High Enthalpy Shock Tunnel Research Articles

In-situ calibration for temperature-sensitive-paint heat-flux measurement on a finite base

This work incorporates in-situ calibration into the analytical inverse heat transfer solution for a finite base in temperature-sensitive-paint (TSP) measurements in high-enthalpy short-duration hypersonic wind tunnels. The thermal penetration parameter of the polymer layer (TSP) in the analytical inverse solution is determined by in-situ calibration based on the data obtained using a reliable heat flux sensor at a reference location. This method is applied to TSP images obtained on a wedge in a high-enthalpy shock tunnel.

Influence of transpiration cooling on second-mode instabilities investigated on hypersonic, conical flows

In the present study, the influence of active cooling on hypersonic boundary-layer transition at different Mach numbers, from 7 up to 10, is investigated. The analyses are carried out on a $$7^\circ$$ half-angle, blunted cone with different nose radii and various gas injection mass flow rates. In all cases, low mass fluxes, which do not inducing visible shocks in the schlieren images, are applied. As injection gas nitrogen is used. At the considered free stream conditions, second modes are the dominant boundary-layer instabilities, which are consequently the focus of this study. The stability analyses are performed by means of the stability code NOLOT, NOnLOcal Transition analysis, of the German Aerospace Center (DLR). The influence of different mass injections on the frequencies and growth rates of the second modes is analyzed in detail. The effect on the transition onset locations is discussed. The numerical predictions are compared with experimental results. The experimental data referred to in the present study were obtained in the DLR High Enthalpy Shock Tunnel Gottingen.

Nozzle design for a large-scale high-enthalpy shock tunnel

The shock tunnels are the important ground-test facilities to perform high-speed aerodynamic experiments, and the nozzle is the key component to generate uniform flow in the shock tunnel test. In this paper, the design method for the axisymmetric contoured nozzles is studied for a large-scale high-enthalpy shock tunnel. The design method includes the design of inviscid contour line and the correction of viscous boundary-layer. Once the inviscid nozzle lines are determined, the corrections for the displacement thickness of boundary layer are performed. The traditional design methods for the hypersonic nozzle ignore the displacement thickness at the throat section, which is small compared to the characteristic length (radial distance at the throat section). This assumption has been applied successfully to many supersonic and hypersonic nozzles. However, the correction of displacement thickness at the throat should be taken into consideration for the large-scale nozzles at the high Mach number. In the high-enthalpy shock tunnels, the real-gas effects and the chemical non-equilibrium are obvious and cannot be ignored. In this study, the necessary corrections for the high-temperature effects and the viscous boundary-layer were carried out using the Navier-Stokes solver, where the effects of chemical non-equilibrium are considered in the simulations. Based on the method of characteristics (MOC), the parameters of characteristic grid region, such as the specific-heat-ratio γ, are determined by results of CFD simulations. Specific-heat-ratio can be calculated by NASA’s fitting method using the present numerical gas individual species. Then, the boundary-layer corrections are performed by using the calculated displacement-thickness iteratively. In this paper, a nozzle for Mach 17 was designed using the improved Sivells’s method and was analyzed based on the computational fluid dynamics (CFD) simulations. The diameter at the nozzle exit is 2.5 m, the total temperature and total pressure are 7400 K and 30 MPa, respectively.

Experimental Investigation of Roughness Effects on Transition on Blunt Spherical Capsule Shapes

Boundary-layer transition plays an important role in aerodynamic heating of blunt capsule shapes. Transition may occur from disturbances created by surface roughness and fluctuations of the freestream. The present paper reports on a number of related experiments on Apollo-type capsule models in a range of hypersonic wind tunnels. The Adjustable Contour Expansion facility of the Texas A&M University provided Mach 6 experiments with distributed roughness at relatively low Reynolds numbers, , with denoting the capsule diameter. Observed boundary-layer transition data align well with current estimates of transient growth theory. Larger Reynolds numbers, , could be assessed in the hypersonic Ludwieg tube Braunschweig of the Braunschweig University of Technology in Germany. Distributed roughness height values were broadly varied during these experiments, and different fluctuation levels of the freestream were considered. The results indicate roughness-induced transition for larger roughness heights, in accordance with correlations based on transient growth theory, and transition due to freestream disturbances, at subcritical roughness. Effects of total flow enthalpy on transition with isolated roughness were investigated in the High Enthalpy Shock Tunnel facility of the Japan Aerospace Exploration Agency. Here, roughness height, Reynolds number, and flow total enthalpy were varied. The paper reviews the most important results of these experiments and provides analysis, taking into account transient growth theory and numerical simulation data.

High-Enthalpy Effects on Hypersonic Boundary-Layer Transition

In the present study, three boundary-layer stability codes are compared based on hypersonic high-enthalpy boundary-layer flows around a blunted 7 deg half-angle cone. The code-to-code comparison is conducted between the following codes: the Nonlocal Transition analysis code of the DLR, German Aerospace Center (DLR); the Stability and Transition Analysis for hypersonic Boundary Layers code of VirtusAero LLC; and the VKI Extensible Stability and Transition Analysis code of the von Kármán Institute for Fluid Dynamics. The comparison focuses on the role of real-gas effects on the second-mode instability, in particular the disturbance frequency, and deals with the question on how far not accounting for real-gas effects compromises the stability analysis. The experimental test cases for the comparison are provided by the DLR High Enthalpy Shock Tunnel Göttingen and the Japan Aerospace Exploration Agency High Enthalpy Shock Tunnel. The focus of the comparison between the stability results and the measurements is, besides real-gas effects, the influence of uncertainties in the mean flow on the stability analysis.

The High Enthalpy Shock Tunnel Göttingen of the German Aerospace Center (DLR)

The High Enthalpy Shock Tunnel Göttingen (HEG) of the German Aerospace Center (DLR) is one of the major European hypersonic test facilities. It was commissioned for use in 1991 and was utilized since then extensively in a large number of national and international space and hypersonic flight projects. Originally, the facility was designed for the investigation of the influence of high temperature effects such as chemical and thermal relaxation on the aerothermodynamics of entry or re-entry space vehicles. Over the last years its range of operating conditions was subsequently extended. In this framework the main emphasis was to generate test conditions which allow investigating the flow past hypersonic flight configuration from low altitude Mach 6 up to Mach 10 in approximately 33 km altitude. The studies performed in HEG focused on the external as well as internal aerodynamics including combustion of hydrogen in supersonic combustion and the investigation of transition from laminar to turbulent hypersonic flow.

Performance of Copper Calorimeter for Heat Transfer Measurement in High Enthalpy Shock Tunnel

Copper calorimeter, based on a calorimetric principle, offers a solution for heat transfer measurement in high enthalpy situation, especially in the erosive flow of high enthalpy shock tunnels. In this study, we numerically investigated the measuring performance of copper calorimeters. Non-ideal effects, such as heat loss to the insulator around and replacement of the average temperature of the copper element by the junction temperature, were discussed in detail. The influences of copper element thickness, copper/constantan wires thickness and sensor diameter were also estimated, with the aim to provide theoretical guidance for the design of copper calorimeter. In addition, corresponding experiments in JF10 high enthalpy shock tunnel were carried out against the data of coaxial thermocouples for verification. Results showed that the non-ideal thermal environment of a copper calorimeter (heat exchange with its surroundings) would result in a smaller measuring heat flux comparing to the one actually loaded; proper thickness of copper element matching the effective test time of shock tunnel was suggested. Besides, preliminary experimental results with corrections showed reasonable agreement with the heat flux of thermocouples, with an average deviation of 8%. Over all, this gauge developed extends and supplements the high enthalpy shock tunnel heat transfer measurements made by other techniques.

Force and Moment Measurements on a Free-Flying Capsule in a Shock Tunnel

Experiments to measure the forces and moments acting on a blunted-cone capsule model are conducted in the High Enthalpy Shock Tunnel Göttingen. A free-flying arrangement is employed, whereby the model is initially suspended by weak threads that are detached by the arrival of the flow, allowing unrestrained flight during the steady test time. High-speed shadowgraphy is used to capture the model motion. Induced forces and moments are measured using both internally mounted accelerometers and visualization-based tracking techniques. Improvements to the visualization-based techniques allow the elimination of several previously identified sources of error. Drag, lift, and pitching-moment coefficients are derived over a range of angles of attack up to 6 deg, and excellent agreement is obtained between the techniques. Results are also compared with those from numerical simulations, with agreement found to lie within the experimental error for the drag and pitching coefficients but slightly outside for the lift coefficients. The precision of the visualization-based results is such that uncertainty in the freestream conditions is the dominant source of experimental error in the drag coefficient, and it becomes comparable to or larger than the standard error in the lift and pitching accelerations at the maximum angle of attack tested.

Determination of heat transfer into a wedge model in a hypersonic flow using temperature-sensitive paint

Heat loads on spacecraft traveling at hypersonic speed are of major interest for their designers. Several tests using temperature-sensitive paints (TSP) have been carried out in long duration shock tunnels to determine these heat loads; generally paint layers were thin, so that certain assumptions could be invoked to enable a good estimate of the thermal parameter ρck (a material property) to be obtained—the value of this parameter is needed to determine heat loads from the TSP. Very few measurements have been carried out in impulse facilities [viz. shock tunnels such as the High Enthalpy Shock Tunnel Gottingen (HEG)], where test times are much shorter. Presented here are TSP temperature measurements and subsequently derived heat loads on a ramp model placed in a hypersonic flow in HEG (specific enthalpy h 0 = ~3.3 MJ kg−1, Mach number M = 7.4, temperature T ∞ = 277 K, density ρ ∞ = 11 g m−3). A number of fluorescence intensity images were acquired, from which, with the help of calibration data, temperature field data on the model surface were determined. From these the heat load into the surface was calculated, using an assumption of a 1D, semi-infinite heat transfer model. ρck for the paint was determined using an insitu calibration with a Medtherm coaxial thermocouple mounted on the model; Medtherm ρck is known. Finally presented are sources of various measurement uncertainties, arising from: (1) estimation of ρck; (2) intensity measurement in the chosen interrogation area; (3) paint time response.

Understanding scramjet combustion using LES of the HyShot II combustor

Detailed understanding of the physical processes occurring in the combustor of a scramjet engine is crucial for enabling this technology and is considered the most promising for hypersonic flight. Laboratory, ground testing, and flight-testing experiments, together with simulations, comprise the tools available to fill the current gap in scramjet combustor knowledge. Here, a computational study has been carried out for the HyShot II scramjet combustor using Large Eddy Simulation (LES) models together with one skeletal and one comprehensive reaction mechanism. Based on a survey of the experimental data available for the HyShot II combustor, we focus on the High Enthalpy Shock Tunnel Göttingen (HEG) operating condition XIII, emulating flight conditions at 28km altitude. To account for slight experimental run-to-run variations, two simulations are performed at different equivalence ratios. The LES are found to capture the experimental wall-pressure and heat-flux data well compared to the measurement data, with marginal influence of the reaction mechanism. The LES results are subsequently used to analyze the flow, mixing, and combustion processes involved. The tools employed include conventional, as well as recently developed methods, such as the Takeno flame index and Chemical Explosive Mode Analysis. These methods give a concise description of the interactions between flow, fuel–air mixing, and combustion, and it is discovered that supersonic combustion is a combination of auto-ignition and non-premixed flame regions and self-igniting fronts. Furthermore, ignition is enabled by shocks, and the supersonic flame is very different in nature to subsonic turbulent flames.

Research on electron density measurements for high enthalpy flows

The high enthalpy shock tunnel, which can provide the high temperature gas conditions for hypersonic flight, is an effective ground facility for the research of reentry problems. It also has the ability of studying real gas effect. The hypersonic test flow condition, with the stagnation enthalpy of 16 MJ/kg and stagnation temperature of 7900 K, was achieved in the JF-10 high enthalpy shock tunnel. The electron density of the flow-field for the 10 degree wedge model was measured using the electrostatic probe, which can obtain sufficient spatial resolution without influencing the flow structure. Results showed that the electrostatic probe has high sensitivity and the spatial electron density can be preferable acquired. The constant bias method can get the dimensionless electron density coupling with the flow field parameters, but the new developed sweeping circuit can get the quantitative value on the basis of effectively reducing the induced noise in the circuit.

Comparative study on aerodynamic heating under perfect and nonequilibrium hypersonic flows

In this study, comparative heat flux measurements for a sharp cone model were conducted by utilizing a high enthalpy shock tunnel JF-10 and a large-scale shock tunnel JF-12, responsible for providing nonequilibrium and perfect gas flows, respectively. Experiments were performed at the Key Laboratory of High Temperature Gas Dynamics (LHD), Institute of Mechanics, Chinese Academy of Sciences. Corresponding numerical simulations were also conducted in effort to better understand the phenomena accompanying in these experiments. By assessing the consistency and accuracy of all the data gathered during this study, a detailed comparison of sharp cone heat transfer under a totally different kind of freestream conditions was build and analyzed. One specific parameter, defined as the product of the Stanton number and the square root of the Reynold number, was found to be more characteristic for the aerodynamic heating phenomena encountered in hypersonic flight. Adequate use of said parameter practically eliminates the variability caused by the deferent flow conditions, regardless of whether the flow is in dissociation or the boundary condition is catalytic. Essentially, the parameter identified in this study reduces the amount of ground experimental data necessary and eases data extrapolation to flight.

Numerical study on wall pressure over cone region of blunt-nosed body in high enthalpy shock tunnel HIEST

A cause of an overestimation of the computed surface pressure on a blunted cone in high-temperature hypersonic flow is explored. The overestimation was observed in a free-piston shock tunnel at the stagnation of H0=15.6 and 10.1 MJ/kg. The sensitivity analysis reveals that a reduction of the upstream translational temperature in the range of 100 to 300 K substantially improves the agreement of the surface pressure with the measured data. As the cause of the lower upstream translational temperature, radiative cooling effect is included in the thermochemical nonequilibrium calculation in the nozzle, the translational temperature at the nozzle exit is reduced to about 250 K. Using the obtained flow variables as the upstream boundary condition, the computed pressure agrees quite well with the experimental data. In order to clarify whether other variables such as translational–vibrational relaxation time, chemical reaction rates, and upstream chemical composition could be the cause of the discrepancy, uncertainty quantification is employed. It is shown that these parameters of the thermochemical model and upstream chemical composition have minor effect on the agreement of surface pressure. It is concluded that the observed discrepancy in the surface pressure is due to radiative cooling effect of high temperature gas in the nozzle region.

Passive Hypersonic Boundary Layer Transition Control Using an Ultrasonically Absorptive Coating With Random Microstructure: Computational Analysis Based on the Ultrasonic Absorption Properties of Carbon-Carbon

Previous investigations in the High Enthalpy Shock Tunnel Göttingen of the German Aerospace Center (DLR) show that carbon fiber reinforced carbon ceramic (C/C) surfaces can be utilized to damp hypersonic boundary layer instabilities resulting in a delay of boundary layer transition onset. Linear stability analyses were performed using the DLR stability code NOLOT, NOnLocal Transition analysis code. To adapt the boundary condition to account for the characteristics of porous C/C material, the ultrasonic absorption properties of C/C were investigated experimentally and theoretically. Therefore, a test rig was set up to directly measure the reflection coefficient in the frequency and pressure range corresponding to the test conditions in HEG. In this frame, the reflection of ultrasonic waves from flat plate test samples with different porous layer thicknesses was investigated and compared to an ideally reflecting surface.The obtained results were used to improved the boundary condition used for stability analysis above porous surfaces. The numerical results, using the original as well as the improved boundary condition, were compared with wind tunnel tests. These experiments were performed at Mach 7.5 and different unit Reynolds numbers. A 7∘ half-angle cone model with a nose radius of 2.5mm and a total length of 1077mm was used. One-third of the metallic model surface in circumferential direction was replaced by C/C ceramics. The comparison between numeric and experiments includes the investigations of the second modes, the damping of the these modes and the resulting transition shift.

Backscattering measurements of plasma coated target in high-enthalpy wind tunnel

When high-speed vehicles enter into the atmosphere, plasma sheath may be excited around due to aerodynamic heating, resulting in difficulties in communicating and changes of electromagnetic scattering properties. Those facts have received lots of attention due to their influences on the aerospace communication and radio telemetry applications. While analytic and numerical studies have been carried out by many native institutions on the electromagnetic radiation/scattering problems in the presence of plasma sheath, there remains the lack of measurement data to support and verify those researches. This work reports the backscattering measurements for the target surrounded by plasma sheath in the ground high-enthalpy shock tunnel facility. Using the step frequency sweeping mode of a commercial instrument, i.e., vector network analyzer, we conduct the experiments in the JF-10 high-enthalpy shock tunnel. The dynamic electromagnetic scattering measurement must be completed on a time scale of ms while the shock tunnel is running. The implementation details are demonstrated in this work, including the experimental configurations, data processing procedures, timing synchronization, and discussion on the relationships between the air flow status and measured target scattering signals. The influences of the plasma sheath on the target RCS (radar cross section) in the C band are successfully and clearly observed. The influence of the air flow status on the measured data can be concluded as follows: the front section of high-speed air flow lasting about 0.5-1 ms will change the measured signal dramatically, which should be avoided in observation due to its instability; the effective plasma sheath lasts only about 2 ms, resulting in an overall reduction on the target RCS by about 2 dB in the measurements. Afterwards, the effects by the plasma sheath on the target scattering vanish quickly.

S1910301 感温塗料を用いた高温衝撃風洞におけるHIFiRE模型の境界層遷移可視化

A visualization experiment of hypersonic boundary layer transition by Temperature-Sensitive Paint (TSP) was performed. The test facility was High-Enthalpy Shock Tunnel (HIEST) at Japan Aerospace Exploration Agency (JAXA) Kakuda Space Center. An elliptic cone model with a 2:1 aspect ratio, which is a full-scale model of HIFiRE-5, was used as the test model. This study visualized the hypersonic boundary layer transition on the side of the test model. TSP was excited by a Blue-LED and the luminescence from TSP was detected by a high-speed camera. Temperature distribution on the side of the test model was obtained by TSP. Moreover, boundary layer transition along the centerline of the model and in approximately halfway between the centerline and leading edge were visualized. The trend of temperature rise along the centerline is same in TSP and thermocouples, which is inserted on the centerline of the test model.

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.

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.

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.

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.