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This paper presents the analysis of optical emission spectroscopic measurements from vacuum ultraviolet (120 nm) to near infrared (960 nm) in a high-enthalpy air plasma flow corresponding to superorbital reentry conditions. The vacuum ultraviolet measurements have been realized with a new experimental setup allowing measurements through a bore hole in the sample. Using the commonly applied optical emission spectroscopy from the side, the radiation transport along the line of sight of the vacuum ultraviolet measurements has been assessed. This allowed the determination of the local ground state densities of atomic oxygen and atomic nitrogen from a branching ratio analysis and the blackbody limiting correction of the vacuum ultraviolet radiation. Through the analysis of the absorption in the borehole, the spectra have been corrected, resulting in a stagnation point radiative heat flux in the vacuum ultraviolet of 235 kW∕m 2 .= local mass-specific enthalpy, MJ∕kg I = arc current, A K = proportionality constant for air, W∕MJ∕kg·Pa 0.5 ·m 1.5 k = Boltzmann constant, J∕K k cal = calibration factor, W∕m 2 · nm · sr∕counts∕s L λ = spectral radiance, W∕m 2 · nm · sr L λ;Limit = blackbody limit spectral radiance, W∕m 2 · nm · sr M = magnification _ m = mass flow, g∕s n = number density, m −3 n i = excited state population density, m −3 n 0 = ground state population density, m −3 P = electric power, kW P k = spectral line shape, nm −1 p = pressure, hPa _ q = heat flux, kW∕m 2 R eff = effective nose radius, mm r = radial distance, mm S = raw signal, counts/s s = distance from the probe surface, mm s 0 = distance between object and focusing element, mm= vibrational quantum number x = axial distance from the nozzle, mm y = horizontal distance from the nozzle, mm z = vertical distance from the nozzle, mm α λ = absorption coefficient, m −1 δ = full width at half-maximum of line broadening, nm ε λ = emission coefficient, W∕m 3 · nm · sr ε = spectrally integrated emission coefficient, W∕m 3 · sr λ = wavelength, nm σ λ = absorption cross section, m 2 Subscripts BB = blackbody DL = deuterium lamp exc = electronic excitation IS = integrating sphere l = lower state rot = rotational tot = total trans = translational u = upper state vib = vibrational

Three different types of rocks were tested in a high enthalpy air plasma flow. Two terrestrial rocks, basalt and argillite, and an ordinary chondrite, with a 10 mm diameter cylindrical shape were tested in order to observe decomposition, potential fragmentation, and spectral signature. The goal was to simulate meteoroid ablation to interpret meteor observation and compare these observations with ground based measurements. The test flow with a local mass-specific enthalpy of 70 MJ kg −1 results in a surface heat flux at the meteorite fragment surface of approximately 16 MW m −2 . The stagnation pressure is 24 hPa, which corresponds to a flight condition in the upper atmosphere around 80 km assuming an entry velocity of 10 km s −1 . Five different diagnostic methods were applied simultaneously to characterize the meteorite fragmentation and destruction in the ground test: short exposure photography, regular video, high-speed imaging with 10 kHz frame rate, thermography, and Echelle emission spectroscopy. This is the first time that comprehensive testing of various meteorite fragments under the same flow condition was conducted. The data sets indeed show typical meteorite ablation behavior. The cylindrically shaped fragments melt and evaporate within about 4 s. The spectral data allow the identification of the material from the spectra which is of particular importance for future spectroscopic meteor observations. For the tested ordinary chondrite sample a comparison to an observed meteor spectra shows good agreement. The present data show that this testing methodology reproduces the ablation phenomena of meteoritic material alongside the corresponding spectral signatures.

This paper experimentally examines the internal and external flow characteristics of porous zirconium diboride (Z r B 2 ), an Ultra-High-Temperature-Ceramic (UHTC) and a potential candidate for transpiration cooling of hypersonic vehicles. This is performed for both partially sintered material and fully densified material with cast features. The Darcy and Forchheimer permeability coefficients of these samples are determined using an ISO standard test rig. The outflow of the transpiring porous samples is investigated where no hypersonic cross-flow is involved using hot-wire anemometry and focused Schlieren visualisation. The velocity maps obtained from the hot-wire data show significant non-uniformities across the UHTC's outflow region, both at low and high differential pressures. The focused Schlieren using carbon dioxide as the injected gas reveals unsteady structures at high differential pressures as the outflowing gas interacts with the surrounding air.

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The mixing between the coolant and the boundary-layer gas downstream of an injector—for transpiration/film cooling—has been extensively studied for turbulent flows; however, only a handful of studies concerning laminar mixing exist, particularly in hypersonic flows. In this paper, the concentration of the coolant gas at the wall and the heat flux reduction downstream of a transpiring injector in a hypersonic laminar flow are experimentally measured and examined. Experiments are performed in the Oxford High Density Tunnel at Mach 7. A flat-plate model is coated with pressure-sensitive paint (PSP) to spatially resolve the film and obtain a film effectiveness based on coolant concentration. Thin-film arrays are installed to measure the heat flux reduction. Six different cases are studied featuring nitrogen and helium as the coolant gas, where the blowing ratio is varied from 0.0406% to $$0.295\%$$ 0.295 % . The unit Reynolds number of the flow is $$12.9\times 10^6\;\mathrm {m^{-1}}$$ 12.9 × 10 6 m - 1 . A coolant concentration of up to $$95\%$$ 95 % is achieved immediately (2 mm) downstream of the injector. The film concentration drops in a monotonic fashion farther downstream; however, a constant film coverage of 5–20 mm immediately downstream of the injector is observed in cases with a higher blowing ratio. A film coverage above 15% over three injector lengths is present even for the lowest blowing ratio. Heat flux reduction is achieved in all cases. The concentration effectiveness obtained from PSP is compared with the thermal film effectiveness calculated from the heat flux reduction. The latter is found to be higher than the former for all data points. Finally, a collapse of the thermal effectiveness is achieved and a modified analytical correlation is proposed. Graphical abstract

This article presents the full operational experimental capabilities of the plasma wind tunnel facilities at the Institute of Space Systems at the University of Stuttgart. The simulation of the aerothermodynamic environment experienced by vehicles entering the atmosphere of Earth is attempted using three different facilities. Utilizing the three different facilities, the recent improvements enable a unique range of flow conditions in relation to other known facilities. Recent performance optimisations are highlighted in this article. Based on the experimental conditions demonstrated a corresponding flight scenario is derived using a ground-to-flight extrapolation approach based on local mass-specific enthalpy, total pressure and boundary layer edge velocity gradient. This shows that the three facilities cover the challenging parts of the aerothermodynamics along the entry trajectory from Low Earth Orbit. Furthermore, the more challenging conditions arising during interplanetary return at altitudes above 70 km are as well covered.

This paper presents results of a system study of transpiration cooled thermal protection systems for Earth re-entry. The cooling performance for sustained hypersonic flight and transient re-entry of a blunt cone geometry is assessed. A simplified numerical model is used to calculate the transient temperature of a transpiration cooled heat shield. The performance of transpiration cooling is assessed by calculating the overall required coolant mass for different steady state and transient flight scenarios. Spatially and temporally optimised mass injection is presented for various flight conditions. The majority of the injection is required on the spherical nose segment of the blunted cone. Carbon/Carbon composite ceramic and the ultra high temperature ceramic Zirconium diboride are considered as wall materials. Both materials require similar amounts of coolant injection. In continuous hypersonic cruise, transpiration cooling is highly effective for flight conditions with velocities below 8 km s −1 and altitudes above 40 km. For transient re-entry, transpiration cooling is most viable for trajectories of entry velocities below 8.5 km s −1 and ballistic coefficients below 2.1 kg m −2. Nomenclature A Area, m 2 C D Drag coefficient c p Specific heat capacity, J kg −1 K −1