The application of coherent anti-Stokes Raman scattering to temperature measurements in a pulsed high enthalpy supersonic flow

Broadband single pulse coherent anti-Stokes Raman scattering (CARS) experiments employing a folded box phase matching geometry in a shock tunnel flow are presented. Rovibrational spectra of molecular nitrogen, produced at the exit of a pulsed supersonic nozzle for a range of flow enthalpies, are examined. Difficulties peculiar to the application of the optical technique to a high enthalpy pulsed flow facility are discussed and measurements of flow temperatures are presented. Theoretically calculated values for temperatures based upon algorithms used to determine shock tunnel flow conditions agree well with experimental measurements using the CARS technique.     

Ultrafast vibrational spectroscopy imaging of nanoshock planar propagation

Variations of the velocity and pressure of repetitive 4 GPa laser-driven shock fronts in a polymer thin film are studied using spatially resolved ultrafast coherent anti-Stokes Raman Spectroscopy (CARS) measurements of shock compression of an anthracene optical gauge. Even though the radial profile of the laser pulse that generates the shock is a Gaussian that falls off steeply at the edges, optical saturation effects in the shock generation layer flatten out the shock front. Detailed measurements show the shock pressure, as indicated by the blueshift of an anthracene vibrational transition, and the shock velocity, as indicated by the arrival time at the anthracene gauge layer, remain constant within better than 5% over the central region probed by CARS, over a run distance of at least m. Received 30 July 2001 / Accepted 13 March 2002 Published online 17 June 2002     

Aerodynamic calibration of TCM2 facility and study of a bow shock layer by emission and laser spectroscopy

Flow properties in the TCM2 free piston shock tube/tunnel are determined by time-resolved pressure and heat flux measurements in numerous points of the shock tube and the nozzle, and in the free flow for two stagnation enthalpy conditions (3.5 and 11 MJ/kg). These measurements demonstrate the homogeneity of the flow during more than 1 ms. The cleanness of the useful test time is shown with time-resolved emission measurements at critical wavelengths. NO fluorescence profiles are established with local and planar laser-induced fluorescence in the shock layer around a cylindrical model. It allows to determine the shock stand-off distance for both enthalpy conditions. The problems of quenching and amplified spontaneous emission are considered. The importance of atomic oxygen and atomic nitrogen densities as well as temperature effects is also shown. Evaluation of the temperatures behind the shock front through spectroscopic data agrees with calculations. The proof of the presence of vibrationally excited NO ahead of the shock layer is given. Received 14 March 2000 / Accepted 18 June 2001     

Simulation of the Effect of the Freestream Turbulence Parameters on Heat Transfer in an Unsteady Boundary Layer

A variant of the two-parameter turbulence model which makes it possible continuously to calculate a flow region with laminar, transition and turbulent regimes is proposed for investigating the flow under conditions of high freestream turbulence intensity. It is shown that the properties of the thermal transition can be theoretically described using the quasi-steady turbulence model in the case of periodic freestream velocity distribution. The numerical results are compared with theoretical and experimental data. The approach proposed is developed for determining the combined effect of the parameters of harmonic fluctuations of the external velocity and freestream turbulence on the heat transfer characteristics on a flat plate with different boundary conditions for the enthalpy.     

Comparison of velocity and temperature measurements with simulations in a hypersonic wake flow

A hypersonic shock-tunnel flow around an axisymmetric model of a planetary entry probe is analyzed. Planar laser-induced fluorescence is applied to measure both the velocity and the rotational temperature everywhere in the central plane of the flow field. The experimental test case is compared to simulations using the direct simulation Monte Carlo (DSMC) method. While the Mach 9.7 flow at a freestream Reynolds number based on the model diameter of 35,000 is chemically frozen, effects of thermal non-equilibrium and localized rarefaction cannot be neglected. DSMC and measurements agree well within the outer wake, but disagree close to the centerline, where in particular the measured velocity is higher than values predicted by the simulations. The experimental results indicated a shorter recirculation region and increased local fluctuations in the free shear layer upstream of the wake recompression shock when compared to the simulation. These effects are attributed to incipient transition, which is not observed in the simulations, as the simulations did not model the effects of freestream fluctuations. Furthermore, measured and simulated vorticities are compared with theoretical predictions.     

高速双锥绕流中热化学与输运模型影响研究

激波与边界层之间相互作用是高超声速飞行中的常见现象,对飞行器气动性能与飞行安全至关重要.对于高焓来流,流场中通常存在复杂的物理化学现象,此时准确模拟流场中激波边界层相互作用的难度大,相关物理化学建模仍有待进一步考察和研究.本文针对最近文献中纯净空气高超声速双锥绕流实验开展数值研究,分别研究了不同热化学模型与输运模型对壁面压力与热流的影响.热力学模型包括完全气体、热力学平衡和非平衡模型,化学模型包括冻结和非平衡化学模型,输运模型包括经典的Wilke/Blottner/Eucken模型与更加复杂的Gupta/SCEBD模型,以及考虑壁面催化/非催化影响的模型.计算了6个不同算例,涵盖了低焓至高焓来流等不同工况.壁面压力与热流的数值计算结果与实验结果符合较好;对于低焓来流,计算结果主要受到分子内能分布的影响,输运模型对计算结果的影响不大;对于高焓来流,一方面计算结果受到化学反应与壁面催化的影响较大,另一方面不同输运模型对计算结果的影响也更加明显.      

Experimental and Numerical Investigations of the Characteristics of Pulsed Thermal Actuators

The purpose of the study is to numerically, experimentally, and analytically investigate the characteristics of plasma pulsed thermal actuators (PT actuators) and to assess their possibilities in controlling flow around airfoils, wings, and configurations at large subsonic freestream velocities. For the PT actuators of the types considered the mathematical models adequately describing their effect on flow past bodies are developed. The characteristics of a prototype PT actuator are experimentally investigated on a specially developed rig. A new type of the PT actuator equippedwith a channel (PTC actuator) is proposed; it is designed to operate at a high pulse repetition frequency and at large flow velocities. Numerical investigations show that the PTC actuators are free of the essential and fundamental shortcoming of the PT actuators which consists in the working zone superheating at high pulse repetition frequencies.     

Nature of the surface heart transfer fluctuation in a hypersonic separated turbulent flow

This paper presents the results of an experimental study of the unsteady nature of a hypersonic separated turbulent flow. The nomimal test conditions were a freestream Mach number of 7.8 and a unit Reynolds number of 3.5×10⁷/m. The separated flow was generated using finite span forward facing steps. An array of flush mounted high spatial resolution and fast response platinum film resistance thermometers was used to make multi-channel measurements of the fluctuating surface heat trtansfer within the separated flow. Conditional sampling analysis of the signals shows that the root of separation shock wave consists of a series of compression wave extending over a streamwise length about one half of the incoming boundary layer thickness. The compression waves converge into a single leading shock beyond the boundary layer. The shock structure is unsteady and undergoes large-scale motion in the streamwise direction. The length scale of the motion is about 22 percent of the upstream influence length of the separation shock wave. There exists a wide band of frequency of oscillations of the shock system. Most of the frequencies are in the range of 1–3 kHz. The heat transfer fluctuates intermittently between the undisturbed level and the disturbed level within the range of motion of the separation shock wave. This intermittent phenomenon is considered as the consequence of the large-scale shock system oscillations. Downstream of the range of shock wave motion there is a separated region where the flow experiences continuous compression and no intermittency phenomenon is observed. The project supported by National Natural Science Foundation of China     

On the influence of elevated surface temperatures on hypersonic shock wave/boundary layer interaction at a heated ramp model

Although important flow parameters as Mach number, Reynolds number and total enthalpy can be reproduced in most hypersonic experiments quite well, due to different surface temperature effects in wind tunnel and flight, scaling as well as specific flow properties of shock wave/boundary layer interactions are different. This especially holds for short-duration facilities like, e.g. shock tunnels where due to short running times the models remain more or less at ambient temperature. To overcome this shortcoming, an experimental study has been conducted using a preheatable ramp model with 15° ramp angle. This allowed us to adjust the surfaces to an arbitrary temperature just before the experiment started. Pressure and heat flux measurements clearly showed the effect of varying surface and free stream temperatures. These results are supported by schlieren pictures and infrared measurements. The comparison of the measurements with theoretical and numerical results shows a good agreement. Separation bubble scaling laws proposed by Katzer and Davis have been applied and partially confirmed using the local conditions of the boundary layer at separation.     

Near-wall imaging using toluene-based planar laser-induced fluorescence in shock tube flow

A quantitative thermometry technique, based on planar laser-induced fluorescence (PLIF), was applied to image temperature fields immediately next to walls in shock tube flows. Two types of near-wall flows were considered: the side wall thermal boundary layer behind an incident shock wave, and the end wall thermal layer behind a reflected shock wave. These thin layers are imaged with high spatial resolution (15μm/pixel) in conjunction with fused silica walls and near-UV bandpass filters to accurately measure fluorescence signal levels with minimal interferences from scatter and reflection at the wall surface. Nitrogen, hydrogen or argon gas were premixed with 1–12% toluene, the LIF tracer, and tested under various shock flow conditions. The measured pressures and temperatures ranged between 0.01 and 0.8 bar and 293 and 600 K, respectively. Temperature field measurements were found to be in good agreement with theoretical values calculated using 2-D laminar boundary layer and 1-D heat diffusion equations, respectively. In addition, PLIF images were taken at various time delays behind incident and reflected shock waves to observe the development of the side wall and end wall layers, respectively. The demonstrated diagnostic strategy can be used to accurately measure temperature to about 60 μm from the wall.     

Mass spectrometric measurements of the freestream composition in the T4 free-piston shock-tunnel

The freestream composition is an important parameter in ground-based aerodynamic testing, and direct measurement of it is very important. This paper reports extensive composition measurements in the freestream of the T4 free-piston shock-tunnel, employing a recently improved time-of-flight mass spectrometer. A wide range of nozzle reservoir conditions were used. The results show good agreement between measured and theoretical values for nitric oxide over the entire enthalpy range reported (2.5–13 MJ/kg). This provides confidence that the chemistry model is correctly predicting sudden freezing of NO in the nozzle expansion. On the other hand, no monatomic species have been measured other than those produced by dissociative ionisation within the mass spectrometer, even at flow conditions where significant freestream dissociation is expected. Furthermore, excess diatomic oxygen is detected at high enthalpies. These observations are consistent with the possibility that oxygen recombination is not correctly predicted in the nozzle expansion, with sudden freezing occurring significantly later than predicted. However, the observations are also consistent with possible catalytic recombination in the skimmer system. The possibility for producing an empirical correlation between the freestream composition and the reservoir entropy has also been observed. PACS 07.75.th; 47.40.-x     

Superorbital expansion tube tests of a caret waverider

A computational and experimental study was conducted to assess the potential of testing waverider configurations in a high-performance, short-duration expansion tube facility. The tests were performed in the newly commissioned X3 superorbital expansion tube and provide the first experimental data of a waverider tested at a stagnation enthalpy and equivalent flight speed exceeding 40 MJ/kg and 9 km/s, respectively. Two simple caret configurations were chosen as benchmark test cases to test the use of the facility, instrumentation and numerical models to investigate these flows. The general performance of the sharp and blunt leading edge waveriders at angles of attack ranging from 0° to 5° were analyzed and compared to CFD and theoretical predictions. For the conditions tested, the presence of a strong viscous interaction caused the shock wave to be detached from the leading edge of the models resulting in a significant loss in performance. An analytical model was developed to account for the strong coupling between the shock wave and boundary layer. Results were shown to be in very good agreement with CFD estimates for both configurations at all angles of attack considered. Finite-rate chemistry CFD simulations indicated that real gas effects other than the residual levels of nonequilibrium freestream dissociation present in the expansion tube flow were negligible for the conditions tested. The study also revealed that a past flow visualization technique gave a false indication of the leading edge shock location. An improved experimental visualization technique was successfully tested with results from these tests correlating well with computational estimates. This study successfully demonstrated the use of the facility to study waverider performance at speeds representative of orbital flight.      

Supersonic Flow Past a Thermal Source

The results of an experimental investigation and numerical simulation of a gasdynamic structure formed as a result of supersonic flow past a pulsating thermal source are presented. Heat was supplied to the flow by producing a limited plasma volume as a result of the breakdown of the focused radiation of a CO₂ laser operating in the pulse periodic regime. On the basis of the experimental data obtained, a thermal source model was developed and accepted for further numerical calculations. The calculations were carried out within the framework of the inviscid gas model using the TVD scheme and nonreflecting boundary conditions. The effect of the relevant gasdynamic and energetic parameters on the flow pattern associated with the studied phenomenon is established. Data on the flow parameter distributions in the wake of the thermal source are obtained as a function of the freestream Mach number.     

PLIF thermometry in shock tunnel flows using a Raman-shifted tunable excimer laser

Planar laser-induced fluorescence is performed in a free-piston shock tunnel by using a Raman-shifted tunable excimer laser to excite nitric oxide molecules in the flow. Two different flowfields are examined to test the difficulties associated with applying the technique to shock tunnels: the bluff body flow produced by a 25 mm diameter cylinder; and the oblique shock and expansion fan produced by a 35° half-angle wedge. For the cylinder, the maximum flow enthalpy was limited to 4.1 MJ kg due to high flow luminosity which is produced by metallic contaminants in the flow. A reflective filter is used to reduce the influence of flow luminosity making these measurements feasible. Freestream temperature measurements are in excellent agreement with those predicted from numerical flow calculations. Large uncertainties were observed for the high-temperature post-shock results. Several higher enthalpy shots (14 MJ kg) were also performed with the wedge and showed an insignificant amount of contaminant emission. Received 5 June 1996 / Accepted 8 February 1998     

High enthalpy nozzle flow sensitivity study and effects on heat transfer

The sensitivity of the flow along the nozzle and in the test section of high enthalpy wind tunnels to the thermochemical response of the nozzle expansion process, as well as effects on the pressure and heat transfer distributions over the Electre blunt cone standard test model, are examined in the framework of properly characterizing the test section flow field in such facilities. Particularly sensitive to the thermochemical behaviour of the nozzle flow, in the facilities under consideration, are the static pressure, static temperature and Mach number, whereas stagnation point (pitot) pressure and heat transfer data or freestream velocity are inadequate for the characterization of the thermochemical state of the flow. The Electre and nozzle wall pressure data in the F4 arc jet wind tunnel suggest, in contrast to nonequilibrium computations, that the flow in the F4 nozzle is close to equilibrium. In the HEG and, to some extent, the T5 piston-driven shock tunnels, there are indications that significant heat losses occur in the reservoir. Lastly, simple semi-empirical formulations for stagnation point heating are shown to perform reasonably well in high enthalpy flow conditions.     

Experiments in hand-operated,hypersonic shock tunnel facility

Experiments were conducted using the newly developed table-top, hand-operated hypersonic shock tunnel, otherwise known as the Reddy hypersonic shock tunnel. This novel instrument uses only manual force to generate the shock wave in the shock tube, and is designed to generate a freestream flow of Mach 6.5 in the test section. The flow was characterized using stagnation point pressure measurements made using fast-acting piezoelectric transducers. Schlieren visualization was also carried out to capture the bow shock in front of a hemispherical body placed in the flow. Freestream Mach numbers estimated at various points in the test section showed that for a minimum diameter of 46 mm within the test section, the value did not vary by more than 3 % along any cross-sectional plane. The results of the experiments presented here indicate that the device may be successfully employed for basic hypersonic research activities at the university level.     

Heating enhancement caused by a transverse control jet in hypersonic flow

Experiments are reported in which the heat flux distribution near a single circular, sonic transverse jet on a flat plate exposed to a hypersonic (Mach 6.7) freestream flow was quantitatively measured using thermochromic liquid crystals. The freestream conditions were such that the boundary layer growth on the plate ahead of the jet was laminar. The results indicate that the interaction of the jet with the freestream flow created a complex flowfield with regions of separation and reattachment which caused localised enhancements to the heat flux upstream and to the side of the jet, the magnitudes of which were sensitive to both jet plenum pressure and jet gas composition. Received 28 August 1996 / Accepted 6 June 1997