Supersonic Wind Tunnel Research Articles

Experimental investigation of shock train unsteady movement in a mixing duct

The structure and dynamic of the shock train in a mixing duct are investigated experimentally in a supersonic air breathing wind tunnel. High-speed schlieren technique and pressure measurements are utilized for the data acquisition. Nozzle clapboards are applied to separate the incoming flow into upper secondary flow, primary flow, and lower secondary flow. The Mach number of all incoming flows is 2.05, whereas the static pressures between primary and secondary flow stay unmatched. Experimental results indicate that nozzle clapboards lead to complex background waves and mixing layers in the mixing duct. As the throttling valve rotates at a constant speed, two motion patterns alternatively govern the shock train movement, namely jump movement and slow movement. Judging from the path of shock train leading edges, adverse pressure gradient at the wall of background flow induces jump movement, while the slow movement only appears at the zone of favorable pressure gradient. Besides, leading shock structures continuously change in the forward propagation process. Four types of leading shock structures are discovered namely regular reflection, convex-stem Mach reflection, straight-stem Mach reflection, and concave-stem Mach reflection. Based on the numerical results of the background flow field, these special structures are explained theoretically by analyzing the variation of the wavefront Ma contour lines.

Measurements of a Mach 3.4 turbulent boundary layer using stereoscopic particle image velocimetry

Experiments were performed within the Rutgers University Mach 3.4 supersonic wind tunnel to quantify the turbulent boundary layer in the test section, using stereoscopic particle image velocimetry (SPIV). Seeding of the tunnel was accomplished using atomized mineral oil, resulting in a Stokes number of 0.09. The Reynolds number based on momentum thickness and the friction-velocity Reynolds number were 64,000 and 4320, respectively, marking this as a high Reynolds number boundary layer. SPIV measurements were capable of resolving the mean velocity profile, along with turbulence quantities, permitting comparison against the known literature. Furthermore, the convection of vorticity concentrations representing hairpin vortices was observed within instantaneous vector fields. The resulting Reynolds stresses demonstrate that the boundary layer had typical values of anisotropy, while comparing well with data obtained within a slightly adverse pressure gradient. This was particularly noticeable from the trends in the Reynolds stresses, which exhibited higher-than-expected peak levels. These data provide a unique measurement of a supersonic turbulent boundary layer at Reynolds numbers not often encountered within existing data sets.

Spatially resolved mean and unsteady surface pressure in swept SBLI using PSP

Strong crossflow and swept separation are key aspects of the flow dynamics of 3-D shock wave/boundary layer interactions. This study explores the global surface pressure field beneath the canonical interaction produced by a sharp fin with deflection angle $$15^\circ$$ with a turbulent incoming boundary in a Mach 2 freestream flow. This corresponds to an interaction of moderate strength, and the unsteady pressure distribution captures pressure fluctuations associated with separation shock motion upstream of the interaction. Details of the PSP calibration are also described where the calibration process combines both a priori (with separately painted test coupon) and in situ calibration (with pressure tap measurements during test). Flow features are identified directly from the quantitative pressure distribution and compared to qualitative surface oil flow visualizations. The technique facilitates measurement of the pressure distribution on surfaces that have been difficult or impossible to instrument, such as the face of the shock-generating fin. The unsteady paint response is captured simultaneously with unsteady pressure transducers on the surface underneath the interaction, and a frequency response function based on this comparison is presented. As the results discussed herein demonstrate, the use of PSP allows one to capture significantly more information about this complex, highly three-dimensional interaction with details that are not easily obtained using traditional sensors, while also providing a more informed global view of the interaction. The utility and limitations of the technique in application to supersonic wind tunnel experiments are discussed.

Design of a continuous portable indraft supersonic wind tunnel

Programs in mechanical and aeronautical engineering commonly include courses in compressible fluid flow. As such, learning can be greatly enhanced if theory is taught in conjunction with hands on experimentation. While supersonic wind tunnels are not uncommon at many universities, such facilities are generally of the blow down configuration. Consequently, run time is very short and ear protection is required during operation, potentially hindering instruction. Furthermore, blow down configurations are typically expensive and large. This article presents the design and manufacture of a continuous, indraft, miniature supersonic wind tunnel. The tunnel was designed for a nominal test section Mach number of 2; validation indicated a Mach number of 1.96 was achieved. Vacuum was provided by a regenerative blower. The facility is portable and quiet; measurements indicated that the sound level around the tunnel when operational was less than 81 dB (compared to 119dB generated by the department’s blow down supersonic wind tunnel).

Optimum design of the inlet port profile of a ramjet extended range projectile

The air-inlet is one of the important components of the stamped extended range engine. Its performance directly affects the thrust of the ramjet engine, which in turn affects the range of the projectile. Because the stamped extended range projectile has the characteristics of simple structure, light weight and small volume, the inlet profile of the stamped extended range projectile is generally unadjustable, and the air-inlet surface quality directly determines the performance of the air-inlet. And the resistance characteristics of the shells. Since the air-inlet is associated with the curved portion of the projectile, the curved portion is one of the main sources of resistance to the projectile, so it is necessary to optimize the design of the air-inlet profile. Firstly, based on the CFD flow field simulation, the optimal design method of the air-inlet profile is studied. The parametric modeling, flow field simulation and optimization strategy of the air-inlet are integrated on the platform to realize automatic optimization. The optimization of the nozzle is optimized by this optimized design method, which proves its feasibility. Then, the single-objective optimization and multi-objective optimization of the air-inlet profile are carried out, and the geometric model is established to more accurately describe the flow field characteristics and obtain a more accurate air-inlet profile. Finally, a supersonic wind tunnel test was carried out on the optimized air-inlet, and the wall static pressure and the total outlet pressure of the windward side and the leeward side were measured and compared with the simulated Mach number of 2.5 and 2.0. Comparative Results. The two are basically the same, indicating that the numerical simulation results can basically reflect the pressure change and the change of the total pressure in the air-inlet. Therefore, the reliability of simulation-based optimization results is high.

Oblique Shock Control with Steady Flexible Panels

Flexible panels deforming under pressure loads have been suggested as a passive form of adaptive oblique shock control. This study investigates oblique shock–boundary-layer interactions on a steady flexible panel in a Mach 2.0 flow. Experiments were performed in the Imperial College supersonic wind tunnel, where shock generators were used to produce an oblique shock followed by a corner expansion. A parametric study was conducted, exploring different shock impingement positions and shock–expansion distances. The steady aerostructural response is studied using schlieren photography, static pressure distributions, photogrammetry measurements, and surface oil flow visualization. Two-dimensional numerical simulations were performed to assess the effects of the flexible panel on downstream total pressure recovery. These were validated against experimental wall pressure distributions and measurements from a Pitot rake. Results show reductions in both separation length (of up to 40%) and stagnation pressure losses (of up to 10%) if the flexible plate is used. These improvements occur for a range of shock positions spanning approximately 50% of the panel length and for all the shock–expansion distances considered. A model that captures the flow physics responsible for these trends is proposed. The results highlight the potential of flexible panels for practical oblique shock control.

Morphology of Quasi-Direct-Current Discharges Collocated with Fuel Jets in a Supersonic Crossflow

This study is focused on the morphology of a quasi-DC discharge generated concurrently with a jet of fuel injected normally into Mach 2 crossflow. It was observed that the filamentary plasma follows the local flowfield and localizes in the shear layer between the core flow and fuel jet. The details of behavior were explored by means of stereoscopic two-dimensional imaging with subsequent image processing and three-dimensional reconstruction of the plasma filament shape. On average, the plasma filament generated by the plasma-injection module has a total length over 100 mm and a plasma power on the order of 5 kW with a typical duration of 100 ms. Specifically, the discharge is located in a mixing layer, where the fuel–oxidizer ratio varies from down to stoichiometric ratio further downstream. Such a behavior of filamentary plasma promotes the fuel–air mixing, and is favorable for fuel ignition and flameholding. Tests were performed in the SBR-50 supersonic wind tunnel at the University of Notre Dame.

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.

Design and analysis of five probe flow analyser for subsonic and supersonic wind tunnel calibration

The flow angularity in a wind tunnel plays a major role in test section when testing models. The models are being tested at a mixture of flow velocity. Usually a subsonic wind tunnel is calibrated using a normal Pitot tube and supersonic wind tunnels with high efficiency probes. The calibration of a wind tunnel includes determining the Mach number range ad pressure range throughout the test section of the flow throughout the pattern of operating velocity. In the analysis it is noted that the sensitivities of the calibrating instruments are maximum when it is either cone or wedge shaped. When the wedge or cone angles are maximum, more accurate results are also obtained, but it’s not possible to have a maximum angle cone or wedge because of the detached wave and tunnel blockage factors. Hence in order to overcome this difficulties in measurement, a five probe flow analyzer has been designed. The five probe flow analyzer is used to determine the wind tunnel test section parameters like flow angularity, Pitot pressures, static pressures, wave angles, and the presence of test section noise. Over to these parameter, the stagnation states can also be determined using this analyzer. The proposed model is designed by getting optimum solutions from previous analysis and the existing data by considering all the aerodynamic and mechanical loading. This instrument designed is to be tested experimentally after carrying two and three dimensional computational analysis for Mach number ranging from 0.5 of 3. Also it is designed to fit successfully for 0.3m and 0.6m test section wind tunnel to obtain reasonably merging solution with theoretical and experimental results.

Konvergencija rezultata transoničnih aerotunelskih ispitivanja standardnog modela AGARD-B

AGARD-B is a widely-used configuration of a standard wind tunnel model. Beside its originally intended application for correlation of data from supersonic wind tunnel facilities, it was tested in a wide range of Mach numbers and, more recently, used for assessment of wall interference effects, validation of computational fluid dynamics codes and validation of new model production technologies. The researchers and wind tunnel test engineers would, naturally, like to know the "true" aerodynamic characteristics of this model, for comparison with their own work. Obviously, such data do not exist, but an estimate can be made of the dispersion of test results from various sources and of the probable "mean" values of the aerodynamic coefficients. To this end, comparable transonic test results for the AGARD-B model at Mach numbers 0.77, Mach 1.0 and Mach 1.17 from six wind tunnels were analyzed and average values and dispersions of the aerodynamic coefficients were computed.

Experimental and Numerical Investigation on Nozzle Exit Shape and Plume for Aft-Boom Mitigation of Supersonic Transport

This paper describes effects of nozzle geometry and a jet plume on the near-field signature and the thrust performance to explore concepts of aft-boom mitigation for future supersonic transport. Numerical analyses of four kinds of convergent-divergent nozzles were conducted to investigate the sensitivity of the near-field signature and the thrust. The results showed that reducing expansion region at the boat-tail is important to reduce both the peak pressure of near-field signature and the nacelle drag. A nozzle shape, which has an inclined exit shape without a boat-tail portion at lower side, demonstrated a good potential to reduce aft-boom. Supersonic wind tunnel tests using a jet with high-pressure air were also performed to validate the numerical analyses and the results showed good agreement with numerical analyses.

Experimental study of the laminar-turbulent transition on models of wings with subsonic and supersonic leading edge at M = 2

The paper is devoted to an experimental study of laminar-turbulent transition in supersonic boundary layers on swept wings at Mach 2. The experiments were fulfilled in the low nose supersonic wind tunnel T-325 of Khristianovich Institute of Theoretical and Applied Mechanics on the models with supersonic or subsonic leading edges. The transition location was determined using a constant temperature hot-wire anemometer. while experiments for the case of a supersonic leading edge are a continuation of previous studies, then for the case of a subsonic leading edge, the first data have been obtained that show the possibility of carrying out studies on the development of unstable perturbations. It was obtained that transition take place parallel to the leading edge for both models. The destabilizing effect of the surface curvature of the model on the position of the laminar-turbulent transition for the case of a supersonic leading edge was confirmed. This experimental data is the room for further investigation.

Free-stream preserving linear-upwind and WENO schemes on curvilinear grids

Applying high-order finite-difference schemes, like the extensively used linear-upwind or WENO schemes, to curvilinear grids can be problematic. If the scheme doesn't satisfy the geometric conservation law, the geometrically induced error from grid Jacobian and metrics evaluation can pollute the flow field, and degrade the accuracy or cause the simulation failure even when uniform flow imposed, i.e. free-stream preserving problem. In order to address this issue, a method for general linear-upwind and WENO schemes preserving free-stream on stationary curvilinear grids is proposed. Following Lax-Friedrichs splitting, this method rewrites the numerical flux into a central term, which achieves free-stream preserving by using symmetrical conservative metric method, and a numerical dissipative term with a local difference form of conservative variables for neighboring grid-point pairs. In order to achieve free-stream preservation for the latter term, the local differences are modified to share the same Jacobian and metric terms evaluated by high order schemes. In addition, this method allows a simple hybridization switching between linear-upwind and WENO schemes is proposed for improving computational efficiency and reducing numerical dissipation. A number of testing cases including free-stream, isentropic vortex convection, double Mach reflection, flow past a cylinder and supersonic wind tunnel with a step are computed to verify the effectiveness of this method.

Fine structures of self-sustaining dual jets in supersonic crossflow

The interaction between self-sustaining dual jets and supersonic laminar boundary layer is experimentally studied using the nanoparticle-based planar laser scattering in a Ma 2.95 supersonic low-turbulence wind tunnel for the first time. The typical fine flow structures, such as separation shock, “W” bow shock, barrel shock, horseshoe vortex, large-scale vortex structures and counter-rotating vortex pairs, are clearly identified by streamwise and spanwise flowfield images. The curves of jet penetration depth are fitted at the momentum ratio of 0.55 between self-sustaining dual jets and supersonic crossflow. The results play a significant role in further and comprehensive exploration of the active flow control for a supersonic flow.

Improved representation of destructive spacecraft re-entry from analysis of high enthalpy wind tunnel tests of spacecraft and equipment

The assessment of casualty risk in destructive re-entry has historically been performed purely by simulation using heating correlations, which only have verification on basic shapes, and estimated phenomenology for fragmentation. As the application of space debris mitigation requirements is expected to result in a higher number of re-entries, this has brought a stricter enforcement of the casualty risk guidelines to proposed missions. In turn, this has increased the interest in designing spacecraft to demise in re-entry in order to allow uncontrolled re-entries to be performed. Initial testing of spacecraft materials and basic structures has demonstrated that the destructive re-entry tools do not capture the correct physics to be able to assist design in a meaningful way, and therefore, some means to improve the representativeness of the tools is required. The current understanding of the phenomenology of the fragmentation and demise processes is limited. As a consequence, it is vital to perform appropriate tests in order to improve the capability of the tools to assist the design process. To this end, a set of destructive tests on spacecraft materials, structures and components has been performed in an arc-heated supersonic wind tunnel. These tests include the first destructive wind tunnel tests ever performed on a complete nano-satellite and a reaction wheel. From these tests, it has been determined that the failure of aluminium structures is highly dependent upon the behaviour of the protective metal oxide layer, and that this can be catastrophic in nature. The tests on the nano-satellite have shown that the structure can be supported by stainless steel spacers between the electronics cards, and that glass fibre reinforced plastic PCBs are more resistant to melting than had been anticipated. The reaction wheel test has shown that the connections between parts are critical to the fragmentation and demise processes, as the glued housing separates quickly, well before melt temperature is reached at the joint. It has also demonstrated the importance of radiative cooling, as the flywheel and ball-bearing unit have survived a test at over 800 kW/m2 with little damage.

Effects of steady and pulsed discharge arcs on shock wave control in Mach 2.5 flow

The effects of the steady and pulsed arc discharge on the oblique shock wave control are explored through an experimental study in a supersonic wind tunnel with the maximum design Mach number of 2.5. The oblique shock is formed by a compression ramp with an angle of 7 degree mounted on the floor of the wind tunnel. Four electrodes are placed upstream of the compression ramp and spaced equally in the spanwise direction to generate the discharge arcs. The steady discharge is generated between the electrodes and downstream ramp corner, forming streamwise arcs. The arc length can be modulated between 10 mm and 40 mm by moving the electrodes in the streamwise direction. With a High Frequency Switch (HFS), the pulsed discharge is achievable with a frequency range from 5 kHz to 50 kHz. The pulsed discharge is generated between two adjacent electrodes, forming so-called transversal arcs, which are blown downstream by the main flow. Results show that the steady discharge arcs act like a uniformly distributed conductor. With an increase of the arc length, the arc power increases, and the weakening effect on the shock is enhanced. For the pulsed discharge arcs, the shorter arcs are observed with higher discharge frequency, which implies a lower arc power. Overall, the weakening effect of the steady arcs on the shock is more effective with shock strength reduced by 4%, compared to the pulsed ones with only 0.35% reduction.

Effect of hypersonic flow chemical composition on the oxidation behavior of a super-strong UHTC

A super-strong ZrB2 ceramic containing WC and SiC was tested in a supersonic arc-jet wind tunnel by exposure to flows with two chemical compositions, simulated air or pure nitrogen, at temperatures of 2650 and 2800 K. Temperature jumps of 500–600 K were observed in both environments at constant flow conditions. SEM analyses revealed that oxygen in the high-enthalpy flow retards the material consumption owing to the formation of partially protective glass. Then, it appears that even 5 vol% of W-compounds is sufficient to modify the subscale oxide configuration and form a cell-like multi-layered architecture typical of tungsten, rather than that typical of ZrB2.

Peculiarities of low-Reynolds-number supersonic flows in long microchannel

The characteristics of low-Reynolds-number supersonic flows in a long microchannel having a rectangular cross section are investigated computationally. The channel is composed of a Laval nozzle and a straight duct. The design Mach number of the nozzle is 2.0 and the Reynolds number calculated at the nozzle exit is 3100. The length of the straight duct is changed from 2 to 18 h, where h is the duct height. In the computations, the Navier–Stokes equations are numerically solved. The computational code is validated using the experimental data measured by the laser-induced fluorescence (LIF) technique. The computational results demonstrate that neither a normal shock wave nor a pseudo-shock wave, which corresponds to the starting shock wave in a supersonic wind tunnel, appears in microchannel flows. Namely, a low-Reynolds-number supersonic flow is created in a channel without the starting shock wave passing along the duct, although it has been believed that a supersonic internal flow should have been formed through the starting shock wave. In addition, it is found that the microchannel flow changes gradually its supersonic state with the channel length under an underexpanded condition, although a starting shock wave for high-Reynolds-number flows suddenly appears in a channel just as its length exceeds a certain specific length. Such unexpected phenomena originate from the peculiarity that the low-Reynolds-number flows can expand (accelerate) along a straight duct at supersonic speeds, although the high-Reynolds-number flows cannot.

Accurate 2-D Modelling of Transonic Compressor Cascade Aerodynamics

Modern aeronautic fans are characterised by a transonic flow regime near the blade tip. Transonic cascades enable higher pressure ratios by a complex system of shockwaves arising across the blade passage, which has to be correctly reproduced in order to predict the performance and the operative range. In this paper, we present an accurate two-dimensional numerical modelling of the ARL-SL19 transonic compressor cascade. A large series of data from experimental tests in supersonic wind tunnel facilities has been used to validate a computational fluid dynamic model, in which the choice of turbulence closure resulted critical for an accurate reproduction of shockwave-boundary layer interaction. The model has been subsequently employed to carry out a parametric study in order to assess the influence of main flow variables (inlet Mach number, static pressure ratio) and geometric parameters (solidity) on the shockwave pattern and exit status. The main objectives of the present work are to perform a parametric study for investigating the effects of the abovementioned variables on the cascade performance, in terms of total-pressure loss coefficient, and on the shockwave pattern and to provide a quite large series of data useful for a preliminary design of a transonic compressor rotor section. After deriving the relation between inlet and exit quantities, peculiar to transonic compressors, exit Mach number, mean exit flow angle and total-pressure loss coefficient have been examined for a variety of boundary conditions and parametrically linked to inlet variables. Flow visualisation has been used to describe the shock-wave pattern as a function of the static pressure ratio. Finally, the influence of cascade solidity has been examined, showing a potential reduction of total-pressure loss coefficient by employing a higher solidity, due to a significant modification of shockwave system across the cascade.

HB-2 high-velocity correlation model at high angles of attack in supersonic wind tunnel tests

Responding to a need for experimental data on a standard wind tunnel model at high angles of attack in the supersonic speed range, and in the absence of suitable reference data, a series of tests of two HB-2 standard models of different sizes was performed in the T-38 trisonic wind tunnel of Vojnotehnički Institut (VTI), in the Mach number range 1.5–4.0, at angles of attack up to +30°. Tests were performed at relatively high Reynolds numbers of 2.2 millions to 4.5 millions (based on model forebody diameter). Results were compared with available low angle of attack data from other facilities, and, as a good agreement was found, it was assumed that, by implication, the obtained high angle of attack results were valid as well. Therefore, the results can be used as a reference database for the HB-2 model at high angles of attack in the supersonic speed range, which was not available before. The results are presented in comparison with available reference data, but also contain data for some Mach numbers not given in other publications.