Shock Tunnel Research Articles

Rate-dependent energetic processes in hypersonic flows

In celebration of the first 60 years of the Air Force Office of Scientific Research, several studies of hypersonic flows dominated by rate-dependent energetic processes are revisited. The work presented shows the evolution and advancement of computational capabilities in this area, and illustrates some key lessons learned over the previous decade or so. Early work with Leyva and Hornung in the California Institute of Technology T5 Free-Piston Shock Tunnel had the goal of validating thermochemical models for high-enthalpy flows. Several of these flows are re-analyzed with more advanced numerical methods, resulting in improved comparisons with the experimental measurements. This work was followed by a series of experiments in the Calspan-University at Buffalo Research Center (now CUBRC Inc.) facilities at lower enthalpy conditions. Initial comparisons were poor, but with a better understanding of the facility behavior and the inclusion of key finite-rate processes, excellent agreement was obtained for nitrogen flows. An interesting study related to plasmadynamics and finite-rate processes in a different type of flow is discussed. Finally, it is shown that recent advances in numerical methods that are beginning to enable the direct numerical simulation of key rate-dependent energetic processes in hypersonic flows.

Radical recombination in a hydrocarbon-fueled scramjet nozzle

To reveal the radical recombination process in the scramjet nozzle flow and study the effects of various factors of the recombination, weighted essentially non-oscillatory (WENO) schemes are applied to solve the decoupled two-dimensional Euler equations with chemical reactions to simulate the hydrocarbon-fueled scramjet nozzle flow. The accuracy of the numerical method is verified with the measurements obtained by a shock tunnel experiment. The overall model length is nearly 0.5m, with inlet static temperatures ranging from 2000K to 3000K, inlet static pressures ranging from 75kPa to 175kPa, and inlet Mach numbers of 2.0±0.4 are involved. The fraction Damkohler number is defined as functions of static temperature and pressure to analyze the radical recombination progresses. Preliminary results indicate that the energy releasing process depends on different chemical reaction processes and species group contributions. In hydrocarbon-fueled scramjet nozzle flow, reactions with H have the greatest contribution during the chemical equilibrium shift. The contrast and analysis of the simulation results show that the radical recombination processes influenced by inflow conditions and nozzle scales are consistent with Damkohler numbers and potential dissociation energy release. The increase of inlet static temperature improves both of them, thus making the chemical non-equilibrium effects on the nozzle performance more significant. While the increase of inlet static pressure improves the former one and reduces the latter, it exerts little influence on the chemical non-equilibrium effects.

Incipient thermal choking and stable shock-train formation in the heat-release region of a scramjet combustor. Part I: Shock-tunnel experiments

A series of experiments is performed in the High Enthalpy Shock Tunnel Göttingen to investigate the response of the HyShot II scramjet combustor to equivalence ratios close to the critical value at which the onset of thermal choking occurs. This critical equivalence ratio is first identified (for simulated Mach-8, 27-km altitude conditions) as 0.38–0.39. Subsequent experiments cover the range between this value and 0.50. As the HyShot II combustor has a constant-area cross section, the expected behavior in this equivalence-ratio range is the formation and upstream propagation of an unsteady shock train that would eventually lead to inlet unstart. Instead, however, the shock train is observed to become lodged in the heat-release zone of the combustion chamber, with a quasi-stable position that shifts upstream with increasing equivalence ratio; this behavior is distinct from the steady isolator shock trains seen in traditional dual-mode scramjet configurations. The shock-train development in the present experiments is characterized through fast-response surface pressure and heat-flux measurements, as well as simultaneous high-speed schlieren and OH∗ chemiluminescence visualization. Possible explanations for this unexpected behavior are proposed in terms of the increased heat-transfer to the combustor walls and a possible increased uniformity of heat release across the combustor cross-section caused by the presence of the shock train.

충격파 터널을 이용한 고속추진기관 시험 연구

충격파 터널은 초음속 및 극초음속 비행조건의 유동장을 지상에서 가장 유사하게 모사할 수 있는 시험설비로서, 거의 반세기 동안 전 세계적으로 유용하게 운용되고 있다. 국내의 극초음속 비행체에 대한 많은 관심의 결과로 충격파 시험장치에 대해 주목하고 있으며, 최근에 KAIST에서는 중간 크기의 충격파 터널을 갖추게 되었다. 본 논문은 개발된 충격파 터널의 성능과 최근 수행된 모델 스크램제트 시험자료를 제시하고 있다. 본 연구에서 모사된 자유흐름 유동조건은 마하수 5.7 비행속도에 해당하며, 이러한 비행조건에서 스크램제트의 초음속 연소기내에 유동 환경을 모사하였다. Shock tunnels are known to be capable of simulating flow-field environments of supersonic and hypersonic flights. They have been operated successfully world-wide for almost half a century. As a consequence of the strong interest in hypersonic vehicles in Korea, attention has been given on this type of facility and so an intermediate-sized shock tunnel has lately been built at KAIST. In the light of this, this paper presents our tunnel performance and some of the model scramjet test data. The freestream flow used in this work replicates a supersonic combustor environment for a Mach 5.7 flight speed.

Conjugate Heat Transfer Analysis for a Finite Thickness Cylinder in a Hypersonic Flow

Prediction of surface heating rates is of prime importance for the hypersonic flow regime. Experimental and conventional computational efforts overlook the heat transfer phenomenon in the solid due to the rigid assumptions involved in the solution methodologies. In order to address this fact, conjugate heat transfer (CHT) studies are carried out using various coupling techniques to examine their implementation abilities. Three types of solution methodologies are adopted, namely, decoupled, strongly coupled, and loosely coupled analysis. This study is also focused on looking into the effect of a hypersonic flow field on wall heat flux for a finite thickness insulating cylinder at moderately large time scales. Increase in wall temperature and decrease in surface heat flux have been noticed using strong and loose coupling techniques with an increase in simulation time. Decoupled fluid and solid domain analysis is found to be useful for typical shock tunnel test durations (∼1 ms) while investigations with loose coupling techniques are advisable for time scales corresponding to flight testing (∼1 s). Efforts are also made to reason the discrimination in prediction of stagnation point heat flux using conventional computational and experimental analysis.

Fluorescence Imaging of Reaction Control Jets and Backshell Aeroheating of Orion Capsule

Planar laser-induced fluorescence of nitric oxide was used to visualize the interaction of reaction control system jet flows on the afterbody of a hypersonic capsule reentry vehicle at the Calspan-University at Buffalo Research Center’s Large Energy National Shock Tunnel I reflected shock tunnel facility. The interaction of pitch and roll jets with the flowfield was investigated. Additionally, thin-film sensors were used to monitor heat transfer on the surface of the model to detect localized heating resulting from the firing of the reaction control system jets. Visualizations of the capsule shear layer using both planar laser-induced fluorescence and schlieren imaging compared favorably. The structure of the roll jet was found to be significantly altered due to interactions with the flowfield. Additionally, the presence of the roll jet appeared to change the nature of the shear layer from steady laminar to unsteady. The pitch jet structure was only disturbed in the far field. Comparison of the planar laser-induced fluorescence jet-fluid visualizations and the surface heat flux distributions indicate that the regions of enhanced aeroheating are not caused by the jet fluid itself impinging on the surface, but rather by the presence of jet-induced horseshoe vortices and shock wave/boundary-layer interactions.

Schlieren-based techniques for investigating instability development and transition in a hypersonic boundary layer

Three variants of schlieren techniques are employed to investigate the development of second-mode instability waves in the hypersonic boundary layer of a slender cone in a reflected shock tunnel. First, a previously proposed technique using high frame rate (i.e., at least as high as the dominant instability frequency) schlieren visualization with a continuous light source is shown to provide repeatable measurements of the instability propagation speed and frequency. A modified version of the technique is then introduced whereby a pulsed light source allows the use of a higher-resolution camera with a lower frame rate: this provides significant benefits in terms of spatial resolution and total recording time. A detailed picture of the surface-normal intensity distribution for individual wave packets is obtained, and the images provide comprehensive insight into the unsteady flow structures within the boundary layer. Finally, two-point schlieren deflectometry is implemented and shown to be capable of providing second-mode growth information in the challenging shock tunnel environment.

Transitional shock-wave/boundary-layer interactions in hypersonic flow

AbstractStrong interactions of shock waves with boundary layers lead to flow separations and enhanced heat transfer rates. When the approaching boundary layer is hypersonic and transitional the problem is particularly challenging and more reliable data is required in order to assess changes in the flow and the surface heat transfer, and to develop simplified models. The present contribution compares results for transitional interactions on a flat plate at Mach 6 from three different experimental facilities using the same instrumented plate insert. The facilities consist of a Ludwieg tube (RWG), an open-jet wind tunnel (H2K) and a high-enthalpy free-piston-driven reflected shock tunnel (HEG). The experimental measurements include shadowgraph and infrared thermography as well as heat transfer and pressure sensors. Direct numerical simulations (DNS) are carried out to compare with selected experimental flow conditions. The combined approach allows an assessment of the effects of unit Reynolds number, disturbance amplitude, shock impingement location and wall cooling. Measures of intermittency are proposed based on wall heat flux, allowing the peak Stanton number in the reattachment regime to be mapped over a range of intermittency states of the approaching boundary layer, with higher overshoots found for transitional interactions compared with fully turbulent interactions. The transition process is found to develop from second (Mack) mode instabilities superimposed on streamwise streaks.

Hypersonic Scramjet Testing via Diode Laser Absorption in a Reflected Shock Tunnel

A wavelength-multiplexed two-color tunable diode laser absorption spectroscopy sensor probing transitions near was developed to measure temperature and column density simultaneously across three lines of sight in a ground-based model scramjet combustor. High-enthalpy scramjet conditions equivalent to Mach 10 flight were generated with a reflected shock tunnel. The sensor hardware development and optical engineering to overcome noisy combustor conditions, including short test time, beam steering, and mechanical vibration, are discussed. A new scanned-wavelength-modulation spectroscopy technique was used to acquire the complete spectral lineshape while maintaining the high signal-to-noise ratio characteristic of wavelength-modulation spectroscopy measurements. Two combusting flow experiments were conducted, and the results compared with steady-state computational fluid dynamics calculations, with both the simulations and the measurements finding formation of from combustion during the test time. This work represents the first use of wavelength-modulation spectroscopy in a reflected shock tunnel and the first use of scanned-wavelength-modulation spectroscopy in a scramjet combustor.

Ultrasonic absorption characteristics of porous carbon–carbon ceramics with random microstructure for passive hypersonic boundary layer transition control

Preceding studies in the high enthalpy shock tunnel Gottingen of the German Aerospace Center (DLR) revealed that carbon fibre reinforced carbon ceramic (C/C) surfaces can be utilized to damp hypersonic boundary layer instabilities leading to a delay of boundary layer transition onset. To assess the ultrasonic absorption properties of the material, a test rig was set up to measure the reflection coefficient at ambient pressures ranging from 0.1 × 105 to 1 × 105 Pa. For the first time, broadband ultrasonic sound transducers with resonance frequencies of up to 370 kHz were applied to directly cover the frequency range of interest with respect to the second-mode instabilities observed in previous experiments. The reflection of ultrasonic waves from three flat plate test samples with a porous layer thickness between 5 and 30 mm was investigated and compared to an ideally reflecting surface. C/C was found to absorb up to 19 % of the acoustic power transmitted towards the material. The absorption characteristics were investigated theoretically by means of the quasi-homogeneous absorber theory. The experimental results were found to be in good agreement with the theory.

Shock tunnel experiments on control of shock induced large separation bubble using boundary layer bleed

Shock–Boundary Layer Interaction (SBLI) often occurs in supersonic/hypersonic flow fields. Especially when accompanied by separation (termed strong interaction), the SBLI phenomena largely affect the performance of the systems where they occur, such as scramjet intakes, thus often demanding the control of the interaction. Experiments on the strong interaction between impinging shock wave and boundary layer on a flat plate at Mach 5.96 are carried out in IISc hypersonic shock tunnel HST-2. The experiments are performed at moderate flow total enthalpy of 1.3 MJ/kg and freestream Reynolds number of 4 million/m. The strong shock generated by a wedge (or shock generator) of large angle 30.96° to the freestream is made to impinge on the flat plate at 95 mm (inviscid estimate) from the leading edge, due to which a large separation bubble of length (75 mm) comparable to the distance of shock impingement from the leading edge is generated. The experimental simulation of such large separation bubble with separation occurring close to the leading edge, and its control using boundary layer bleed (suction and tangential blowing) at the location of separation, are demonstrated within the short test time of the shock tunnel (∼600 μs) from time resolved schlieren flow visualizations and surface pressure measurements. By means of suction – with mass flow rate one order less than the mass flow defect in boundary layer – a reduction in separation length by 13.33% was observed. By the injection of an array of (nearly) tangential jets in the direction of mainstream (from the bottom of the plate) at the location of separation – with momentum flow rate one order less than the boundary layer momentum flow defect – 20% reduction in separation length was observed, although the flow field was apparently unsteady.

Roughness-Induced Transition of Hypervelocity Boundary Layers

Experiments were performed in the T4 free-piston reflected shock tunnel to explore the transition process of hypervelocity boundary layers. The experiments took place on a 1000-m-long by 300-m-wide flat plate with a sharp leading edge at representative conditions pertaining to those of an unpublished test of a Mach 12 scramjet. A streamwise row of thin-film heat transfer gauges along the centerline of the plate was used to determine the state of the boundary layer. It was found that, without any boundary-layer trips, the boundary layer remained laminar for the entire length of the plate. Multiple combinations of trip geometry and size were then used in an attempt to force the boundary layer into turbulence. The most successful geometry was the diamond trip configuration, although transition did not occur directly behind the trip. Transition occurred after a period of relaminarization, with the length of the delay region agreeing well with the literature.

Free-flight measurement technique in the free-piston high-enthalpy shock tunnel.

A novel multi-component force-measurement technique has been developed and implemented at the impulse facility JAXA-HIEST, in which the test model is completely unrestrained during the test and thus experiences free-flight conditions for a period on the order of milliseconds. Advantages over conventional free-flight techniques include the complete absence of aerodynamic interference from a model support system and less variation in model position and attitude during the test itself. A miniature on-board data recorder, which was a key technology for this technique, was also developed in order to acquire and store the measured data. The technique was demonstrated in a HIEST wind-tunnel test campaign in which three-component aerodynamic force measurement was performed on a blunted cone of length 316 mm, total mass 19.75 kg, and moment of inertia 0.152 kgm(2). During the test campaign, axial force, normal forces, and pitching moment coefficients were obtained at angles of attack from 14° to 32° under two conditions: H0 = 4 MJ/kg, P0 = 14 MPa; and H0 = 16 MJ/kg, P0 = 16 MPa. For the first, low-enthalpy condition, the test flow was considered a perfect gas; measurements were thus directly compared with those obtained in a conventional blow-down wind tunnel (JAXA-HWT2) to evaluate the accuracy of the technique. The second test condition was a high-enthalpy condition in which 85% of the oxygen molecules were expected to be dissociated; high-temperature real-gas effects were therefore evaluated by comparison with results obtained in perfect-gas conditions. The precision of the present measurements was evaluated through an uncertainty analysis, which showed the aerodynamic coefficients in the HIEST low enthalpy test agreeing well with those of JAXA-HWT2. The pitching-moment coefficient, however, showed significant differences between low- and high-enthalpy tests. These differences are thought to result from high-temperature real-gas effects.

Flow Characterization of a Detonation Gun Facility and First Coating Experiments

A computer-controlled detonation gun based spraying device has been designed and tested to obtain particle velocities over 1200 m/s. The device is able to be operated in two modes based on different flow-physical principles. In one mode, the device functions like a conventional detonation gun in which the powder is accelerated in a blast wave. In the other mode, an extension of the facility with a nozzle uses the detonated gas for an intermittently operated shock tunnel process in which the particles are injected into and accelerated by a quasi-steady high enthalpy nozzle flow with high reservoir conditions. Presented are experimental results of the operation without nozzle in which the device generates moderate to high particle velocities in an intermittent process with a frequency of 5 Hz. A hydrogen/oxygen mixture and Cu and WC-Co (88/12) powders are used in the experiments. Operation performance and tube outflow are characterized by time-resolved Schlieren images and pressure measurements. The particle velocities in the outflow are obtained by laser Doppler anemometry. Different substrate/powder combinations (Al/Cu, Steel/Cu, Al/WC-Co, and Steel/WC-Co) have been investigated by light microscopy and measurements of microhardness.

Interaction of single-pulse laser energy with bow shock in hypersonic flow

Pressure sensing and schlieren imaging with high resolution and sensitivity are applied to the study of the interaction of single-pulse laser energy with bow shock at Mach 5. An Nd:YAG laser operated at 1.06μm, 100mJ pulse energy is used to break down the hypersonic flow in a shock tunnel. Three-dimensional Navier–Stokes equations are solved with an upwind scheme to simulate the interaction. The pressure at the stagnation point on the blunt body is measured and calculated to examine the pressure variation during the interaction. Schlieren imaging is used in conjunction with the calculated density gradients to examine the process of the interaction. The results show that the experimental pressure at the stagnation point on the blunt body and schlieren imaging fit well with the simulation. The pressure at the stagnation point on the blunt body will increase when the transmission shock approaches the blunt body and decrease with the formation of the rarefied wave. Bow shock is deformed during the interaction. Quasi-stationary waves are formed by high rate laser energy deposition to control the bow shock. The pressure and temperature at the stagnation point on the blunt body and the wave drag are reduced to 50%, 75% and 81% respectively according to the simulation. Schlieren imaging has provided important information for the investigation of the mechanism of the interaction.

Boundary-Layer Stability Analysis of the High Enthalpy Shock Tunnel Transition Experiments

Recent experiments focused on the hypersonic boundary-layer transition process have been conducted at Japan Aerospace Exploration Agency’s free-piston High Enthalpy Shock Tunnel facility. Numerical simulations of the experiments were conducted to generate mean-flow solutions upon which boundary-layer stability analyses were performed. These computations included the effects of chemical reactions and internal energy relaxation. The disturbances in the boundary layer were calculated by solving the linear parabolized stability equations. Comparisons of transition location, disturbance amplification rates, and disturbance frequencies were made between those measured experimentally and those calculated from the stability analysis. The results suggest that an factor of 8.0 is the value that indicates the onset of transition for this facility. A disturbance frequency comparison between the computations and experiment showed a reasonable agreement for the locations at which the flow remained laminar. Additional theoretical operational conditions were examined to investigate effects of finite-rate chemistry and vibrational excitation. For the conditions examined, the inclusion or exclusion of chemical kinetics modeling in the mean-flow solution influenced boundary-layer stability to a greater extent than the inclusion or exclusion of chemical kinetics in the stability analysis alone. In general, the inclusion of these effects increased disturbance frequencies, peak amplification rates, and the distance over which the disturbance frequencies were amplified.

Experimental Investigation of Wall and Total Temperature Influence on a Shock Train

Experiments are carried out to investigate the shock train in an isolator model. In one model, the wetted surfaces can be heated up to 1000 K. The tests are performed in a hypersonic shock tunnel facility at Mach 7.7 with varying total temperature. The wall temperature is varied between 300 and 1000 K. The experiments show the influence of the wall temperature on the pressure distribution in the shock train. This effect is not covered by the correlation of Waltrup and Billig. The correlation is expanded with the influence of the wall heat flux to include these effects.

Particle Acceleration in a High Enthalpy Nozzle Flow with a Modified Detonation Gun

The quality of thermal sprayed coatings depends on many factors which have been investigated and are still in scientific focus. Mostly, the coating material is inserted into the spray device as solid powder. The particle condition during the spray process has a strong effect on coating quality. In some cases, higher particle impact energy leads to improved coating quality. Therefore, a computer-controlled detonation gun based spraying device has been designed and tested to obtain particle velocities over 1200 m/s. The device is able to be operated in two modes based on different flow-physical principles. In one mode, the device functions like a conventional detonation gun in which the powder is accelerated in a blast wave. In the other mode, an extension with a nozzle transforms the detonation gun process into an intermittent shock tunnel process in which the particles are accelerated in a high enthalpy nozzle flow with high reservoir conditions. Presented are experimental results of the operation with nozzle in which the device generates very high particle velocities up to a frequency of 5 Hz. A variable particle injection system allows injection of the powder at any point along the nozzle axis to control particle temperature and velocity. A hydrogen/oxygen mixture is used in the experiments. Operation performance and nozzle outflow are characterized by time resolved pressure measurements. The particle conditions inside the nozzle and in the nozzle exit plane are calculated with a quasi-one-dimensional WENO-code of high order. For the experiments, particle velocity is obtained by particle image velocimetry, and particle concentration is qualitatively determined by a laser extinction method. The powders used are WC-Co(88/12), NiCr(80/20), Al2O3, and Cu. Different substrate/powder combinations for varying particle injection positions have been investigated by light microscopy and measurements of microhardness.

Development of Laser-Induced Grating Spectroscopy for Application in Shock Tunnels

An optical setup for laser-induced grating spectroscopy is presented. This setup provides the possibility to induce and analyze electrostrictive gratings as well as thermal gratings. Switching between the techniques can be realized smoothly. Temperature measurements were carried out at different test gas conditions in a temperature range from 20 to 70°C and a pressure range from 0.1 to 2 MPa using a test chamber made of stainless steel. The optically measured temperature was compared with simultaneously recorded thermocouple data to demonstrate the feasibility of the presented method. As a sufficient amount of an absorbing species in the test gas is required for the formation of a thermal grating, the experiments were performed on nitrogen monoxide with concentrations of 10 and 125 ppm, respectively, diluted in molecular nitrogen. Thereby, the higher concentration favors the formation of a thermal grating. Because of the planned application of the setup on a shock tube and a shock tunnel, data gained from single-shot measurements were also compared with thermocouple data.

Dynamic Calibration and Validation of an Accelerometer Force Balance for Hypersonic Lifting Models

An accelerometer-based force balance was designed and developed for the measurement of drag, lift, and rolling moment on a blunt-nosed, flapped delta wing in a short-duration hypersonic shock tunnel. Calibration and validation of the balance were carried out by a convolution technique using hammer pulse test and surface pressure measurements. In the hammer pulse test, a known impulse was applied to the model in the appropriate direction using an impulse hammer, and the corresponding output of the balance (acceleration) was recorded. Fast Fourier Transform (FFT) was operated on the output of the balance to generate a system response function, relating the signal output to the corresponding load input. Impulse response functions for three components of the balance, namely, axial, normal, and angular, were obtained for a range of input load. The angular system response function was corresponding to rolling of the model. The impulse response functions thus obtained, through dynamic calibration, were operated on the output (signals) of the balance under hypersonic aerodynamic loading conditions in the tunnel to get the time history of the unknown aerodynamic forces and moments acting on the model. Surface pressure measurements were carried out on the model using high frequency pressure transducers, and forces and moments were deduced thereon. Tests were carried out at model angles of incidence of 0, 5, 10, and 15 degrees. A good agreement was observed among the results of different experimental methods. The balance developed is a comprehensive force/moment measurement device that can be used on complex, lifting, aerodynamic geometries in ground-based hypersonic test facilities.