Ludwieg Tube Research Articles

Investigation of streamwise streak characteristics over a compression ramp at Mach 4

Experiments of shock wave/boundary layer interactions over a nominally two-dimensional compression ramp are conducted in a Mach 4 Ludwieg tube tunnel. Measurements of Schlieren, Rayleigh scattering, and surface pressure are performed to present the relevant flow features. The effects of two parameters, namely the Reynolds number based on the length of the flat plate and the ramp angle, on the flow stabilities are focused on. Four ramp angles of 6°, 8°, 10°, and 12° are tested under a Reynolds number of 7.22 × 105, while two other Reynolds numbers (3.66 × 105 and 9.19 × 105) are investigated with a ramp angle of 10°. Streamwise streaks are observed downstream of the reattachment point. The spanwise wavelength of the streaks remains unchanged with different ramp angles, whereas it slightly decreases as the Reynolds number increases. Power spectral density results show that the flow is transitional in the streak region and becomes turbulent where streaks break down. When increasing the ramp angle or the Reynolds number, the streamwise length of streaks shrinks. Two different patterns are distinguished at the breakdown, resembling the two unstable modes observed in the breakdown of Görtler vortices. To clarify the underlying physics of the formation of streaks, global stability analysis and resolvent analysis are carried out. Two regions of maximum optimal gain are identified, which are associated with Mack's first mode and streaks. The former can serve as an initial seed of Görtler instability via nonlinear interaction, while the latter can be associated with transient growth due to the lift-up mechanism and Görtler instability.

Nonclassical wave propagation measurements in the high temperature vapor of D6\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\hbox {D}_{6}$$\\end{document} with the asymmetric shock tube for experiments in rarefaction waves (ASTER)

A novel test setup called the asymmetric shock tube for experiments on nonideal rarefaction waves (ASTER) has been commissioned at Delft University of Technology. The ASTER, which works according to the principle of Ludwieg tubes, is designed to generate and measure the speed of small and finite amplitude waves propagating in the dense vapors of fluids formed by complex organic molecules, therefore in the nonideal compressible fluid dynamics regime. The ultimate goal of the associated research is to prove the existence of nonclassical gasdynamics. The setup consists of a high-pressure charge tube and a vacuum tank separated by a glass disk equipped with a breaking mechanism for rarefaction waves experiments. When the glass disk is broken, an expansion wave propagates into the tube in the direction opposite to the fluid flow. The propagation speed of this wave is measured using a time-of-flight method with the help of four fast-response pressure sensors placed equidistantly in the middle of the tube. The charge tube can withstand pressures and temperatures of up to 15 bar and 400∘C\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$^{\\circ }\\mathrm{C}$$\\end{document}. Preliminary rarefaction experiments were successfully conducted using dodecamethylcyclohexasiloxane, D6\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$\\hbox {D}_{6}$$\\end{document}, as the working fluid and at pressures and temperatures of up to 9.4 bar and 372∘C\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$^{\\circ }\\mathrm{C}$$\\end{document} , respectively. The results of an experiment featuring the initial state for which a theoretical model predicts the nonclassical acceleration of rarefaction waves show that the propagation is qualitatively different from that put into evidence by experiments for which the propagation is classic. Upcoming setup improvements and experimental campaigns are planned with the objective of experimentally verifying the existence of nonclassical gasdynamics.Graphical abstract

Crossflow Instability on a Swept Fin–Cone with Variable Nose Bluntness in Mach 6 Flow

Experiments and computational analysis have been conducted to study instabilities in the boundary layer over a swept fin–cone geometry in hypersonic flow. Experiments were carried out at the United States Air Force Academy’s Mach 6 Ludwieg Tube. Infrared thermography was employed to assess the surface temperature distribution, particularly the thermal striations on the fin that indicated a crossflow-dominated transition. Increasing the unit Reynolds number increased heating and moved the crossflow-transition front closer to the fin leading edge. Stability analysis based on the linear parabolized stability equations identified unstable stationary crossflow modes across the range of tested unit Reynolds numbers, and the most amplified disturbances were found to have wavelengths of 2–4 mm. The experimentally observed thermal striations were analyzed and found to have initial wavelengths of 2–5 mm but eventually formed higher-amplitude structures with a wavelength of approximately 10 mm. Discrete roughness elements of different diameters and spacings were implemented near the fin leading edge. Boundary-layer transition onset was found to be delayed with a subsequent lowering of the fin surface heating, with discrete roughness elements having a wavelength of 13.33 mm.

Multi-point FLEET velocimetry in a Mach 4 Ludwieg tube using a diffractive optical element.

A diffractive optical element was paired with femtosecond laser electronic excitation tagging (FLEET) velocimetry and used to probe multiple locations in a high-speed wind tunnel. Two configurations were explored, one that uses the traditional method of viewing from a perspective orthogonal to the beam axis and another that uses a perspective parallel to the beam axis. In the latter, the FLEET emissions are viewed as points that can allow for FLEET measurements in a wall normal fashion without the laser needing to impinge upon the surface. The configurations are demonstrated in a Mach 4 Ludwieg tube, highlighting their utility in high-speed flow measurements.

Influence of deep gaps on laminar–turbulent transition in two-dimensional boundary layers at subsonic Mach numbers

Systematic experimental investigations on the influence of deep gaps on the location of laminar–turbulent transition are reported. The tests were conducted in the Cryogenic Ludwieg–Tube Göttingen, a blow-down wind tunnel with good flow quality, at eight different unit Reynolds numbers ranging from Re_1 = {17.5,,times 10^{6},mathrm{{m}^{-1}}} to 80,times,10^{6},mathrm{m}^{-1}, three Mach numbers, M= 0.35, 0.50 and 0.65, and various pressure gradients. A flat-plate configuration, the extended two-dimensional wind tunnel model PaLASTra was modified in order to allow the installation of gaps with nominal widths of 30 upmu m, 100 upmu m and 200 upmu m and a depth of d = {9,mathrm{mm}}. A maximum Reynolds number based on the gap width Re_{w} = Re_1 cdot w approx {16{,}000} was reached. Transition Reynolds numbers ranging from Re_{tr}approx 1 × 106 to 11 × 106 were measured, as a function of gap width, pressure gradient and Mach and Reynolds number. This systematic investigation facilitates a linear approximation of Re_{tr} dependent on the boundary layer shape factor H_{12} for various flow conditions and gap widths. It was therefore possible to conduct an investigation of Re_{tr} depending on Re_{1} and the relative change of the transition location depending on the gap width w. Incompressible linear stability analysis was used to calculate amplification rates of Tollmien–Schlichting waves and determine critical N-factors by correlation with measured transition locations. The change in the critical N-factor varDelta N by installation of the gap is investigated as a function of w and Re_w. It was found that a gap width of 30 upmu m reduces the critical N-factors in the range of varDelta N approx 0.5 pm 0.25, while gap widths of 100 upmu m and 200 upmu m reduce the critical N-factor in the range of varDelta N approx 1.5 pm 1. Interestingly, an increase in gap width from 100 to 200 upmu m was not found to induce smaller transition Reynolds numbers or reduced N-factors, which might be due to resonance effects.

Fast-Responding Pressure-Sensitive Paint Measurements of the IC3X at Mach 7.2

Global surface pressure measurements of a 5.7% scale AFRL Initial Concept 3.X vehicle (IC3X) were obtained using a fast-responding ruthenium-based pressure-sensitive paint (PSP) at the UTSA Mach 7 Ludwieg Tube Wind Tunnel at two different angles of attack, 0° and 2.5°. Static calibration of the paint was performed over a range of 0.386 kPa to 82.7 kPa to relate luminescent intensity to pressure. Details on the facility, paint preparation, application, calibration, and image processing techniques are provided in the manuscript. The results from statistical, spectral, and proper orthogonal decomposition (POD) analyses are presented to characterize the pressure field observed on the model. The experimental results qualitatively follow the expected trends and correspond to the occurrence of shock waves and expansion fans, which were visualized via Schlieren imaging. The theoretical pressure range obtained from conical shock analysis for 0° agrees with the experimentally derived pressure range for the model, and the outliers are attributed to errors in image registration. This study presents preliminary pressure measurements that pave the way for obtaining time-resolved global PSP measurements to train and validate aerothermodynamic machine learning models.

Flow Characterization of the UTSA Hypersonic Ludwieg Tube

The characterization of a hypersonic impulse facility is performed using a variety of methods including Pitot probe scans, particle image velocimetry, and schlieren imaging to verify properties such as the velocity, Mach number, wall boundary layer thickness, and freestream turbulence intensity levels. The experimental results are compared to the numerical simulations of the facility performed with Ansys Fluent to compare the design and operational conditions. The presentation of results in this manuscript is prefaced by a description of the facility and its capabilities. The UTSA Ludwieg tube facility can produce a hypersonic freestream flow with a Mach number of 7.2 ± 0.2 and unit Reynolds numbers of up to 200 × 106 m−1. The Pitot probe profiles of the 203-mm-square test section indicate a 152 ± 10 mm square freestream core with turbulence intensity values ranging from 1% to 2%. Schlieren imaging of the oblique shockwaves on a 15° wedge model provided an alternate means of verifying the Mach number. Particle image velocimetry and previous molecular tagging velocimetry results showed a good agreement with the Pitot probe data and numerical simulations in the key parameters including freestream velocity, wall boundary layer velocity profiles, and wall boundary layer thickness.

Wind tunnel measurements of dynamic aerodynamic coefficients using a freely rotating test bench

PurposeThis paper aims to propose a dedicated measurement methodology able to simultaneously determine the stability derivative Cmα̇ and the pitch damping coefficient sum Cmq + Cmα̇ in a wind tunnel using a single and almost non-intrusive metrological setup called MiRo.Design/methodology/approachTo assess the MiRo method’s reliability, repeatability and accuracy, the measurements obtained with this technique are compared to other sources like aerodynamic balance measurements, alternative wind tunnel measurements, Ludwieg tube measurements, free-flight measurements and computational fluid dynamics (CFD) simulations. Two different numerical approaches are compared and used to validate the MiRo method. The first numerical method forces the projectile to describe a pure oscillation motion with small amplitude along the pitch axis during a rectilinear flight, whereas the second numerical approach couples the one degrees of freedom simulation motion equations with CFD methods.FindingsMiRo, a novel and almost non-intrusive technique for dynamic wind tunnel measurements, has been validated by comparison with five other experimental and numerical methodologies. Despite two completely different approaches, both numerical methods give almost identical results and show that the holding system has nearly no impact on the dynamic aerodynamic coefficients. Therefore, it could be assessed that the attitude of MiRo model in the wind tunnel is very close to the free-flight one.Originality/valueThe MiRo method allows studying the attitude of a projectile in a wind tunnel with the least possible impact on the flow around a model.

High-Speed Fluid-Structure Interactions on a Compliant Panel Under Shock Impingement

This paper reports experimental results of fluid/structure interactions induced by shock impingement on a clamped-free/clamped-free compliant panel in a Mach 5.8 Ludwieg tube. A shock generator is used to statically impinge a shock on the panel with the aim of producing flow-induced vibrations. In turn, these vibrations create an aeroelastic coupling between the panel and the flowfield, which is dominated by the shock-wave/boundary-layer interactions. The influence of the structural compliance and shock impingement strength is assessed by testing a nominally rigid panel and compliant panels with thickness-to-length ratios of 0.0034 and 0.0068 with a shock generator at flow deflection angles of 0, 2, 3, 8, and 10 deg. One of the primary findings of the present work is that the aeroelastic response of the panel tends to shift from a structurally dominated to a fluid-dominated regime, which is sustained by either the freestream fluctuations or the fluid/structure coupling, depending on the structural compliance and shock impingement strength and location. This work is of fundamental nature, and its main goal is to analyze the interactions between high-speed flow and a thin compliant panel of canonical geometry. However, studies like this are crucial to characterize those complex couplings and provide experimental data that can help enhance the accuracy of aeroelastic modeling tools that feed into the design process of supersonic and hypersonic airframes.

Laser-Induced Diaphragm Rupture for Improved Sequencing and Repeatability in a Hypersonic Facility

For hypersonic facilities where the flow conditions are established through the rupture of a diaphragm, such as in the University of Southern Queensland’s hypersonic wind-tunnel facility, the variability in the flow conditions is related to the uncertainty of the pressure at which the diaphragm ruptures. Variability in the diaphragm rupture pressure also results in uncertainty of the time at which the diaphragm will rupture. For experiments that require knowledge of when the test flow will be initiated, the sequencing of events relative to the flow onset is difficult when the flow is initiated using the natural rupture of a diaphragm. The challenge of experiment sequencing that arises due to rupture pressure variability is addressed by introducing a laser for rapid thermal weakening of the diaphragm. Event sequencing challenges are discussed in the context of free-flight testing, including model release strategies for such testing. The work proceeds through a review of Ludwieg tube flow initiation strategies and a discussion of the present context, which requires a reliable method for sequencing the retraction of the free-flight model holder. The natural variability of strength of the Mylar diaphragms in the present work is found to result in around uncertainty in rupture pressure. This rupture pressure variability is demonstrated to have a significant temperature dependence through empirical results and engineering models. Implementation of the laser-induced diaphragm rupture method is demonstrated to enhance repeatability in generating the flow conditions; the variability in rupture pressure was reduced to when the laser method was used. Based on the remaining sequencing uncertainties with the laser-induced rupture method and practical speeds for model platform retraction, uncertainty in the positioning of the free-flight models at the time of flow onset is shown to be .

Velocity measurements in a hypersonic flow using acetone molecular tagging velocimetry.

In the present work, a non-intrusive diagnostic technique known as molecular tagging velocimetry was used to collect quantitative freestream velocity measurements in the Mach 7 Ludwieg Tube Wind Tunnel located at The University of Texas at San Antonio. This laser-based diagnostic technique used a single Nd:YAG 4th harmonic laserline to excite acetone molecules seeded in the flow field. From the resulting emitted light, mean and instantaneous velocity profiles in the hypersonic freestream flow and facility boundary layer were measured. Uncertainty in the velocity measurements for individual test runs is estimated at ≤ 8% while overall 1D freestream mean velocity measurements were recorded with ±2.4% (± 21 m/s) accuracy. The effect of acetone seeding on the speed of sound was also quantified.

Experimental study of bluntness effects on hypersonic boundary-layer transition over a slender cone using surface mounted pressure sensors

In this work, we studied the bluntness effect on the hypersonic boundary-layer transition over a slender cone at Mach 6 with interchangeable tips in a noisy Ludwieg tube tunnel before the so-called “transition reversal” phenomenon occurs. The evolution of instability waves is characterized using surface flush-mounted pressure sensors deployed along the streamwise direction within unit Reynolds number from 4E+ 6/m ≤ Reunit ≤ 10E+ 6/m, and the bluntness of the cone nose ranges from 0.1 mm to 5 mm. Power spectral density (PSD) of pressure fluctuation indicates that small nose bluntness (ReR ≤ 2000) has little influence on the evolution of instability waves along the hypersonic boundary-layer, whereas with a moderate nose size (2000 ≤ ReR ≤ 5000), the hypersonic boundary layer transition is delayed monotonically as the nose radius increases before the boundary-layer turns into fully laminar without instability waves. The delaying effect can be attributed to the increased entropy-layer swallowing distance with a large tip radius. Instability wave characterization reveals that the second mode instability wave plays a dominant role before the transition reversal happens. The quadratic phase locking of second mode instabilities can be identified by bispectral analysis, and it attenuates as the nose tip radius increases.

Assimilation of wall-pressure measurements in high-speed flow over a cone

A nonlinear ensemble-variational data assimilation is performed in order to estimate the unknown flow field over a slender cone at Mach 6, from isolated wall-pressure measurements. The cost functional accounts for discrepancies in wall-pressure spectra and total intensity between the experiment and the prediction using direct numerical simulations, as well as our relative confidence in the measurements and the estimated state. We demonstrate the robustness of the predicted flow by direct propagation of posterior statistics. The approach provides a unique first look at the flow beyond the sensor data, and rigorously accounts for the role of nonlinearity, unlike previous efforts that adopted ad hoc inflow syntheses. Away from the wall, two- and three-dimensional assimilated states both show rope-like structures, qualitatively similar to independent schlieren visualizations. Despite this resemblance, and even though the planar second modes are the most unstable upstream, three-dimensional waves must be included in the assimilation in order to accurately reproduce the wall-pressure measurements recorded in the AFRL Ludwieg Tube facility. The results highlight the importance of three-dimensionality of the field and of the base-state distortion on the instability waves in this experiment, and motivate future measurements that probe the three-dimensional nature of the flow field.

Calibration and performance characterization of a Mach 5 Ludwieg tube.

Calibration, commissioning, and design features of a new Mach 5 Ludwieg Tube wind tunnel at the University of Arizona are discussed. Mach number uniformity and free-stream noise levels are measured using a Pitot rake at a range of unit Reynolds numbers and at multiple spanwise and streamwise positions. The wind tunnel is shown to have a free-stream Mach number of 4.82 with maximum variance less than 0.8% (and less than 0.5% at most streamwise positions). The average free-stream acoustic noise level in the core (based on Pitot pressure) is shown to be less than 1.2% at an intermediate Reynolds number with some regions dropping locally below 1.0%. The core flow region is measured to be 282.4mm (11.1 in.) in diameter at the nozzle exit.

Experimental study on surface arc plasma actuation-based hypersonic boundary layer transition flow control

Effective control of hypersonic transition is essential. In order to avoid affecting the structural profile of the aircraft, as well as reducing power consumption and electromagnetic interference, a low-frequency surface arc plasma disturbance experiment to promote hypersonic transition was carried out in the Φ0.25 m double-throat Ludwieg tube wind tunnel at Huazhong University of Science and Technology. Contacting printed circuit board sensors and non-contact focused laser differential interferometry testing technology were used in combination. Experimental results showed that the low-frequency surface arc plasma actuation had obvious stimulation effects on the second-mode unstable wave and could promote boundary layer transition by changing the spectral characteristics of the second-mode unstable wave. At the same time, the plasma actuation could promote energy exchange between the second-mode unstable wave and other unstable waves. Finally, the corresponding control mechanism is discussed.

Impact of Wavy Wall Surface on Hypersonic Boundary-Layer Instability of Sharp Cone Model

A comprehensive investigation of the impact of wavy wall on hypersonic boundary layers is carried out in a Mach 6 Ludwieg tube based on a 7 deg half-angle sharp cone. Linear stability theory is applied on the cone base flow with both smooth and wavy surfaces first. The evaluation of amplification rate as well as the frequency of instability waves reveals that the second mode instability is suppressed downstream of the wavy wall region. The development of instabilities is then characterized using surface flush-mounted pressure sensors and focused laser differential interferometer (FLDI). The FLDI measurement is conducted within the wavy wall zone both at troughs and at crests. Good agreement is found between the surface pressure measurement and computation, except that instability waves are invisible shortly downstream of the wavy wall. Additionally, bispectral analysis indicates that the nonlinear interaction of instability waves downstream of the wavy wall region is suppressed. FLDI measurement along the streamwise direction shows that the second mode instability waves disappear in the wavy wall region, and the boundary-layer transition process is delayed. Spatial measurement with FLDI of boundary layer identifies various types of disturbance at the troughs and crests; moreover, the co-existence of low-frequency and high-frequency disturbances is observed off the wall surface downstream of the wavy wall.

Experimental investigation of surface roughness effect on a free-flight sphere in a Ludwieg tube

Experimental investigation of surface roughness effect on a free-flight sphere in a Ludwieg tube

A-Variant Design of Hypersonic Ludwieg Tube Wind Tunnel

To improve the operation efficiency of conventional Ludwieg tube tunnel, a fast-acting valve is usually deployed between the storage tube and Laval nozzle in order to separate the high- and low-pressure air. However, the disturbance produced by the fast-acting valve as well as the nonuniform heating of storage tube affects the flow quality of the hypersonic Ludwieg tube. Stimulated by the motivation of overcoming the mentioned problems, the feasibility of a novel configuration of hypersonic Ludwieg tube is explored in this research. Upon this study, we introduced an additional Laval nozzle and settling chamber to the conventional Ludwieg tube tunnel, while the cross-section area of the main Laval nozzle is smaller than that of the introduced Laval nozzle. To begin with, the flow evolution along the dual-throat Ludwieg tube was analyzed theoretically; subsequently, numerical simulations were conducted to study the starting process of the dual-throat Ludwieg tube tunnel, and emphasis was placed on the influence of different throat area ratios and settling chamber lengths upon the startup time of wind tunnel. Lastly, the flowfield of the dual-throat Ludwieg tunnel was characterized by measuring the Mach number distribution and pitot pressure fluctuation in the test section as well as the pressures along the overall tunnel. The results show that this novel configuration of Ludwieg tube is applicable and it offers the opportunity for flow quality improvement for conventional Ludwieg tubes with small penalty of reducing the tunnel’s efficient running time.

Linear Instabilities over Ogive-Cylinder Models at Mach 6

Computational investigations of a tangent ogive-cylinder geometry with varying forebodies at zero degrees angle of attack are presented. The model geometry and conditions are selected to match experiments conducted in the Air Force Research Laboratory (AFRL) Mach 6 Ludwieg Tube. Specifically, five forebodies of interest consisting of sharp and blunt ogives of 4 and 2 caliber and a hemispherical shape are studied at a freestream unit Reynolds number of . Boundary-layer-edge properties and wall-normal profiles of velocity and temperature are compared across streamwise locations past the ogive-cylinder junction. Modal stability analysis is used to characterize the most amplified frequencies corresponding to Mack’s first and second modes and those are found to agree with experimental results for the sharp forebodies. The entropy layer induced by blunt forebodies envelopes the boundary layer and stabilizes modal disturbances. Nonmodal analysis revealed a broadband set of disturbances present for the blunter forebodies, in agreement with experimental observations. Flow perturbation contours of most amplified planar and oblique disturbances are shown to qualitatively match wind tunnel schlieren images, with a switch from rope-like to elongated structures, i.e., from high-frequency Mack’s second modes to low-frequency Mack’s first modes, as the forebody angle for the sharp tip is increased, and from boundary-layer to entropy-layer disturbances as the bluntness is increased.

Experimental Measurements of Hypersonic Instabilities over Ogive-Cylinders at Mach 6

Hypersonic boundary-layer instability disturbance and transition experiments were performed in the Air Force Research Laboratory Mach-6 Ludwieg Tube on a 1-m-long ogive-cylinder model with interchangeable nose-tip geometries of varying ogive radii and nose bluntnesses. Boundary-layer disturbances were measured via focused laser differential interferometry, surface-mounted pressure sensors, and high-speed schlieren videography at unit Reynolds numbers ranging from to . Three primary disturbance structures were observed that were dependent on ogive radius and nose bluntness. Each disturbance structure exhibits unique characteristics that are closely tied to the leading-edge geometry of the model. The addition of nose bluntness is shown to have an overall delay in transition onset, and a possible blunt nose transition reversal is observed as bluntness is further increased. It is also shown that the effects of the ogive radius in addition to nose bluntness alter the flow instability disturbance and transition behaviors and are likely tied to a modification of the entropy layer structure. A novel curvature-based Reynolds number is proposed that successfully correlates the effects of multiple leading-edge radii of curvature to dominant downstream boundary-layer disturbances and transition onset locations.