The tunnel noise in a Mach 5.9 Ludwieg tube is determined by two methods, a newly developed cone-probe-DNS method and a reliable hot-wire-Pitot-probe method. The new method combines pressure and heat flux measurements using a cone probe and direct numerical simulation (DNS). The modal analysis is based on transfer functions obtained by the DNS to link the measured quantities to the tunnel noise. The measurements are performed for several unit-Reynolds numbers in the range of 5 ⋅ 106 ≤ Re/m ≤ 16 ⋅ 106 and probe positions to identify the sensitivities of tunnel noise. The DNS solutions show similar response mechanisms of the cone probe to incident acoustic and entropy waves which leads to high condition numbers of the transfer matrix such that a unique relationship between response and source mechanism can be only determined by neglecting the contribution of the non-acoustic modes to the pressure and heat flux fluctuations. The results of the cone-probe-DNS method are compared to a modal analysis based on the hot-wire-Pitot-probe method which provides reliable results in the frequency range less than 50 kHz. In this low frequency range the findings of the two different mode analyses agree well. At higher frequencies, the newly developed cone-probe-DNS method is still valid. The tunnel noise is dominated by the acoustic mode, since the entropy mode is lower by one order of magnitude and the vorticity mode can be neglected. The acoustic mode is approximately 0.5% at 30 kHz and the cone-probe-DNS data illustrate the acoustic mode to decrease and to asymptotically approach 0.2%.

Although low-disturbance ("quiet") hypersonic wind tunnels are believed to provide more reliable extrapolation of boundary-layer transition behavior from ground to flight, the presently available quiet facilities are limited to Mach 6, moderate Reynolds numbers, low freestream enthalpy, and subscale models. As a result, only conventional ("noisy") wind tunnels can reproduce both Reynolds numbers and enthalpies of hypersonic flight configurations and must therefore be used for flight vehicle test and evaluation involving high Mach number, high enthalpy, and larger models. This paper outlines the recent progress and achievements in the characterization of tunnel noise that have resulted from the coordinated effort within the AVT-240 specialists group on hypersonic boundary-layer transition prediction. The new experimental measurements cover a range of conventional wind tunnels with different sizes and Mach numbers from 6 to 14 and extend the database of freestream fluctuations within the spectral range of boundary-layer instability waves over commonly tested models. New direct numerical simulation datasets elucidate the physics of noise generation inside the turbulent nozzle wall boundary layer, characterize the spatiotemporal structure of the freestream noise, and account for the propagation and transfer of the freestream disturbances to a Pitot-mounted sensor.

Since supersonic test facilities have tunnel noise that strongly influences boundary layer transition experiments, the determination of tunnel noise is of great significance to properly evaluate and interpret experimental results. The composition of tunnel noise, which consists of acoustic, entropy and vorticity modes, highly influences the boundary layer receptivity. The measurement of the separate modes is a major goal of ongoing research. In this study, the properties of stagnation point probes for a newly developed modal decomposition method for tunnel noise are investigated by direct numerical simulation. Pressure and heat flux responses of a stagnation point probe to various entropy and acoustic mode input functions are analysed to investigate how tunnel noise is perceived by corresponding sensor types. The interaction of the incident mode and the shock wave upstream of the probe is analysed and the resulting wave pattern in the subsonic region between shock wave and probe is evidenced. It is found that pure incident acoustic or entropy modes cause acoustic and entropy waves downstream of the shock wave whose strengths differ depending on the incident mode. The resulting wave pattern downstream of the shock wave is determined by postshock acoustic waves propagating bidirectionally between shock wave and probe. Formulating a model equation linking pressure and heat flux fluctuations to the initially caused postshock acoustic and entropy wave, a criterion for the applicability of stagnation point probes measuring pressure and heat flux fluctuations in the new modal decomposition method can be deduced: to distinguish between the incident mode types based on their pressure and heat flux signal the perception of initially generated entropy waves downstream of the shock wave by the heat flux sensor is crucial. The transfer function between entropy waves and heat flux is shown to have low pass filter characteristics and the cutoff Strouhal number could be estimated by control theory. The analysis of the frequency response to continuous incident waves corroborated this cutoff Strouhal number.

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Tunnel noise in supersonic testing facilities is known to be a decisive factor in boundary layer transition experiments. It defines initial conditions for the growth of modal instabilities by the receptivity mechanism. That is, to interpret experimental results, the determination of tunnel noise is of crucial importance. It is common to use stagnation point probes equipped with pressure transducers in supersonic flows, but since tunnel noise undergoes modulation during the measurement, the probes must be calibrated. The predominant component of tunnel noise is caused by the nozzle boundary layer which radiates highly inclined slow acoustic waves. Therefore, the calibration of stagnation point probes for these disturbances is essential. For quasi-steady deviations from the free stream, an analytic reduced-order method holds. A corresponding conflicting model derived by Stainback & Wagner (1972, AIAA Paper 72-1003) is revised and corrected. Inclined slow acoustic waves generate higher pressure perturbations at the probe than non-inclined waves. In general, costly three-dimensional direct numerical simulations can be used for calibration. In this study, however, new axisymmetric boundary conditions are proposed to reduce the problem to two dimensions to efficiently investigate the detection of incident inclined slow acoustic waves by stagnation point probes. A cylindrical probe with a rounded edge is investigated in supersonic flow at a Mach number $Ma=5.9$. For the inclination angle of radiated slow acoustic waves, stagnation point pressure fluctuations abruptly decay with increasing Strouhal number and a similar behaviour can be seen at constant Strouhal number with increasing inclination angle. Two simple criteria for the onset of decay based on the radial wavenumber are deduced. Furthermore, stagnation point pressure fluctuations were decomposed into an initial pulse impact and resonant amplification to separately investigate the effects. The initial pulse determines the overall pressure signal. At high inclination angles, a new mechanism for resonance caused by a surface pressure wave travelling at the phase speed of the incident wave was found to supersede resonance caused by oscillating acoustic waves prevailing at low inclination angles.