Blowdown Wind Tunnel Research Articles

Surface Pressure Measurements on a Forebody Using Pressure Sensitive Paint

Pressure sensitive paint (PSP) gives the full-field pressure pattern on the model surface with a high spatial resolution and is an alternative to traditional pressure taps used in a wind tunnel model. This study determines surface pressure patterns on the forebody of an AGARD B model in a compressible flow and the data is in comparison with a numerical simulation using ANSYS Fluent software. The lift and drag coefficients are calculated and compared with data in literature. The experiments use a blowdown wind tunnel. The freestream Mach number is 0.83 and the angle of attack ranges from -4° to 8°. The difference between the results using PSP and the numerical simulation is less than 5%.

Role of the crown in tree resistance against high winds.

Rather than using wooden sticks to simulate the breakage of trees in high winds as in most research, we employ fresh samples from camphor and Formosa gum with branches and leaves to certify the crucial role of the tree crown. By using a blowdown wind tunnel with a maximum wind speed of 50 m/s, we purposely reduce the number of leaves and show that the drag force will drop by as much as two thirds when half pruned. Based on real observations, we model the leaf by an open and full cone in the presence of light and strong winds, and calculate how their corresponding cross-sectional area A and drag force F vary with wind speed v. Different slopes before and after the formation of a full cone are predicted and confirmed when these two quantities are plotted in full-log scale. Compared to the empirical value, our simple model gave α=2/5 and 2/3 for A∝v^{-α} and β=4/5 and 2/3 for F∝v^{β} at low and high winds. Discrepancies can be accounted for by including further details, such as the reorientation of open cones and the movement of branches.

Measurements of Metastable Molecules in Nonequilibrium Supersonic Flows

Nonintrusive laser diagnostics are used for the measurements of metastable nitrogen molecules in the lowest excited electronic state, , in a nonequilibrium flow, blowdown supersonic wind tunnel. The tunnel is operated with nitrogen at plenum pressures of Torr, with the flow expanding through a two-dimensional contoured nozzle to the Mach number of . The steady-state run time of the tunnel is approximately 10 s. The flow is excited by a repetitive nanosecond pulse discharge operated in the plenum or in the nozzle throat. Time-resolved absolute populations in the plenum are measured by single-pass, continuous wave tunable diode laser absorption spectroscopy (TDLAS). In the supersonic test section, absolute populations are measured by cavity ring down spectroscopy (CRDS), using a tunable pulsed laser system operated at 10 Hz. During each run, 50 single-shot ring down traces are acquired, demonstrating good shot-to-shot reproducibility. The results demonstrate that the cavity ring down time is not affected by the supersonic flow. populations and the flow temperature are inferred from the single-shot CRDS data. At the conditions when the flow is excited by the discharge in the nozzle throat, population in the supersonic test section is measured by both CRDS and TDLAS diagnostics. The two diagnostic techniques are complementary and can be used for characterization of nonequilibrium reacting flows over a wide range of pressures, including short-run-time high-enthalpy flow facilities.

Numerical Analysis and Wind Tunnel Validation of Low-Temperature Ablators undergoing Shape Change

Ground testing of ablative materials aims at providing critical data on the material behavior under hypersonic reentry conditions. This is normally done in plasma wind tunnel facilities. However, non-negligible technical challenges are faced in order to duplicate the real flight conditions, such as inducing the recession of space-relevant ablative materials, which requires sufficiently high inflow total enthalpies, and/or reproducing the actual hypersonic flow velocity, which requires sufficiently high inflow Mach numbers. Often, ground facilities which are providing one requirement are lacking the other one and vice-versa. A possible solution is to use low-temperature ablators in continuous hypersonic blow-down tunnels, where aerodynamic and ablative tests with considerable shape change effects may be performed under reasonably low total temperature conditions and with affordable test durations. These substances are readily available, and they sublimate or ablate in a fashion that can be described fairly accurately by theory. This work has the objective to numerically characterize the shape change of such materials in hypersonic conditions, concurrently providing a validation against literature data and from a dedicated experimental ground test campaign. The numerical procedure relies on ad-hoc mesh generation/evolution strategies taking into account the material shape change, and is based on subsequent steady-state Computational Fluid Dynamics (CFD) computations coupled with a customizable gas-surface interaction wall boundary condition. Preliminary numerical simulations helped the design of the experiments to be carried out in the von Karman Institute (VKI) H-3 hypersonic wind tunnel, in particular for the identification of capsule geometry and size in order to maximize the shape change caused by ablation. Subsequently, camphor is identified as the most suitable low-temperature ablator to be used in the experimental campaign after a thorough analysis of its surface reaction thermodynamics and kinetics. Results from the CFD approach are first compared with a literature experimental test case and then with those of the previously designed experiments, featuring a camphor sub-scale capsule, underlying advantages and limits of the numerical procedure adopted. The obtained numerical and experimental results underline how it is possible to obtain a relevant shape change for relatively small exposure times by using low-temperature ablators in continuous hypersonic blow-down wind tunnels. Hence, results from this work can be used to support the design and sizing of the actual heat shield and the analysis of the capsule’s aerodynamics and stability, accounting for shape change effects, by establishing an appropriate similitude between in-flight and on-ground conditions.

Cylindrical Focused Laser Differential Interferometer

A novel configuration of focused laser differential interferometry (FLDI) was developed, simulated, and tested. The new configuration is referred to as cylindrical FLDI or CFLDI. Using cylindrical lenses in place of standard lenses changes the three-dimensional shape of the focusing beam from a cone shape for FLDI to a triangular sheet for CFLDI, allowing for probing of flat-plate boundary layers closer to the wall. The beam propagation and interference pattern were modeled to determine the transfer functions for one-dimensional sinusoidal disturbances of varying wavelength () and varying spanwise thickness () of the tested area. CFLDI was employed to measure the density perturbations in supersonic blowdown wind tunnel SBR-50 at the University of Notre Dame with the test section width of 0.076.2 m. It was determined that the CFLDI system sufficiently filtered the sidewall boundary layers to measure the spectra of density fluctuations in a plane wall boundary layer. However, accurate measurements of the density disturbances in the freestream flow must account for the contribution of the sidewall boundary layers. This serves as an initial validation of the CFLDI maintaining spatial filtering similar to a regular FLDI.

Aerotunelska verifikacija efekata isticanja mlaza na aerodinamičke karakteristike osnosimetričnog projektila

Experimental verification of base-bleed effects on aerodynamic characteristics for the supersonic flight stages was requested during the development of an axi-symmetric projectile. The experimental setup for testing base-bleed configurations of the projectile was designed, built and tested in the T-38 trisonic blowdown wind tunnel of the Military Technical Institute (VTI), Belgrade. Gas-generator was formed from a simple chamber with a single manifold pneumatically connected to a high-pressure air supply. Wind tunnel tests comprised measurement of aerodynamic forces and moments of the projectile with active base-bleed. Flow visualization was performed using the Schlieren method. It was of interest to verify the base drag reduction in the wind tunnel tests and to compare it with numerical predictions. In addition, the experience gained during these tests has been found to be helpful in the design of the future similar wind tunnel setups

Mode decomposition of disturbances in a supersonic flow

Controlled disturbances are inserted into the wind tunnel flow, and probing measurements of the fields of the amplitudes of pressure and phase velocity oscillations of these disturbances are performed. Based on the relations for inviscid interaction of long-wave vortex, entropy, and acoustic disturbances with the shock wave on a wedge and numerical simulations, the coefficients of conversion of various modes to pressure oscillations on the model surface are determined for the test conditions of a supersonic flow in the T-327B blowdown wind tunnel located at the Khristianovich Institute of Theoretical and Applied Mechanics of the Siberian Branch of the Russian Academy of Sciences. The method of mode decomposition for controlled disturbances with the use of disturbance conversion coefficients is implemented in the case of a flat plate model with a sharp leading edge aligned at different angles of attack and side-slip in the flow.

Modal Parameter Estimation of a Compliant Panel Using Phase-based Motion Magnification and Stereoscopic Digital Image Correlation

This paper demonstrates the use of broad-band phase based motion magnification (PMM) to improve the modal parameter estimation from high-speed stereoscopic digital image correlation (DIC). PMM is used as a diagnostic technique to investigate the free vibration response of a panel. The compliant panel, consisting of a thin polycarbonate sheet, forms a test section wall in a supersonic blow-down wind tunnel, where it is used to investigate supersonic fluid-structure interaction in the presence of shock wave boundary layer interaction. The panel is excited by an impact hammer and the transient deformation is captured using high-speed cameras. The original and motion-magnified images are input to a digital image correlation algorithm to calculate the out-of-plane deformation of the panel. The measured deformation is used to extract the modal parameters of the compliant panel. By using PMM as a preprocessing tool in a broad frequency band containing multiple structural modes, the signal to noise ratio of the measured deformation is improved. The use of PMM improves the estimated mode shapes, increasing the MAC value of the first mode compared to FEM predictions from 0.29 to 0.99. Motion magnification also improves the coherence between measured input force and panel deformation by up to 13% if suitable parameters are chosen.

Transonic Turbine Vane Endwall Film Cooling Using the Pressure-Sensitive Paint Measurement Technique

Abstract This work focuses on the parametric study of film cooling effectiveness on the turbine vane endwall under various flow conditions. The experiments were performed in a five-vane annular sector cascade facility in a blowdown wind tunnel. The controlled exit isentropic Mach numbers were 0.7, 0.9, and 1.0, from high subsonic to transonic conditions. The freestream turbulence intensity is estimated to be 12%. Three coolant-to-mainstream mass flow ratios (MFR) in the range 0.75%, 1.0%, and 1.25% are studied. N2, CO2, and Argon/SF6 mixture were used to investigate the effects of density ratio (DR), ranging from 1.0, 1.5, to 2.0. There are eight cylindrical holes on the endwall inside the passage. The pressure-sensitive paint (PSP) technique was used to capture the endwall pressure distribution for shock wave visualization and obtain the detailed film cooling effectiveness distributions. Both the high-fidelity effectiveness contour and the laterally (spanwise) averaged effectiveness were measured to quantify the parametric effect. This study will provide the gas turbine designer more insight on how the endwall film cooling effectiveness varies with different cooling flow conditions including shock wave through the endwall crossflow passage.

Unscented Kalman filter (UKF)–based nonlinear parameter estimation for a turbulent boundary layer: a data assimilation framework

A turbulent boundary layer is a ubiquitous element of fundamental and applied fluid mechanics. Unfortunately, accurate measurements of turbulent boundary layer parameters (e.g. friction velocity and wall shear τw) are challenging, especially for high-speed flows (Smits et al ). Many direct and/or indirect diagnostic techniques have been developed to measure wall shear stress (Vinuesa et al ). However, based on various principles, these techniques generally give different results with varying uncertainties. The current study introduces a nonlinear data assimilation framework based on the unscented Kalman filter (UKF) that can fuse information from (i) noisy and discretized measurements from stereo particle image velocimetry (SPIV), a Preston tube, and a MEMS shear stress sensor, as well as (ii) the uncertainties of the measurements to estimate the parameters of a turbulent boundary layer. A direct numerical simulation of a fully developed turbulent channel flow is used first to validate the data assimilation algorithm. The algorithm is then applied to experimental boundary layer data at Mach 0.3 obtained in a blowdown wind tunnel facility. Drag coefficients from control volume analysis of the SPIV and wall pressure data and laser interferometer skin friction measurements are used for independent cross-validation. The UKF-based data assimilation algorithm is robust to the uncertain and discretized experimental data and is able to provide accurate estimates of turbulent boundary layer parameters with quantified uncertainty.

Mixing enhancement in a subsonic-supersonic shear layer with a cavity splitter plate

The development of more efficient ramjet engines has motivated the study of enhanced mixing in compressible shear layers. Cavity-mixer is a widely used passive flow control tool, but its mixing enhancement mechanisms in compressible subsonic-supersonic shear layers has yet to be comprehensively explored. In this work, a cavity-mixer is used to increase mixing in a subsonic-supersonic shear layer, and the underlying flow control mechanisms are examined. Parametric experiments were performed on a laboratory-scale blowdown wind tunnel and an extraction wind tunnel. Numerical simulations were performed in a large eddy simulation framework. Schlieren and nanoparticle-based planar laser scattering techniques were adopted for mixing flow visualization. It is found that the growth rate of the cavity (length-to-depth ratio of 5) disturbed shear layer is 31% higher than the benchmark configuration at convective Mach number of 0.49∼0.56. It is demonstrated that the compressibility and length-to-depth ratio affect the cavity-mixer performance. High convective Mach number suppresses the cavity disturbance and weakens its mixing enhancement efficiency. Cavity-mixer with a length-to-depth ratio of 3 shows the highest experimental and numerical growth rate, compared to 5 and 7 under a convective Mach number around 0.49. Large-scale coherent vortices are confirmed in the subsonic-supersonic mixing process. Cavity disturbance promotes the three-dimensional characteristics of the flow field, increases the vorticity of large coherent structures, and advances the K-H instability position of shear layer, which plays a key contribution to mixing enhancement.

Off-design performance of 2D adaptive shock control bumps

Adaptive shock control bumps can exploit the on-design drag-reducing potential of 2D bumps, while mitigating their off-design performance deterioration through geometric modifications. In this study, experiments and simulations have been employed to investigate the wave-drag reducing potential of (actuated and unconstrained) 2D adaptive shock control bumps over a wide range of shock positions. Experiments were carried out in the Imperial College supersonic wind tunnel, modelling the adaptive bump as a flexible surface placed beneath a Mach 1.4 shock wave. 2D RANS CFD simulations of the flow in a parallel channel with a solid bump complement experiments. Wave drag was demonstrated to be proportional to the ratio of inlet to exit stagnation pressure in a blow-down wind tunnel for a given shock position. The shock exhibits a hysteretic behaviour when travelling in the wind tunnel working section, governed by the wave drag reducing potential of the bump. The actuated adaptive bump tested reduces wave drag over a wider operational envelope than solid bumps as experiments revealed the presence of three preferred structural configurations, which lead to a significantly enlarged hysteresis region. Finally, tests on unconstrained bumps were shown to increase wave drag, both on- and off-design, due to the unfavourable bump shapes that result from (only) passive actuation, suggesting that some constraints are required to achieve desirable surface deformations.

Effect of Dust-Proof Filters Attached to Fan Suction Port of Blowdown Wind Tunnel on Its Performance

Effect of Dust-Proof Filters Attached to Fan Suction Port of Blowdown Wind Tunnel on Its Performance

Flow characteristics of Axisymmetric Cavity Rear Wall Divergence Angle in a Scramjet Combustor

Non-reacting experimental study was performed on a rear wall angled cavity actuated supersonic flow of Mach 1.5 from a convergent divergent nozzle using a blowdown wind tunnel test setup. Ten different model combinations of double angled rear wall cavities is preferred for the study of improvements in the geometrical design of the combustor. Flow field properties of various cavity geometries were analyzed based on the key parameters like, wall static pressures, stagnation pressure loss to the flow and qualitative mixing of flow using momentum flux distribution. The static pressure is found to decrease inside the combustor with a decrease in the secondary dual rear wall angle below 90 degrees whereas value increases at the rear wall OWING to oscillation and recompression of shear layers inside the cavity region. In addition, the decrement in primary rear wall angle, an enhancement in mixing profile and a reduction in stagnation pressure loss are also observed.

Supersonic flow around a cylinder with a permeable high-porosity insert: experiment and numerical simulation

Results of experimental and numerical investigations of a supersonic flow around a cylinder with a frontal gas-permeable insert made of a high-porosity cellular material are presented. The measurements are performed in a T-327 supersonic blowdown wind tunnel at the free-stream Mach numbers $M_{\infty }=4.85$, 7 and 21 in the range of the unit Reynolds numbers $Re_{1\infty }=(0.6{-}13.5)\times 10^{6}~\text{m}^{-1}$. The drag coefficients for a cylinder with an aerospike and a cylinder with a frontal gas-permeable porous insert are obtained. For the cylinder with the frontal gas-permeable porous insert, variations of the insert length, cylinder diameter and pore size are considered, and the mechanism of drag reduction is found, which includes two supplementary processes: attenuation of the bow shock wave in a system of weaker shock waves, and formation of an effective pointed body. The experiments are accompanied by numerical simulations of the flow around the cylinder with the frontal high-porosity insert: the fields of parameters of the external flow and the flow inside the porous insert are obtained, the drag coefficients are calculated, and the shape of the effective body for the examined model is found. The structure of the high-porosity material is modelled by a system of staggered rings of different diameters aligned in the radial and longitudinal directions (skeleton model of a porous medium). Numerical simulations of the problem are performed by means of solving two-dimensional Reynolds-averaged Navier–Stokes equations written in an axisymmetric form. The experimental and numerical data reveal significant drag reduction in a wide range of supersonic flow conditions. The results obtained on the drag coefficient for the cylinder are generalized with the use of a parameter which includes the ratio of the cylinder diameter to the pore diameter in the insert and the Mach number. This parameter is proposed as a similarity criterion for the problem of a supersonic flow around a cylinder with a frontal high-porosity insert.

Turbine Vane Endwall Film Cooling Comparison From Five Film-Hole Design Patterns and Three Upstream Injection Angles

Abstract The effects of upstream injection angle on film cooling effectiveness of a turbine vane end wall with various endwall film-hole designs were examined by applying pressure-sensitive paint (PSP) measurement technique. As the leakage flow from the slot between the combustor and the turbine vane is not considered an active source to protect the vane endwall in certain engine designs, discrete cylindrical holes are implemented near the slot to create an additional controllable upstream film to cool the vane end wall. Three potential injection angles were studied: 30 deg, 40 deg, and 50 deg. To explore the optimum endwall cooling design, five different film-hole patterns were tested: axial row, cross row, cluster, midchord row, and downstream row. Experiments were conducted in a four-passage linear cascade facility in a blowdown wind tunnel at the exit isentropic Mach number of 0.5 corresponding to inlet Reynolds number of 380,000 based on turbine vane axial chord length. A freestream turbulence intensity of 19% with an integral length scale of 1.7 cm was generated at the cascade inlet plane. Detailed film cooling effectiveness for each design was analyzed and compared at the design operation conditions (coolant mass flow ratio (MFR) 1% and density ratio 1.5). The results are presented in terms of high-fidelity film effectiveness contours and laterally (spanwise) averaged effectiveness. This paper will provide the gas turbine designers valuable information on how to select the best endwall cooling pattern with minimum cooling air consumption over a range of upstream injection angle.

Effect of Wall Blowing on Hypersonic Boundary-Layer Transition

An investigation of outgassing effects on boundary-layer transition was carried out. This joint computational–experimental work mimicked heat-shield pyrolysis outgassing in atmospheric reentry conditions. A slender 7 deg half-angle cone with air wall blowing through a porous section near the apex was tested in the VKI-H3 Mach 6 hypersonic blowdown noisy wind tunnel. The steady transition onset location was measured using infrared thermography, whereas gaseous-naphthalene-based planar laser-induced fluorescence imaging enabled the visualization and quantification of the spatial characteristics of the instabilities. Linear stability theory, combined with the semiempirical method with , was used to predict the location of transition onset in the same configurations. Studies were performed at different freestream unit Reynolds numbers and blowing rates. Numerical and experimental results agreed within the experimental uncertainties, and they coincided with the previously reported advancement of the transition location due to wall blowing. The wave numbers of the most amplified second-mode instabilities obtained with the linear stability theory matched the observations done with planar laser-induced fluorescence, suggesting it was the growth of such perturbations that led to the transitioning of the boundary layer. The porous section was seen to destabilize the boundary layer for nonblowing configurations. Regarding the upstream advancement of transition associated with wall blowing, the numerical analysis suggested that it is a consequence of the increase in the range of unstable frequencies in the wall-blowing region. Blowing nonuniformities in the cone’s longitudinal direction were observed to have no influence on the local transition location under noisy conditions.

Реализация течения в гиперзвуковом воздухозаборнике с пространственным сжатием

The results of a numerical and experimental study of a hypersonic air inlet with spatial compression and a compact cross section of the internal channel are presented. Reducing the surface area of the channel simplifies the heat protection of the air inlet and the combustion chamber. Experimental studies were carried out in a blow-down wind tunnel for Mach numbers M∞ = 2-6. The pressure distributions on the compression surfaces and in the air inlet channel are obtained, the coefficients of the total pressure recovery and flow rate, the Mach number in the throat of the air inlet are determined. The influence of the side cheeks and of the boundary layer bleed has been found. It is determined that in a convergent air inlet significantly higher compression ratios are realized than in a two-dimensional air inlet with less drag. It was established that on the basis of solving the three-dimensional Euler equations and the boundary layer, it is possible to reliably predict the characteristics of the air inlet in off-design modes and determine the structure of the flow.

Design and Characterization of a High-Speed Subsonic Aeroacoustic Wind Tunnel

The high-speed aeroacoustic wind tunnel at Carleton University is a new facility commissioned with the purpose of facilitating experimental studies of turbulent boundary-layer (TBL)-induced surface pressure fluctuations. This research is primarily intended for applications related to aircraft cabin noise generation from structures exposed to high-speed flow. This open-jet blowdown wind tunnel is a unique facility in Canada and one of a few aeroacoustic wind tunnels in the world capable of achieving speeds up to Mach 0.8. Flow is delivered from a nozzle with dimensions of 6.1 by 15 cm to a test section enclosed within an anechoic chamber with internal dimensions of 1.9 by 0.88 by 1.95 m. This study details the complete design methodology for all major wind-tunnel components, including the numerical simulations performed in the validation of the designed components. The results of the test section background noise levels and TBL surface pressure fluctuation spectral behavior developed over a flat test section plate are compared with established data and empirical models available in the literature.

Effects of Upstream Step Geometry on Axisymmetric Converging Vane Endwall Secondary Flow and Heat Transfer at Transonic Conditions

This paper presents a detailed experimental and numerical study on the effects of upstream step geometry on the endwall secondary flow and heat transfer in a transonic linear turbine vane passage with axisymmetric converging endwalls. The upstream step geometry represents the misalignment between the combustor exit and the nozzle guide vane endwall. The experimental measurements were performed in a blowdown wind tunnel with an exit Mach number of 0.85 and an exit Re of 1.5×106. A high freestream turbulence level of 16% was set at the inlet, which represents the typical turbulence conditions in a gas turbine engine. Two upstream step geometries were tested for the same vane profile: a baseline configuration with a gap located 0.88Cx (43.8 mm) upstream of the vane leading edge (upstream step height = 0 mm) and a misaligned configuration with a backward-facing step located just before the gap at 0.88Cx (43.8 mm) upstream of the vane leading edge (step height = 4.45% span). The endwall temperature history was measured using transient infrared thermography, from which the endwall thermal load distribution, namely, Nusselt number, was derived. This paper also presents a comparison with computational fluid dynamics (CFD) predictions performed by solving the steady-state Reynolds-averaged Navier–Stokes with Reynolds stress model using the commercial CFD solver ansysfluent v.15. The CFD simulations were conducted at a range of different upstream step geometries: three forward-facing (upstream step geometries with step heights from −5.25% to 0% span), and five backward-facing, upstream step geometries (step heights from 0% to 6.56% span). These CFD results were used to highlight the link between heat transfer patterns and the secondary flow structures and explain the effects of upstream step geometry. Experimental and numerical results indicate that the backward-facing upstream step geometry will significantly enlarge the high thermal load region and result in an obvious increase (up to 140%) in the heat transfer coefficient (HTC) level, especially for arched regions around the vane leading edge. However, the forward-facing upstream geometry will modestly shrink the high thermal load region and reduce the HTC (by ∼10% to 40% decrease), especially for the suction side regions near the vane leading edge. The aerodynamic loss appears to have a slight increase (0.3–1.3%) because of the forward-facing upstream step geometry but is slightly reduced (by 0.1–0.3%) by the presence of the backward upstream step geometry.