter) exists at the University of Stuttgart and was refurbished in the past years.
Considering these facts, a new procedure for aircraft design must be developed. It relies more on theoretical predictions, partial simulation tests in classical wind tunnels (i. c. model tests with turbulent flow) and very carefully selected checks in the high quality Ludwieg tube, mostly to evaluate transition physics in validation experiments. Comparable procedures were developed for reentry vehicles like Hermes: Partial simulation is used, a relatively small number of experimental checks, and it relics strongly on theoretical predictions.