Hypersonic Shock Tunnel Research Articles

Nozzle Flow Characterization of the SNU Hypersonic Shock Tunnel

AbstractThis paper presents the operational capabilities and the characteristics of the nozzle flow of the Seoul National University Hypersonic Shock Tunnel (SHyST), a recently added impulse facility designed to produce high-enthalpy flows reaching up to 5 MJ/kg through hypersonic contoured nozzles. Within this investigation, emphasis is placed on the design and utilization of a pitot rake, serving as an essential instrument for quantifying pitot pressure and Mach number distributions throughout the test section. The experimental results confirm the presence of axis-symmetric behavior and spatial–temporal uniformity of the freestream Mach number at the nozzle outlet and along the test section. In addition, the study demonstrates a good agreement with the predictions of numerical simulation. Mach number validation is further supported by the measurement of shock standoff distance through schlieren visualization.

Carbon-Monoxide Laser Absorption Spectroscopy Measurements over Ablating Graphite in a Shock Tunnel

Measurements of carbon monoxide (CO) during air–carbon ablation of graphite are performed in the Sandia Hypersonic Shock Tunnel. Resistive heating is used to bring graphite samples to wall temperatures (Tw=1250–1630 K) as measured by an imaging pyrometer. The heated models are subjected to a flow condition of H0=10.6 MJ/kg, Tr(∞)=630 K, P∞=300 Pa, and u∞=4240 m/s. Tunable diode laser absorption spectroscopy measures CO temperature and concentration within the boundary layer at 50 kHz. Ablation product concentrations from experiments are presented as a function of time, including the transient startup of the tunnel. The measured CO concentration increases by approximately a factor of 2 as surface temperatures increase from 1250 to 1630 K. Finally, the CO concentration is observed to be lower than that predicted by several air–carbon ablation models, with the best agreement occurring at the higher surface temperature.

On the determination of the standing oblique detonation wave in an engine combustor using laser absorption spectroscopy of hydroxyl radical

Freejet experiments were conducted in the hypersonic flight-duplicated shock tunnel (JF-12) to understand the detonation combustion mode in the combustor of a standing oblique detonation ramjet (Sodramjet) engine prototype. Besides the traditional schlieren imaging to capture the oblique shock positions in the combustor, we implemented laser absorption spectroscopy of hydroxyl (OH) radical with three laser beams around the shock front. The Sodramjet engine combustor provides an ideal condition to measure OH because its molar fraction within the detonation front is estimated to exceed 0.03, which is much higher than other combustors. However, the harsh environment of freejet shock tunnel test poses a great challenge to the reliability of measurement systems. In this work, near-infrared tunable laser diode sensors centered at 1528 nm were chosen to measure the line-of-sight average of OH partial pressure. Continuous absorption measurements were conducted near the shock front, indicating a thin OH region after shock compression that could only be achieved by detonation combustion. The time-varied OH partial pressure was well predicted by the computational fluid dynamics (CFD) prediction. Our measurement results verified the detonation combustion of Sodramjet engine prototype and showed that in situ, real-time detonation diagnostics based on laser absorption spectroscopy is possible in freejet experiments of shock tunnel.

Experimental Observations on Film Cooling Effectiveness on Aerodynamic Heating in Hypersonic Flow

An experimental campaign was carried out to examine the effectiveness of film cooling on aerodynamic heating at hypersonic velocities. A flat-faced cylindrical model was tested at flow velocity of 2.5 km/s in a hypersonic shock tunnel. The test model was equipped with temperature sensors for convective heat flux measurement. The coolant gas, nitrogen, was injected through a 2 mm orifice housed at the center of the test model at different pressure ratios. Two different flow regimes were observed, steady and unsteady, with the steady flowfield having overall reduction in heat transfer compared to the unsteady one.

The Measurement of Radar–Plasma Signatures in a Hypersonic Shock Tunnel: Simulation and Experiment

The Measurement of Radar–Plasma Signatures in a Hypersonic Shock Tunnel: Simulation and Experiment

Arrays of glass wedges for multi-dimensional optical diagnostics.

There is a common need in the advancement of optical diagnostic techniques to increase the dimensionality of measurements. For example, point measurements could be improved to multi-point, line, planar, volumetric, or time-resolved volumetric measurements. In this work, a unique optical element is presented to enable multi-dimensional measurements, namely, an array of glass wedges. A light source is passed through the wedges, and different portions of the illumination are refracted by different amounts depending on the glass wedge angle. Subsequent optics can be used to focus the light to multiple points, lines, or planes. Basic characterization of a glass wedge array is presented. Additional wedge-array configurations are discussed, including the use of a periodic intensity mask for multi-planar measurements via structured illumination. The utility of this optical element is briefly demonstrated in (a)multi-planar flame particulate measurements, (b)multi-point femtosecond-laser electronic excitation tagging for flow velocimetry, and (c)multi-line nitric oxide molecular tagging velocimetry in a hypersonic shock-tunnel. One significant advantage of this optical component is its compatibility with high-energy laser sources, which may be a limiting factor with other beam-splitting or beam-forming elements such as some diffractive optics. Additionally, an array of glass wedges is simple and easily customizable compared to other methods for forming multiple closely spaced illumination patterns. Suggestions for further development and applications are discussed.

Experimental demonstration of forced initiation of kerosene oblique detonation by an on-wedge trip in an ODE model

The first free-jet experiment of a kerosene-fueled oblique detonation engine, conducted in a large-scale hypersonic shock tunnel, is reported in this paper. A novel initiation-control technique using a small on-wedge trip is proposed to overcome the initiation issue of oblique detonation waves (ODWs) encountered when liquid hydrocarbon fuel is used. The results show that the kerosene-fueled ODW fails to initiate within the length-limited combustor without the trip but is successfully initiated when the trip is used, which demonstrates the effectiveness of the proposed initiation-control method. The feasibility of kerosene-fueled oblique detonation propulsion technology is also demonstrated.Novelty and significance statement The feasibility of liquid-hydrocarbon-fueled oblique detonation propulsion technology was demonstrated by performing free-jet shock tunnel experiments of a kerosene-fueled oblique detonation engine. To address the initiation issue of oblique detonation waves encountered when liquid hydrocarbon fuel such as kerosene is used, a novel but simple initiation-control technique was proposed and experimentally validated. Stabilized oblique detonation waves were implemented in the combustor, and the detailed flow structures of different oblique detonation combustion modes were clarified. The proposed technique and the results provide significant reference to the future development of oblique detonation engines.

Viscous damping theory based recovery algorithm for force measurement in hypersonic flow regime

Prediction of aerodynamic forces is challenging task ahead aero-dynamists, specifically in the hypersonic flow regime. As per available literature in this field, there is a scarcity of recovery techniques to predict the applied aerodynamic load from a known response. Therefore, a Viscous Damping based novel algorithm is proposed for recovery of the unknown force from known response. Novelty of the presently proposed Viscous Damping technique resides in the fact that it is based on the physical principles those govern the motion of model-balance assembly. On the other hand, the contemporary techniques are data driven and hence provide approximate solution to the problem. For considering applicability of this technique, a bicone-shaped test model is designed and fabricated. This test model, integrated with an accelerometer force balance assembly, is calibrated using multiple point loads. Successful calibration experiments are then followed by experiments in IITB- Hypersonic Shock Tunnel (HST) at 0° Angle of Attack (AOA). Spike is also mounted on the model during force measurement experiments to assess the sensitivity of the Viscous Damping based technique. This technique has shown excellent agreement in force recovery with the conventional accelerometer force balance theory for the calibration as well as shock tunnel experiments. Therefore, presently proposed method is noticed to provide a definitive recovery and it is also seen to be sensitive since it has precisely recovered the reduced drag force in the presence of spike.

Hypersonic flow study in a pneumatically operated academic shock tunnel.

A hypersonic shock tunnel is a primary tool used for basic experimental research and may be used in engineering and university courses to study compressible flows involving shock waves. In the present study, a pneumatically operated shock tunnel is demonstrated for hypersonic flow studies. The high-pressure nitrogen gas is used to drive a pneumatic cylinder, which is used to burst the thin metal diaphragms. Tunnel-free stream conditions are quantified using the measured pressure values and by applying shock tube relations. The free-stream Mach number of 5.5-7.2 is achieved by varying the bursting pressure and test gas pressure from 2.1 to 4.5 bars and 0.2 to 0.5bar, respectively. The simulation is performed and the shock standoff distance quantified, and the stagnation pressures measured. The results demonstrate that the pneumatically operated tunnel enhanced operation capacity compared to the manually operated tunnel and well suits the academic hypersonic research and developmental activities.

Peak heat flux prediction of hypersonic flow over compression ramp under vibrationally excited free-stream condition

In hypersonic shock tunnel experiments, the high-temperature reservoir gas expands and accelerates so rapidly that there is not enough time for vibrational energy relaxation. As a result, thermal nonequilibrium gas flow is frequently encountered in the test section, and this significantly affects the measured heat flux. In this paper, hypersonic compression-ramp flows are studied numerically to investigate the effect of incomplete vibrational energy accommodation on the separation flow structure and peak heat flux in the reattachment region under low-to-medium Reynolds number and high Mach number conditions. Numerical results and theoretical analysis suggest that the vibrational energy accommodation has no noticeable impact on the length scale of the separation zone, but strongly influences the peak heat flux of the separated ramp flows. Decomposing the peak heat flux into translational–rotational energy and vibrational energy components, qtr and qv, respectively, we find that qv/qtr characterizes the nonequilibrium degree of the vibrational energy accommodation. A formula for predicting the peak heat flux is then proposed, taking the effect of incomplete vibrational energy accommodation into consideration. Finally, surface heat flux measurements in a hypersonic shock tunnel indicate that a deviation of up to 13% in total peak heat flux could arise if vibrational energy accommodation is not considered under the vibrationally excited free-stream condition.

Methodology for the application of velocimetry by molecular tagging of hypersonic flows

A laser-induced NO fluorescence technique was applied to measure velocity in a hypersonic shock tunnel nozzle exit. For the application of the technique, a detailed study of the density and fluorescence lifetime of the tracer radical, flow velocity and effective test time is proposed, resulting in a methodology for the application of the technique in hypersonic pulsed facilities. The study has demonstrated that it is necessary to jointly evaluate the flow velocity, the fluorescence lifetime of the radical and the width at half height of the laser beam, resulting in a kind of indicator for the feasibility of the technique. The variation of the laser incidence time with respect to the Pitot signal showed that it is not enough to select a stable Pitot pressure signal region to define the laser incidence time, preliminary trial and error analysis are necessary for each device used. Furthermore, the analysis of the velocity values calculated from the linear fit method shows that the adoption of such a method eliminates the effect of the systematic error of the measurements.

Experimental study on the effects of the cone nose-tip bluntness

In this Letter, hypersonic boundary-layer transition was investigated on a large-scale cone with a height of 3 m and a half-cone angle of 7° at a zero angle of attack in the JF-12 hypersonic flight duplicate shock tunnel. For the same freestream unit Reynolds number, with the increase in the bluntness Reynolds number, the transition Reynolds number has a trend of first increasing and then decreasing, showing a “transition reversal” phenomenon. As the bluntness increased, the high/low-frequency instability waves in the boundary-layer were modulated, which caused the boundary-layer transition to be delayed and then advanced.

Analytic investigation on error of heat flux measurement and data processing for large curvature models in hypersonic shock tunnels

Due to short test time, heat conduction was considered as transient in hypersonic shock tunnels. The heat flux measurement and data processing were operated basing on one-dimensional semi-infinite heat conduction theory. However, for models with local large curvature or small radius, it resulted in significant compression or expansion of space for heat transfer, or lateral heat conduction, which made the hypothesis of one-dimensional unsatisfied and errors. In this paper, approximate solutions for the unsteady heat conduction in cylindrically convex and concave shells were established, and were used for further analysis of the errors, with forms of heating load, location and curvature radius of heated surface taken into consideration.

Structured illumination-based transport-of-intensity phase imaging for quantitative diagnostics of high-speed flows.

A method for quantitative estimation of density in a high-speed flow field is presented, which uses structured illumination to interrogate the region of interest (ROI). Wavefront distortion suffered by the interrogating beam as it passes through the shock-induced flow field is investigated. Customized camera optics is used to measure the cross-sectional intensities of the exiting wavefront at two different planes, concurrently. A technique is proposed that uses the displacement of the structured light pattern to estimate the axial intensity derivative, ∂I∂z, which is the input to the transport-of-intensity equation (TIE) for phase estimation. The proposed method allows the use of images with larger defocused distances, ∂z, as it does not follow the conventional recipe of finite-difference (FD) approximation for intensity derivative estimation. This flexes two requirements of conventional TIE imaging: first, it brings flexibility in imaging optics, not restricted to small f-number objectives; second, use of a sensitive camera is optional, as ∂I∂z estimation does not use absolute intensities but information from the distorted pattern of structured light. The recovered phase from the TIE is used as an input to the phase tomography algorithm to obtain the refractive index followed by density distributions in the flow field quantitatively. The proposed phase estimation technique is verified through simulations first and then is implemented in experiments conducted in a hypersonic shock tunnel for flow Mach No. of 8.8. Estimated cross-sectional densities of the flow field around the aerodynamic test model are presented and compared with the numerically estimated values, which shows good agreement.

Experimental study on the evolution of mode waves in laminar boundary layer on a large-scale flat plate

In this research, to study the hypersonic boundary-layer transition, experiments were conducted on a large-scale flat plate with a length of 3.2 m at a zero angle of attack in the hypersonic shock tunnel duplicating flight conditions. Surface-mounted piezoelectric pressure sensors and coaxial thermocouples were, respectively, used to measure the pressure fluctuations and wall heat transfer. The spatial distribution of heat transfer was used to distinguish the transition. Under the test conditions of Ma = 7.0, T0 = 2120 K, and Re∞ = 6.08 × 105 m−1, no transition occurred, and under the test conditions of Ma = 7.0, T0 = 2220 K, and Re∞ = 1.23 × 106 m−1, the transition position was s = 2.06 m. The repeatability of the experiment was found to be good. Furthermore, focus was placed on the spectral and spatial/temporal evolution characteristics of pressure fluctuations in the laminar boundary layer. The experiment captured the three frequency distributions of mode waves in the laminar flow zone. Among the mode waves distributed in the three frequency bands, the low-/high-frequency bands were dominant, and the mid-frequency band exhibited a staged contribution. The amplitude energy percentages of the high- and low-frequency mode waves exhibited opposite trends in both time and space, which means that the disturbance energy will be distributed among the various harmonics in the laminar stage.

Experimental investigation on aerothermal effects of forward-facing cylindrical and parabolic cavity in hypersonic flow

Thermal protection of hypersonic vehicles is of great concern during ascent and re-entry conditions due to higher aerothermal loads. The passive forward-facing cavity is exhibited a substantial reduction of aero heating effects. The experimental studies on unsteady aspects of the flow considering the cavity geometries and heat transfer effects within the cavity are not abundant. In the present study, we investigated the heat transfer variations and nature of shock around a blunt body with cylindrical and parabolic cavity geometries. Experiments are conducted in hypersonic shock tunnels using air as the test gas for flow enthalpy conditions of 3.2 and 5.4 MJ/kg for a free-stream Mach number of 10.9 and 10.2 respectively. The cavity flow is established by verifying the measured heat flux, pressure, and shock oscillation results. The unsteady heat flux measurements and high-speed schlieren imaging is performed to study the effects. The results reveal that the flow with the parabolic cavity geometry shows a significant increase in the mean shock standoff position of about 14% and 30% compared to the blunt-nosed body, and about 7%–18% compared to the cylindrical cavity geometry, with a reduction of bow shock oscillations. As a result, a substantial reduction of heat flux about 34%–59% on the nose surface is noticed and found it is effective for mitigating the aerodynamic heating at higher flow enthalpy conditions. The flow with a cylindrical cavity geometry indicated the substantial increase in the heat flux due to the shock interaction and the significant heat flux variation in the cavity region is recognized and observed that such effects were not prevalent with the parabolic cavity geometry.

Experimental investigation of blunt cone model at hypersonic Mach number 7.25

Experiments were carried out in hypersonic shock tunnel in Defence Research and Directorate Laboratory at hypersonic Mach number of 7.25 using an 11.37° apex-angle blunt cone model. Heat flux measurement was carried out on cone model at different angles of attack with angle of rotation ϕ = 0° to 360° in steps of 45° with vacuum sputtered platinum thin film sensors. The measured experimental value of heat transfer data at stagnation point was compared with theoretical value estimated Fay and Riddell correlation. As angle of rotation was increased from ϕ = 0° to ϕ = 180°, the shock wave became closer to model surface due to high density ratio across the shock wave and consequently heat transfer rate became higher.

Gas-Surface Energy Exchange Characterization Around a Cone in the Free-Piston-Driven Shock Tunnel

This work describes numerical and experimental research in the Indian Hypersonic Shock Tunnel 3 on a blunt cone model and related computation of the facility parameters: nozzle reservoir conditions, nozzle transit time, and freestream conditions. Specialized codes accounting for real gas effects and nonequilibrium are used to obtain facility parameters. Starting at freestream conditions, a physicochemical model is developed in Fluent to simulate the dissociating environment within the shock layer over the model, assuming thermodynamic equilibrium. The model is validated, then invoked with a surface reaction of exothermic chromium oxidation, to study model erosion and gas-surface energy exchange. A mathematical analysis is performed to estimate surface oxidation rate, heat release, and its effect on shock layer temperature and aerodynamic heating. Both the quantities show considerable increase. Experimental measurement of surface heat flux and temperature using thin film gauges and two-color ratio pyrometry, respectively, also show an increase. The Fluent model with an iterative technique is used to show that the actual temperature at stagnation point is about 90 K higher than its counterpart based on line-of-sight averaged measurements. Finally, analytical calculations are performed to obtain the reaction rate parameter as a measure of the degree of dissociation in the shock layer.

Experimental investigation of hypersonic flight-duplicated shock tunnel characteristics

Hypersonic air-breathing propulsion is one of the key techniques for future aviation and the ground aerodynamic testing for full scale test models with sufficient test time at flight conditions is of fundamental importance for verifying hypersonic air-breathing engines. Based on the backward detonation-driven concept, the hypersonic flight-duplicated shock tunnel (or JF-12 shock tunnel) has been successfully constructed and calibrated. This facility is capable of reproducing airflow for Mach numbers ranging from 5 to 9 at an altitude of 25–50 km, with a test duration of more than 100 ms. To quantify the performance of the shock tunnel, experiments were conducted to investigate the aerodynamic characteristics of the test flows and the effects of several critical techniques that play important roles in the operation of the shock tunnel. The stagnation pressure was constant within ±5 $$\%$$ and the average stagnation pressure varied by less than 0.048 $$\%/$$ ms. The variation of the stagnation pressure in repeated experiments is less than 2.0 $$\%$$ , indicating the good repeatability of the wind tunnel. The non-uniformity of the Mach number in the core flow field at the nozzle exit was within ±2.5 $$\%$$ . Additional, a uniform flow field is established upstream of the nozzle exit. The axial gradients of the flow field are small since the Mach number varies less than 1.7 $$\%/$$ m. Findings regarding the ignition technology, diaphragm ruptures, detonation driver capacity, incident shock-wave decay, and tunnel operation mode are also presented. The findings of this study are not only helpful for operating the shock tunnel, but can also assist the future development of hypersonic wind tunnels.

Noise reduction for temperature-sensitive paint measurement contaminated by strong background radiation in a high enthalpy hypersonic tunnel

The strong background radiation in high enthalpy hypersonic shock tunnels has posed severe challenges for measurement using luminescent coatings. We proposed a solution for reducing background radiation from time-resolved temperature-sensitive paint (TSP) measurement in a hypersonic flow with Ma = 6.5 and T0 = 3525 K. The TSP was applied on an inlet ramp model, and the images were taken by a high-speed camera at 2 kHz under a modulated excitation. The strong background radiation led to a low signal-to-noise ratio and significant errors for the first half of the 130-ms test duration. Accordingly, three noise reduction methods were developed and evaluated based on temporal reconstruction, spatial reconstruction and robust principal component analysis (RPCA), respectively. The RPCA method showed the best performance that successfully recovered high-quality TSP data for a majority of test duration (t ≥ 40 ms).