Nonequilibrium shock layer temperature profiles from arc jet radiation measurements

During reentry into the atmosphere, a vehicle can encounter extreme convective and possibly radiative heating loads. Such spacecraft must rely on a carefully designed thermal protection system to maintain the integrity of the underlying structure, thus ensuring the survival of any cargo or crew it may contain. Despite many past developments and research efforts, to this day there remain large uncertainties associated with the prediction of heat shield performance. In particular, accurate modelling of ablation phenomena is critical to minimising uncertainties, reducing excess weight, and increasing safety. This includes thermochemical ablation by oxidation, nitridation, and sublimation, as well as the mechanical removal of material via spallation. Due to the high costs associated with in-flight testing, and the difficulty of mounting precision instruments and advanced diagnostics on flight vehicles, ground-based experiments stand as a cost-effective method for reducing modelling uncertainties. Ablation testing is generally conducted in long-duration facilities such as arc jets, which are well suited to investigating in-depth material response due to their ability to provide representative heating rates for extended periods. They are, however, unable to reproduce a realistic in-flight shock layer due to low Reynolds numbers and, depending on the facility type, a high degree of thermal and chemical nonequilibrium in the free-stream. In contrast, impulse facilities such as shock tubes and their variants are well suited to reproducing realistic flight conditions with free-streams that are a closer match for flows dominated by binary processes in terms of Reynolds number, temperature, and chemical composition. Due to their extremely short test times, however, ranging from the order of 10 us to several milliseconds, they are unable to aerothermally heat test models up to representative in-flight temperatures, or recreate the quasi-steady heat and mass flux balance of flight. This key limitation of impulse facilities was recently overcome in experiments using the University of Queensland’s X2 expansion tunnel by electrically pre-heating samples to temperatures approaching 2500K immediately prior to a test. In this work, the pre-heating methodology was enhanced to generate maximum surface temperatures approaching 3300K in order to study sublimation phenomena. It was found that although sublimation-regime temperatures could be achieved, the relevant surface processes themselves were suppressed due to the high pitot pressures of the conditions used. Instead, the nitridation process was studied by observing air shock layer CN emissions with pure graphite models heated to surface temperatures ranging from 1770–3300K. These results were compared with numerical simulations which predicted a monotonic increase of emissions with surface temperature for all models considered, whereas the experiments showed an initial increase from 1770–2500K and then remained relatively constant from 2500–3300K. The simulations also suggested that a significant portion of the observed CN emissions were not a result of direct nitridation, but instead due to surface CO production and subsequent gas-phase interactions to form CN. A final series of experiments was conducted using a novel model design whereby the heated graphite strip was embedded into a steel base structure in order to relieve edge effects, with the aim of producing quasi two-dimensional flow along the strip. The degree of flow two-dimensionality was found to be questionable due to visible cross-flow and a shock stand-off 10% smaller than a nominal two-dimensional case, although it was much closer than a heated strip on its own. These tests were conducted in both air and nitrogen flows to further investigate the coupling effect of surface CO production on the levels of CN emissions via gas phase reactions, as suggested by the previous comparisons with numerical simulations. Based on the relative levels of CN emissions between the two test conditions, it was concluded that although CO interactions were a non-negligible contributor to CN production, the majority was in fact due to direct surface nitridation. In terms of overall trends, the nitrogen condition results showed continually increasing CN emissions with surface temperature. For air, however, they reduced with temperature above 2500K which was consistent with the previous experiments in air. The combination of these observations suggests that the oxidation rate is decreasing with temperature above 2500K. This phenomenon is well known; however, it is the first such observation for an ablating body with the additional influence of a full realistic hypervelocity shock layer. Improved magnification in the high-speed video recordings for these experiments also allowed spallation phenomena to be observed in much greater detail than in previous work. It was clearly seen that spallation can significantly alter the flowfield, and is a phenomenon which must be given greater consideration in future studies.

The vibrational and rotational temperatures in a spark-discharge plasma were measured using optical emission spectroscopy, and the influence of the air flow velocity and ambient pressure on these temperatures was investigated. The optical emissions from the plasma were led to an imaging spectroscope through an optical fiber. The temperature was estimated by fitting a theoretically calculated spectrum to that which had been acquired experimentally, formed by nitrogen molecule emission from 372 to 382 nm. The spark-discharge plasma was examined with a flow of ambient air at a discharge energy of 80 mJ. The air flow caused the spark-discharge channel to elongate downstream. At the center of the spark plug gap, the vibrational temperature in the plasma was 4000 K, whereas the rotational temperature was 2000 K. This plasma can be regarded as being in non-thermal equilibrium because the vibrational temperature was higher than the rotational temperature. At a position approximately 3 mm downstream from the spark plug gap, the vibrational and rotational temperatures increased to 4500 and 4000 K, respectively, while approaching each other. Both temperatures reached a maximum value. These results show that the plasma transitions from non-thermal equilibrium to thermal equilibrium as it is elongated by the air flow. Ignition efficiency improvements can be expected if the time required to transition from non-thermal to thermal equilibrium can be shortened.

PLIF temperature measurements in a free-piston shock tunnel are presented. The viability of performing quantitative temperature measurements is established and application considerations are detailed. A single laser/camera system was used to measure rotational and vibrational temperatures in the freestream and in the shock layers on a wedge, a cylinder and a hemisphere. Two different laser sources were used to excite NO via the AΣ ← XΠ transition between 225 and 245 nm. The first was a Raman-shifted tunable excimer laser which was restricted to excitation of the (0,0) band near to 226 nm, and therefore was only utilised for rotational temperature measurements. The other was an excimer-pumped frequency-doubled dye laser which was used to access the first three vibrational bands (0,v=0–2) and to perform rotational and vibrational temperature measurements. Temperature measurements are compared with computations from CFD codes in an ongoing procedure of testing and validation. The calculations were performed with a quasi-one-dimensional nozzle flow code (STUBE) which has been modified to include vibrational freezing phenomena. The measured rotational temperature of 439 ± 20 K compares well with the calculated freestream rotational temperature of 396 ± 10 K. A vibrational temperature of 4500 ± 270 K was measured in the stagnation region on the hemisphere. This measurement was estimated to be systematically high by 160 K due to variations in the overlap integral between 225 and 245 nm. The temperature calculated using STUBE was 4146 K. The measured freestream and hemisphere temperatures are higher than those calculated, which questions the validity of the nozzle-reservoir temperature used in the CFD calculations and the method used to obtain it. The freestream vibrational temperature of NO was measured to be 890 ± 30K which is greater than the calculated value of 670 K. This was anticipated due to the limited vibrational relaxation data available for the calculation.

Generation and development of disturbances in a hypersonic viscous shock layer on a flat plate is studied both experimentally and numerically. The study is performed at the Mach number M∞ = 21 and the Reynolds number ReL = 1.44 × 105 and is aimed at elucidating the physical mechanisms that govern the receptivity and instability of the shock layer at extremely high hypersonic velocities. The experiments are conducted in a hypersonic nitrogen-driven wind tunnel. An electron-beam fluorescence technique, a Pitot probe and a piezoceramic transducer are used to measure the mean density and Mach number contours, as well as density and pressure fluctuations, their spectra and spatial distributions in the shock layer. Direct numerical simulations are performed by solving the Navier–Stokes equations with a high-order shock-capturing scheme in a computational domain including the leading and trailing edges of the plate, so that the bow shock wave and the wake behind the plate are also simulated. It is demonstrated that computational and experimental data characterizing the mean flow field, intensity of density fluctuations and their spatial distributions in the shock layer are in close agreement. It is found that excitation of the shock layer by external acoustic waves leads to generation of entropy–vortex disturbances with two maxima of density fluctuations: directly behind the shock wave and on the external edge of the boundary layer. At the same time, the pressure fluctuations decay inward into the shock layer, away from the shock, which agrees with the linear theory of interaction of shock waves with small perturbations. Thus, the entropy–vortex disturbances are shown to dominate in the hypersonic shock layer at very high Mach numbers, in contrast with the boundary layers at moderate hypersonic velocities where acoustic modes are most important. A parametric numerical study of wave processes in the shock layer induced by external acoustic waves is performed with variations of frequency, amplitude and angle of propagation of external disturbances. The amplitude of generated disturbances is observed to grow and decay periodically along the streamwise coordinate, and the characteristics of these variations depend on the frequency and direction of incident acoustic waves. The hypersonic shock layer excited by periodic blowing and suction near the leading edge is also investigated; in the experiments, this type of excitation is obtained by using an oblique-cut whistle. It is shown that blowing/suction generates disturbances resembling those generated by external acoustic waves, with similar spatial distributions and phase velocities. This result paves the way for active control of instability development in the shock layer by means of destructive interference of two types of disturbances. Numerical simulations are performed to show that instability waves can be significantly amplified or almost entirely suppressed, depending on the relative phase of blowing/suction and acoustic disturbances. Wind-tunnel experiments completely confirm this numerical prediction. Thus, the feasibility of delaying instability development in the hypersonic shock layer has been demonstrated for the first time.

A major design challenge for re-entry capsules lies in the modelling of convective and radiative heat transfer to the surface of the vehicle. At certain points on superorbital re-entry trajectories, up to 40% of the total radiative heat flux is contributed by the vacuum ultra-violet (VUV) spectral range and it is in this spectral range that the largest uncertainties lie. The high level of uncertainty in the VUV is a result of a lack of published experimental data due to difficulties encountered in measuring radiation in the VUV, such as strong absorption by most optical materials and air. Additional complexities of the VUV spectral range include its strongly self-absorbing nature and spectral line broadening. The primary goal of this study was to obtain calibrated spectral measurements in the VUV that enable the investigation of physical processes occurring in the shock layer that influence the incident radiative heat flux. In particular, the issues to be investigated were the variation in spectral radiance observed across a shock layer compared to the spectral radiance measured through the surface, the effects of self-absorption on spectral line intensity and the broadening of spectral lines in the VUV as a function of depth of radiating flow field. The measurements made across and through the surface of a model provide the first set of calibrated experimental results for the validation of computational codes used to predict incident radiative heat flux. Measurements made with a varying depth of radiating flow field provide a unique set of experimental data for the validation of radiation transport models and broadening coefficients. This study also used computational simulations to investigate the accuracy of a flow field solver coupled with two reaction rate schemes and compared the spectra produced using Specair with experimentally measured values. To achieve these goals, an optical system was designed to measure the VUV radiative emission produced around a blunt two-dimensional model in a spatially resolved manner across the shock layer. Spatial resolution allowed for spectral measurements to be made in both the equilibrium and non-equilibrium parts of the shock layer. A second optical system was designed to obtain measurements of VUV radiation incident on the surface of the model. This system incorporated a window in the surface with a mirror housed within the model to deflect the radiation out of the test section and into the detection system. To effectively vary the depth of the radiating flow field, the length of a two-dimensional model was varied, changing the depth of the shock layer being observed. The X2 expansion tube was used to create the high enthalpy flows required to produce radiating shock layers. Two flow conditions were created for this study that represented flight equivalent velocities of 10.0 km/s and 12.2 km/s. The spectroscopy system utilized for this study consisted of an evacuated McPherson NOVA 225 spectrometer coupled to an Andor iStar VUV enhanced intensified charge coupled device. An evacuated light tube sealed with a magnesium fluoride window was required to extend the evacuated light path to the model and avoid any absorption by molecular oxygen. An in-situ calibration of the VUV spectroscopy system was conducted using a deuterium lamp located in the position of the radiating shock layer. The integrated incident spectral radiance measured through the surface of the model between 115 nm and 180 nm was 0.744 W/cm2 sr for the 10.0 km/s condition and 12.3 W/cm2sr for the faster 12.2km/s condition. [...]

A method for determining rotational and vibrational temperatures based on the emission of the N2' First Negative band system has been extended to take into account the N2 Second Positive band system. The numerical simulation of the emission of both radiating systems and the method for the combination of the separately simulated spectra and the determination of rotational and vibrational temperatures are described. Emission spectroscopic measurements in a subsonic air plasma flow in front of a flat-faced model of 50 mm in diameter are presented. The emission within a wavelength region from 300 nm up to 500 run is used to determine rotational and vibrational temperatures for the combined N2/Nz+-system. The observed volume was defined by axial distances to the probe surface from 0 to 20 mm with a maximum spatial resolution of 0.1 mm. Radial profiles at distances of up to flO0 mm to the probe axis have been measured to preserve the required data for the Abel-inversion. The boundary layer and catalytic characteristics of the surface materials Sic and copper are characterized by the temperature profiles. The experiments were conducted in the plasma wind tunnel PWK2 of the Institut fiir Raumfahrtsysteme of Stuttgart University for different plasma states in front of a water-cooled pressure probe and a glowing Sic material probe.

: The problem considered is that of determining the inviscid, nonadiabatic flow field which results when energy loss by radiation becomes significant. The approach taken is to include the radiation term in the energy equation and solve the resulting conservation equations for the flow past a blunt body. The direct blunt-body problem is solved using an integral method of solution. The flow field is determined in an iterative manner in which the shock shape from the previous iteration is used in the current iteration. Hence, in the final analysis, no a priori information about the shock shape is required. Results have been obtained for the flow about a sphere and a hemispherecone under conditions where large energy loss from the shock layer occurs. The nonadiabatic nature of such a flow was found to significantly reduce the shock layer thickness about the body and alter the enthalpy profiles within the shock layer. As expected, the radiative heating is reduced considerably from the adiabatic value for cases where large nonadiabatic effects are present.

Hypersonic flow of a viscous gas near blunt bodies with a permeable surface is investigated for large Reynolds numbers with intense blow-in (suction) of gas from the surface, including the case when the blow-in (suction) velocity vector is oriented in a different direction from the outer normal to the surface of the body, by asymptotic methods with the help of models of three-dimensional viscous shock and three-dimensional laminar boundary layers. Analytical solutions are presented for the velocity and temperature profiles across the layer and the coefficients of friction and heat transfer on the surface, and the manner in which the basic determining parameters of the problem influence the structure of the perturbed flow and the basic behavior of the integral flow characteristics, which are important for practical applications, and the region of existence of the solution of the initial equations are analyzed.

We recently reported the first identification of rotational transitions of methyl formate (HCOOCH$_{3}$) in the second torsionally excited state toward Orion Kleinmann-Low (KL) observed with the Nobeyama 45 m telescope. In combination with the identified transitions of methyl formate in the ground state and the first torsional excited state, it was found that there is a difference in rotational temperature and vibrational temperature, where the latter is higher. In this study, high spatial resolution analysis by using Atacama Large Millimeter/Submillimeter Array (ALMA) science verification data was carried out to verify and understand this difference. Toward the Compact Ridge, two different velocity components at 7.3 and 9.1 km s$^{-1}$ were confirmed, while a single component at 7.3 km s$^{-1}$ was identified towards the Hot Core. The intensity maps in the ground, first, and second torsional excited states have quite similar distributions. Using extensive ALMA data, we determined the rotational and vibrational temperatures for the Compact Ridge and Hot Core by the conventional rotation diagram method. The rotational temperature and vibrational temperatures agree for the Hot Core and for one component of the Compact Ridge. At the 7.3 km s$^{-1}$ velocity component for the Compact Ridge, the rotational temperature was found to be higher than the vibrational temperature. This is different from what we obtained from the results by using the single-dish observation. The difference might be explained by the beam dilution effect of the single-dish data and/or the smaller number of observed transitions within the limited range of energy levels ($\leq$30 K) of $E_u$ in the previous study.

In the fundamental configuration studied here, a steady hypersonic free stream flows over a thin sharp aligned airfoil or flat plate with a leading-edge shock wave, and the flow field in the shock layer (containing a viscous and an inviscid layer) is steady laminar and two-dimensional, for a perfect gas without real and high-temperature gas effects. The viscous and inviscid layers are analysed and computed simultaneously in the region from the leading edge to the trailing edge, including the upstream-influence effect present, to determine the interactive flow throughout the shock layer and the positions of the shock wave and the boundary-layer edge, where matching is required. Further theoretical analysis of the shock layer helps to explain the computational results, including the nonlinear breakdown possible when forward marching against enhanced upstream influence, for example as the wall enthalpy increases towards its insulated value. Then the viscous layer is computed by sweeping methods, for higher values of wall enthalpies, to prevent this nonlinear breakdown for airfoils including the flat plate. Thin airfoils in hypersonic viscous flow are treated, for higher values of the wall enthalpies and with the upstream-influence effect, as are hypersonic inviscid flows, by modifying the computational methods used for the flat plate. Also, the behaviour of the upstream influence for bodies of relatively large thickness, and under wall velocity slip and enthalpy jump for flat plates, is discussed briefly from a theoretical point of view.Subsequent to the present work, computations based on the Navier–Stokes and on the parabolized Navier–Stokes equations have yielded excellent and good agreement respectively with the present predictions for large Mach and Reynolds numbers.

Shock layer thickness and optimum linear velocity in displacement chromatography

A mid-infrared interference-free laser absorption technique for simultaneously measuring rotational temperature, vibrational temperature, and CO concentration was developed for application to shock-tube studies on thermochemical non-equilibrium CO over 1000–3000 K. Three transition lines in the fundamental vibrational band of CO (P(0, 21), near 4.87 μm, P(1, 21), near 4.93 μm, and P(0, 37), near 5.05 μm) were selected. The P(0, 21)/P(1, 21) line pair was used for vibrational temperature measurements whereas the P(0, 21)/P(0, 37) line pair was used for rotational temperature measurements. Spectroscopic parameters for developing the technique were measured: line strengths and collisional broadening data in Ar were obtained at 1040–2940 K. Validation experiments for the thermometry system were performed in shock-heated thermal-equilibrium CO/Ar mixtures at 1050–3010 K and 1.1–2.8 bar. The time-dependent rotational and vibrational temperatures were measured during the vibrational relaxation processes of CO. The technique showed high sensitivity in detecting the rotational and vibrational temperatures. The measured rotational temperature agreed well with the temperature calculated using the measured pressure and isentropic relationship. The measured vibrational temperature showed good agreement with the predictions using the Landau and Teller theory and Millikan and White relationship. The time-dependent CO concentration during the oxidation processes of n-heptane over a wide temperature range (1350–2750 K) was measured considering n-heptane as one of the alternative fuels for the scramjet. The interference-free laser absorption strategy showed good flexibility in detecting the CO concentration at ultra-high temperatures. The measured results showed overall good agreement with the predictions from two detailed mechanisms and one skeletal mechanism. The reactivity of n-heptane was found to be insensitive to the temperature increase at ultra-high temperatures (>2100 K).

A numerical investigation is carried out within the framework of the multicomponent total viscous shock layer model [6, 7], according to which when Re∞ ≳ 100 the flow near a blunt body can be divided into a shock wave zone and a viscous shock layer. At the inner edge of the shock wave the generalized Rankine-Hugoniot relations are imposed, and in the shock layer the complete system of viscous shock layer equations is solved with allowance for vibrational relaxation and nonequilibrium dissociation and ionization reactions.

A reentry vehicle is exposed to a partially ionized flow during the reentry flight. For such a flight, a strong magnet mounted on the vehicle, which generates the magnetic field around the vehicle, is suggested to affect a surrounding ionized flow and make it possible to control the flow. Such an electromagnetic effect on the flow is investigated experimentally by using a small arc-jet wind tunnel. In the experiment, the translational temperature distribution in the shock layer around a magnetized blunt body in a supersonic, weakly ionized, argon flow is determined by applying an absorption spectroscopic technique. For the absorption spectrum affected by the magnetic field, the temperature determination method was newly developed. The temperature distribution thus determined for the shock layer shows that the applied magnetic field significantly affects the shock layer or, specifically, the shock standoff distance and enhances it.

In the framework of the approximation of local similarity to the Navier-Stokes equations, an investigation is made of the axisymmetric flow of homogeneous gas in a hypersonic shock layer, this including the region of transition through the shock wave. Boundary conditions, which take into account blowing of gas, are specified on the surface of the body and in the undisturbed flow. A numerical solution to the problem is obtained in a wide range of variation of the Reynolds number and the blowing parameter. Expressions are found for the dependences on the blowing parameter usually employed in boundary layer theory of the coefficients of friction and heat transfer on the surface of the body, which are divided by their values obtained for blowing parameter equal to zero. It is shown that these dependences are universal and the same as the dependences obtained from the solution of the equations of a hypersonic viscous shock layer with modified Rankin-Hugoniot relations across the shock wave and from the solution of the boundary layer equations.