Inviscid, nonadiabatic flow about blunt bodies

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. [...]

Shock layer temperature profiles are obtained through analysis of radiation from shock layers produced by a blunt body inserted in arc jet flow. Spectral measurements have been made in a nitrogen flow of 54.4 gm/s at an enthalpy of 8.72 MJ/kg. Vibrational temperatures for N2+ are obtained by matching spectral regions from arc jet spectra with spectra generated using the NEQAIR code. Temperature profiles obtained from the radiation layers show a vibrational temperature higher than the rotational temperature near the front of the shock and both temperatures decrease as the flow approaches the body. The spectral measurements are made and analysis completed for four distances, from the surface of the blunt body. The corresponding shock layer thickness is approximately 3.6 cm. Although the shock layer appears to be in thermal nonequilibrium, the measured rotational temperature approaches the single temperature results of viscous shock layer calculations at this test condition.

The experimental activities at Shock Wave Research Center of Tohoku University using a ballistic range are reviewed. First, the shock stand-off distance for a sphere and the shock shape over a sharp cone are measured in the hypersonic speed regime. These measurements are made for 30 mm diameter models, flight speeds between 2.5 and 4.0 km/sec and ambient pressure 4.2 and 76 mmHg. The flow field is visualized by the schlieren method using a short-pulse Nd-YAG laser as the light source. Shock stand-off distance and shock shape are shown to vary with the ambient pressure, demonstrating the effect of the chemical nonequilibrium. Secondly, hypervelocity impact tests are conducted for simulation of space debris impacts against a bumper shield. The bumpers are made of metal and nonmetal materials. The projectile made of high density polyethylene has a cylindrical shape of 14 mm diameter and 3.7 g in weight. The impact velocity is about 4 km/sec. The debris cloud is visualized by shadowgraph. The nonmetal bumper shields used in this study are more effective than the metal bumper in regard to both the bumper mass and defensive capacity. Introduction The ballistic range is one of the readily available impulse facilities for the fundamental investigation of the hypersonic flow phenomena. The so-called two stage light gas gun can accelerate a model to hypersonic speeds. The facility is used mainly for the study of hypersonic aerothermodynamics and for the hypervelocity impact tests. In Shock Wave Research Center of the Institute of Fluid Science of Tohoku University, experiments are in progress in a ballistic range with a 14 mmand 30 mm-bore two stage light gas gun. First, the shock stand-off distance for sphere and the shock shape over sharp cone are measured in the thermochemical nonequilibrium regime. In the past, a large number of experimental investigations on high temperature real-gas effects have been carried out in the shock tunnel. However, there exist uncertainties in the characteristics of the flow produced in a shock * Graduate student f Professor, Institute of Fluid Science, Tohoku University; Member AIAA. Copyright ©1998 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. tunnel in real-gas regime, and therefore it is necessary to validate the facility.Theoretically, the real-gas effects are being studied using computational fluid dynamics (CFD). However, the CFD calculation of nonequilibrium flow encounters many difficulties because of the complex processes of chemical reactions in the nonequilibrium state, particularly in the case of air. Therefore reliable experimental data are needed for validating shock tunnel results and CFD. One of the most appropriate reference quantity is the shock stand-off distance for a sphere and shock shapes for a sharp cone, because the heat absorption by the real-gas phenomena behind shock wave sensitively affects the shock stand-off distance and shock shapes.'' Ballistic ranges are the only impulse facility capable of producing a reliable data, because, in a ballistic range, a model flies in a quiescent atmosphere. Measurement of shock stand-off distance for sphere has been made by Lobb in the ballistic range (see Figure I). His experiment misses the range of the binary scaling parameter pR above 2X10 kg/m and the flight velocity below 4 km/sec. The accuracy of Lobb's experiment is unclear because of the relatively long pulse duration of the light source used for flow visualization and small size of the models. Therefore, we conduct an experiment using a Nd-YAG laser of very short pulse duration for the light source and use 30 mm-diameter models. The shock shapes have been also measured in real gas regime. However, these measurements are conducted in high enthalpy shock tunnels and expansion tubes.' As mentioned above, the data

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.

The entropy layer in a hypersonic flow over a blunted body is investigated using a high-accuracy spectral shock-fitting algorithm that solves the Euler equations within the shock layer. The base flow is computed via a direct numerical solution of the nonlinear equations. The analysis pays particular attention to the physical phenomena that arise at geometric discontinuities in body curvature. The flow field around a blunted 30-degree half-angle wedge, used as a representative body, is examined in detail and compared to viscous direct numerical solutions to evaluate the effect of viscosity on the profile of the entropy layer. Instabilities associated with the generalized inflection point in the entropy layer are investigated using linear stability theory. The entropy layer instabilities are shown to exhibit Mach number independence under a proper normalization. Our findings may be particularly useful for relating experimental results at different Mach numbers where the cold flow (calorically perfect gas) assumption is applicable.

2 Alien, R. A., Graphs: Spectrally Integrated Fluxes Including Line Contributions and Self Absorption, CR-556, Aug. 1966, NASA. 3 Gruszczynski, J. S. and Warren, W. R., Study of Equilibrium Air Total Radiation, AIAA Journal, Vol. 5, No. 3, March 1967, pp. 517-525. 4 Hunt, B. L. and Sibulkin, M., Transfer in a Gas of Uniform Properties in Local Thermodynamic Equilibrium Part 3: Detailed Calculation for Nitrogen, Nonr-562(35)/18, Dec. 1966, Brown Univ., Providence, R.L 5 Callis, L. B., Time Asymptotic Solutions of Blunt Body Stagnation Region Flows with Nongray Emission and Absorption of Radiation, TR, NASA (to be published). 6 Hoshizaki, H. and Wilson, K. H., Convective and Radiative Heat Transfer During Superorbital Entry, CR-584, Sept. 1966, NASA. 7 Chin, J. H., Effects of Non-Grey Self-Absorption and Energy Loss for Blunt-Body Flows, Journal of Quantitative Spectroscopy & Transfer, Vol. 8, 1968, pp. 503-511. 8 Thomas, M., Radiation Transfer Through High-Temperature Shock Layers, Paper 68-788, 1968, AIAA. 9 Suttles, J. T., A Method of Integral Relations Solution to the Non-Adiabatic Problem for an Inviscid BluntBody Flow Field, M.S. thesis, March 1968, Virginia Polytechnic Institute, Blacksburg, Va. 10 Suttles, J. T., private communication, May 1968. 11 Wilson, K. H., RATRAP-A Transport Code, March 1967, Lockheed Missile & Space Co. 12 Page, W. A. et al., Transport in Inviscid Nonadiabatic Stagnation-Region Shock Layers, Paper 68-784, 1968, AIAA. 13 Callis, L. B., private communication, May 1968.

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.

Along with the rapid growth of cutting-edge petrochemical technology and the pressing demand for efficiency improvement, evaluation of the performance characteristics of high-speed pump is becoming increasingly important. In this paper, numerical simulation is presented on the flow instability of a 16 straight-blade high-speed centrifugal pump with flow rate of 3 m3/h and rotating speed of 8500 rpm. Combined with the analysis of flow stability, the entropy production method is introduced to evaluate regions of high mechanical energy loss and its distribution at different flow rates. Results show that approximately 96% of the energy loss of the pump is produced in the volute, gap, and front and back chambers. Large energy loss is observed near the trailing edge of the blade and volute tongue, which are caused by the small region including both the high and low pressure gradients and large momentum exchange by the flow separation, respectively. Moreover, the rotor–stator interaction causes much energy loss at the wall of the volute and front and back chambers. Owing to the circumferential pressure gradient and the 90° leading edge of the straight blade, the fluid tends to form counter-rotating recirculation vortices. The large number of blades narrows the passage and limits the formation of large vortices in flow channels, thus the backflow phenomenon seems not to worsen with the rise of flow rates. Hence, the entropy production in most of the flow parts are insensitive to flow rates.

Energy loss and nucleon exchange in the reaction of 86Kr with 166Er

Introduction A TOMS and molecules in the hot shock layer formed in front of a hypersonic vehicle often release some of their internal energy by spontaneous emission of electromagnetic radiation. The isotropically emitted radiative energy propagates through the shock layer, sometimes undergoing absorption, until a portion of the remaining energy eventually reaches the body surface and heats the vehicle. In cases of high-speed entries of large vehicles, the radiative heating can be larger than the convective heating from the hot gas. It is therefore crucial to the design of the protective heat shield that this radiative-transport process be modeled accurately. In practice radiative transport is typically reduced to a onedimensional line-of-sight problem by assuming small body curvature and small tangential flow gradients. This approach is known as the tangent slab model. In this Note, we present a three-dimensional transport model based on geometric view factors between the emitting volumes and surface elements, which significantly improves the accuracy of radiative heat-flux estimates for an optically thin shock layer. This method can also be applied to absorbing gases, but the cost can be significantly more than the tangent slab model if fine spectral resolution is required.

The effectiveness of SiC ablation layer was investigated numerically for the radiative heating from the high-temperature shock layer to a wall surface using a very simplified shock layer with the wall-surface reflectance. Considering the entry into the atmosphere of the planet Jupiter, it was assumed that the shock layer consisted only of the chemical species originated from the atmosphere and the ablation layer from the SiC heat shield. In the calculation of the radiative heat flux with the incident angle, molecular bands, atomic lines, and continuum emission were all included. It was found that SiC ablation material was effective in reducing the radiative heat flux to the wall surface, to the same extent as carbon/carbon material.

PurposeThe purpose of this paper is to present a non‐equilibrium viscous shock layer (VSL) solution procedure that considerably improves computational efficiency, especially for long slender bodies.Design/methodology/approachThe VSL equations are solved in a shock oriented coordinate system. The method of solution is spatial marching, implicit, finite‐difference technique, which includes coupling of the normal momentum and continuity equations. In the nose region, the shock shape is specified from an algebraic expression and corrected through global passes through that region. The shock shape is computed as part of the solution beyond the nose region and requires only a single global pass. For this study, a seven‐species (O2, N2, O, N, NO, NO+, e−) air model is used.FindingsThe present approach eliminates the need for initial shock shape, which was required by previous method of solution. This method generates its own shock shape as a part of solution and the input shock shape obtained from a different solution is not required. Therefore, in comparison with the other VSL methods, the present approach dramatically reduces the CPU time of calculations. Moreover, by using the shock oriented coordinate systems the junction point problem in sphere‐cone configurations is solved.Practical implicationsThis method is an excellent tool for parametric study and preliminary design of hypersonic vehicles.Originality/valueThe present method provides a computational capability which reduces the CPU time, and expands the range of application for the prediction of hypersonic heating rates.

Many extensions of the Standard Model predict the existence of charged heavy long-lived particles, such as R-hadrons or charginos. These particles, if produced at the Large Hadron Collider, should be moving non-relativistically and are therefore identifiable through the measurement of an anomalously large specific energy loss in the ATLAS pixel detector. Measuring heavy long-lived particles through their track parameters in the vicinity of the interaction vertex provides sensitivity to metastable particles with lifetimes from 0.6 ns to 30 ns. A search for such particles with the ATLAS detector at the Large Hadron Collider is presented, based on a data sample corresponding to an integrated luminosity of 18.4 fb^{-1} of pp collisions at sqrt{s} = 8 TeV. No significant deviation from the Standard Model background expectation is observed, and lifetime-dependent upper limits on R-hadrons and chargino production are set. Gluino R-hadrons with 10 ns lifetime and masses up to 1185 GeV are excluded at 95 % confidence level, and so are charginos with 15 ns lifetime and masses up to 482 GeV.

Review: 35 refs.

In order to extend the existing approach of computing ionized flows about reentry spacecraft by the Direct Simulation Monte Carlo (DSMC) method, an ionized shock layer in a rarefied gas flow around a reentry vehicle is considered. The effect of various physical processes on the distribution of the electron temperature in the shock layer is analyzed by evaluation of the relevant relaxation and exchange terms in the equation of the electron energy balance.