Pressure-sensitive paint diagnostic to measure species concentration on transpiration-cooled walls

This paper presents the performance of pressure-sensitive paint (PSP) for the direct measurement of species concentration on a porous surface with mass injection. It is used to measure the ability of an injected gas to reduce the mass transfer of freestream species to the surface. A porous alumina sample was sprayed with a PSP luminophore solution. The sample was installed into a flat plate model and exposed to hypersonic cross-flows in the Oxford High-Density Tunnel. Tests were conducted with no coolant injection, air injection, and nitrogen injection at increasing blowing ratios. Oxygen partial pressure maps on the transpiration-cooled surface were obtained for several conditions at unit Reynolds numbers between $$2.58{-}5.0 \times 10^7/ \mathrm{m}$$ 2.58 - 5.0 × 10 7 / m and blowing ratios between $$0.016{-}0.078\%$$ 0.016 - 0.078 % . The oxygen pressure decreases as the unit Reynolds number decreases and the blowing ratio increases. Graphic abstract

Experimental Study of the Influence of Unit Reynolds Number on the Laminar-Turbulent Transition on the Model Swept Wing with a Subsonic Leading Edge at M = 2

Experimental studies of the influence of unit Reynolds number on the laminar-turbulent transition in a supersonic boundary layer of a swept wing with a subsonic leading edge at Mach number 2 are performed. The experiments were performed on a model of a swept wing with a swept angle of the leading edge of 72 degrees and with a 3% profile with a variable chord length in span. The hot-wire measurements showed that a laminar-turbulent transition in a supersonic boundary layer of a swept wing with a subsonic leading edge occurs earlier (~25-30%) than on a model with a supersonic leading edge with the same oncoming flow parameters. It is shown that a change unit Reynolds number insignificant influence the laminar-turbulent transition in the boundary layer of a swept wing with a subsonic leading edge.

Research on the Influence of the Unit Reynold’s Number on the Characteristics of N-Waves at M = 2.5

This paper presents the results of studying the features of the development of weak shock waves generated by a twodimensional roughness on the wall of the working part of a supersonic wind tunnel in a free flow at a Mach number of 2.5. The measurements were performed with a constant resistance thermoanemometer. It is shown that a twodimensional sticker induces weak shock waves into the free flow. They cause distortion of the average flow, the shape of which corresponds to the N-wave. High-intensity pulsations were recorded in the region of passage of a pair of weak shock waves. With an increase in the unit Reynolds number, the level of distortions of the average flow remains practically constant, but an increase in nonstationary disturbances is observed. It was found that the greatest increase in pulsations caused by Mach waves is observed in the area of the maximum gradient of the average flow. It is found that an increase in the number ReLleads to an expansion of the frequency range of unstable disturbances generated by a pair of weak shock waves.

Laminar–turbulent transition reversal on blunt ogive body of revolution at hypersonic speeds

The influence of flow parameters and nose radius on the location of laminar–turbulent transition is investigated. The model is an ogive-conical body of revolution having half angle about 9°. Experiments were conducted in a shock tunnel at Mach number 5. The transition location was diagnosed by the heat transfer rate distribution determined with the aid of luminescent temperature converters. It is shown that transition reversal can occur either (a) in the absence of turbulent wedges or (b) at a constant level of freestream disturbances. Both increasing and decreasing branches of Re∞,t (Re∞,R) dependency were observed at constant nose radius while varying only the unit Reynolds number.

Investigation of the second-mode instability at Mach 14 using calibrated schlieren

Second-mode wave growth within the hypersonic boundary layer of a slender cone is investigated experimentally using high-speed schlieren visualizations. Experiments were performed in AEDC Tunnel 9 over a range of unit Reynolds number conditions at a Mach number of approximately 14. A thin lens with a known density profile placed within the field of view enables calibration of the schlieren set-up, and the relatively high camera frame rates employed allow for the reconstruction of time-resolved pixel intensities at discrete streamwise locations. The calibration in conjunction with the reconstructed signals enables integrated spatial amplification rates ($N$ factors) to be calculated for each unit Reynolds number condition and compared to $N$ factors computed from both pressure transducer measurements and linear parabolized stability equation (PSE) solutions. Good agreement is observed between $N$ factors computed from the schlieren measurements and those computed from the PSE solutions for the most-amplified second-mode frequencies. The streamwise development of $N$ factors calculated from the schlieren measurements compares favourably to that calculated from the pressure measurements with slight variations in the $N$ factor magnitudes calculated for harmonic frequencies. Finally, a bispectral analysis is carried out to identify nonlinear phase-coupled quadratic interactions present within the boundary layer. Multiple interactions are identified and revealed to be associated with the growth of disturbances at higher harmonic frequencies.