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Experiments were performed within Rutgers University’s Mach 3.4 wind tunnel to visualize the unsteady flow over a wall-mounted hemisphere using high-speed schlieren imaging and stereoscopic particl...

Results of experimental and numerical investigations of a supersonic flow of cylindrical models aligned at a zero angle of attack with frontal inserts made of cellular porous nickel are reported. The experiments are performed in a supersonic wind tunnel at Mach numbers $$\text{M}_{\infty } = 4.85$$ and 7.00 and unit Reynolds numbers $$\text{Re}_1 = 2.7\cdot 10^6$$ and $$1.5\cdot 10^6$$ m $$^{ - 1}$$ , respectively. Numerical simulations with the use of a ring-shaped skeletal model of the porous material are also performed. A possibility of drag control is studied for two thermal methods: external heating of the porous insert and internal heating of the insert by a glow discharge. The mechanisms of the thermal action and the efficiency of the thermal methods of drag control are analyzed.

This paper presents the results of studying the features of the development of weak shock waves generated by a twodimensional roughness on the wall of the working part of a supersonic wind tunnel in a free flow at a Mach number of 2.5. The measurements were performed with a constant resistance thermoanemometer. It is shown that a twodimensional sticker induces weak shock waves into the free flow. They cause distortion of the average flow, the shape of which corresponds to the N-wave. High-intensity pulsations were recorded in the region of passage of a pair of weak shock waves. With an increase in the unit Reynolds number, the level of distortions of the average flow remains practically constant, but an increase in nonstationary disturbances is observed. It was found that the greatest increase in pulsations caused by Mach waves is observed in the area of the maximum gradient of the average flow. It is found that an increase in the number ReLleads to an expansion of the frequency range of unstable disturbances generated by a pair of weak shock waves.

Understanding the flow characteristics over a double compression ramp is crucial for high-speed vehicle design. Leading edge bluntness is a key factor influencing the formation of a separation region on a double compression ramp flow. In the present study, the effect of bluntness on a double compression ramp is investigated experimentally at a nominal Mach number of 4. The test model has 13° and 40° inclinations with respect to the freestream. Five different levels of leading-edge radius, varying from 0.0 to 2.0 mm, were subjected to supersonic wind tunnel tests. Shadowgraph and infrared thermography techniques were employed to visualize the flow features of the double ramp model. Measurements of surface heat-transfer along the centerline of the test model were obtained from the acquired infrared images. It is shown that the leading-edge radius alters the separation characteristics as well as the surface heat-transfer. Possible reasons for such flow characteristics are discussed.

Results of experimental and numerical modeling of a supersonic flow around a cylinder with a frontal gas-permeable high-porosity insert aligned at different angles of attack are presented. The experiments are performed in a supersonic wind tunnel at the Mach number $$\mbox{M}_{\infty }=7$$ and unit Reynolds number $$\mbox{Re}_{1}=1.5 \cdot 10^6$$ m $$^{ - 1}$$ in the range of the angles of attack $$0–25^\circ$$ . The numerical simulations are performed by means of solving three-dimensional Reynolds-averaged Navier–Stokes equations with the use of a three-dimensional ring skeleton model of the porous material. The drag and lift coefficients for a cylinder with a 95% porosity and pore diameter of 2 mm are obtained for different values of the insert length and angle of attack.

The transient response of a plate and a cavity is investigated in a supersonic wind tunnel start experiment where the freestream flow inside the test section reaches turbulent flow at Mach 2. Experimentally measured plate displacement time history shows flutter onset, transition to limit cycle oscillation, and stabilization at a static deformed state during the 30 s run. To analyze and interpret the measured plate response, a fully coupled aero-thermal-acousto-elastic analysis is carried out. A theoretical–computational model is formulated with a nonlinear structural plate model, acoustic pressure equation for the stationary fluid in a cavity, and the first-order Piston Theory aerodynamics. A linear stability analysis is performed that includes the nonlinear added stiffness due to an initial deformation to investigate the combined effects of freestream coupling and temperature differential on system stability. Also, direct time integration of the nonlinear fluid structural equations of motion is performed using experimentally measured flow parameters as inputs. All stability transitions are captured using the theoretical model with good agreement with experiment for transitions from no flutter to flutter/limit cycle oscillations (LCO) although the theoretical LCO amplitude is approximately $$50\%$$ larger than measured. The system’s sensitivity to cavity coupling, temperature differential, thickness calibration, static pressure differential, and turbulent pressure fluctuations are investigated. Lastly, snap-through buckling analyses in response to periodic and quasi-static excitations are conducted.

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 follow...

To understand time-dependent aerodynamic loads on single and tandem (fore/aft) slender cylinders translated out into the supersonic freestream from a cavity at Mach 1.5, the normal force and pitchi...

Extensive wind-tunnel tests were performed on a Y-shaped diverterless supersonic inlet (DSI). All tests were conducted at a free stream Mach number of M∞=1.65, the design Mach number for this inlet...

This study aims to develop a simplified approach for the design of high-speed wind tunnels, operating within the range of Mach 1-4. The study is conducted in two stages; design stage-I consists of the basic parts of a high-speed wind tunnel, i.e. test section and nozzle throat sizes along with the calculations of the minimum pressure ratios necessary for the starting and running of the wind tunnel. Design stage-II is the aerodynamic design of the inlet nozzle contour with the focus on the divergent nozzle section only. In this stage, the fundamental method of characteristics (MOC) is employed by using region-to-region solution approach. Due to the long and complicated procedure of classical MOC, a simpler approach of “table of inclination” is proposed. This inclination table approach is simple in use, highly precise, and efficient which reduces more than 75 % of the calculation time of the MOC. Validation of the nozzle contour design is performed with the Area-Mach ratio relation followed by the results verification obtained from the MATLAB programming. The contour design results of the study are highly accurate (less than 1 % error) which manifests the future possible significance of the inclination table approach for supersonic wind tunnel designs.

An image-based flow velocimetry technique coupling pulse-burst laser technology and the cross-correlation Doppler global velocimetry (CC-DGV) technique is presented as a faster variant of the CC-DG...

A control method is developed to suppress pressure oscillations induced by supersonic cavity flow using high-speed upstream injection. The injection is generated with a large blowing coefficient th...

The shock wave boundary layer interactions (SWBLIs) at compression ramps (ramp angle α=20o -30o ) are studied at Ma=2.0 and under two Reynolds numbers (Re1=18,600 and Re2=35,600, Re based on boundary layer thickness). High-speed schlieren operating at 20 kHz is used as the flow diagnostics. The flow structures in the compression ramp SWBLIs, including the shock wave, interaction region and induced turbulent region over the ramp surface, are discussed. Their variation under increasing ramp angle and Reynolds number are further examined. The low-frequency shock wave oscillations are also studied through tracking the shock wave motion. A larger ramp angle increases the spectral intensity of the shock wave’s low-frequency unsteadiness, while increasing the Reynolds number results in a lower peak frequency for the separation and reattachment shock waves.

In supersonic wind tunnels, the airflow at the exit of a convergent-divergent nozzle is affected by the connection between the nozzle and test section, because the connection is a source of disturbance for supersonic flow and the source of disturbance generated by this disturbance propagates downstream. In order to avoid the disturbance, the test can only be carried out in the rhombus area. However, for the supersonic nozzle, the rhombus region is small, limiting the size and attitude angle of the test model. An integrated supersonic nozzle is a nozzle and a test section as a whole, which is designed to weaken or eliminate the disturbance. The inviscid contour of the supersonic nozzle is based on the method of characteristics. A new curve is formed by the smooth connection between the inviscid contour and test section, and the boundary layer is corrected for the overall curve. Integrated supersonic nozzles with Mach number 1.5 and 2 are designed, which are based on this method. The flow field is validated by numerical and experimental results. The results of the study highlight the importance of the connection about the nozzle outlet and test section. They clearly show that the wave system does not exist at the exit of the supersonic nozzle, and the flow field is uniform throughout the test section. Keywords: Boundary layer, Disturbance, Flow field, Integrated supersonic nozzle, Supersonic wind tunnels, The rhombus region

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. Keywords: Base pressure, Experimental aerodynamics, High angle of attack, Standard model, Wind tunnel

Schlieren visualizations of flowfield structures of supersonic film cooling in a backward-facing step were conducted in a Mach 2.95 suction wind tunnel with the film coolant tangentially ejected th...

The base flow of an axisymmetric generic space launcher model is investigated experimentally by means of particle image velocimetry and dynamic pressure measurements at a Mach number of 0.76 and a Reynolds number of 1.5 × 106, based on the main body diameter. The flow separation at the end of the main body forms a highly dynamic recirculation region with strong pressure fluctuations on the reattaching surface. The time-averaged reattachment on the rear sting is at 1.05 main body diameters downstream of the step. This work investigates the application of passive flow control devices for their potential of reducing the loads on the space launcher’s nozzle. It is shown that rectangular or circular grooves at the end of the main body force enhanced mixing in the separated shear layer, leading to a reduction of the reattachment length of 55%. Additionally, the fluctuations of the reattachment are significantly reduced, which results in lower-pressure fluctuations and thus reduced dynamic loads.

Upon its development and initial characterization, the supersonic variant of the nanoscale thermal anemometry probe (S-NSTAP) was deployed in a supersonic wind tunnel facility, where both freestream and boundary layer measurements were obtained at $$M_\infty =2$$ . The low operating stagnation pressures generated reliable data, where the effects of Reynolds number, Mach number and overheat ratio on the sensor’s heat transfer were investigated in detail. The performance of the S-NSTAP was also compared to that of a conventional cylindrical hot-wire and the S-NSTAP was shown to exhibit unparalleled temporal resolution ( $$\sim $$ 300 kHz). The mass flux sensitivity coefficient of both hot-wires was further computed and appeared to vary between probes, yielding a coefficient twice as large for the conventional probe than for the S-NSTAP. The experimental data obtained from both hot-wires were also compared, via spectral analysis and turbulence statistics, to the results of a numerically modelled turbulent boundary layer.

Experiments were performed within Rutgers University’s supersonic wind tunnel to measure the influence of off-axis laser energy deposition on the flow field about an ogive cylinder at a freestream Mach number of 3.4. Perturbation of the flow field was accomplished using an infrared laser source, focused to a point ahead of the ogive cylinder. Stereoscopic particle image velocimetry measurements were performed to quantify the effects of energy deposition on the flow field at discrete time delays following the generation of the spark. The SPIV results showed a measurable change in streamwise velocity downstream of ogive’s shock that appears to be dependent on proximity of the initial spark to the ogive’s surface. In contrast, the spark was shown to have little influence on the vertical velocity component at early times. Data corresponding to later times showed the passage of an induced jet through the flow field. The jet rotated about its axis while passing through the shock structure, in agreement with previous qualitative imaging. These results demonstrate the feasibility of using SPIV to investigate the influence of laser energy deposition on the flow field.

Streamwise-coherent structures were observed in schlieren images of a Mach 2.5 flow in an empty supersonic wind tunnel with a rectangular cross section. These features are studied using Reynolds-averaged Navier–Stokes computations in combination with wind-tunnel experiments. The structures are identified as regions of streamwise vorticity embedded in the sidewall boundary layers. These vortices locally perturb the sidewall boundary layers, and they can increase their thickness by as much as 37%. The vortices are caused by a region of separation upstream of the nozzle where there is a sharp geometry change, which is typical in supersonic wind tunnels with interchangeable nozzle blocks. Despite originating in the corners, the vortices are transported by secondary flows in the sidewall boundary layers so they end up near the tunnel center height, well away from any corners. The successful elimination of these sidewall vortices from the flow is achieved by replacing the sharp corner with a more rounded geometry so that the flow here remains attached.

The aerodynamic characteristics of launch vehicles separable elements, which are cylindrical and conical shells, in supersonic airflow are presented in this paper. Numerical simulation of the flow around shells in the plane of their symmetry is performed for flow Mach numbers from 2.0 to 4.0 using an open-source software package OpenFoam. We got the aerodynamic coefficients of the axial, normal forces and pitch moment, and the structure of the flow around shells. We compared the numerical simulation results with the results of experimental studies conducted in the BMSTU supersonic wind tunnel. The complex nature of the flow around shells with the formation of shock waves, areas of flow separation, and circulation flow are established. We compared the aerodynamic characteristics of cylindrical and conical shells with the characteristics of a solid cone and cylinder, and with the characteristics of rectangular and triangular plates. We found that the flow around the shells (except for the hollow cone) is accompanied by a through flow, the aerodynamic force is created by all the shell surfaces. The flow around the hollow cone is accompanied by the formation of a stagnant area inside the cavity. Features in the flow structures are reflected in the shells’ aerodynamic characteristics, which differ from the solid bodies and plates characteristics.

This study was carried out after the initial divergent nozzle contours of the high-speed wind tunnel inlet nozzle were designed in [1]. The goal of the study is to design a suitable and simplified inlet section (settling chamber, C–D nozzle, and test section) of the supersonic wind tunnel prototype. At first, an optimized design for the settling chamber was created after which the settling chamber could be used for any planned Mach configuration (1.5≤M≤3). In the next step, three different designs for the inlet section of the wind tunnel were created and a suitable design for the future manufacturing of the inlet section was selected with the help of the numerical simulations.

The peculiarities of supersonic unsteady flows forming in the processes of an impulse starting of a wind tunnel including a fore-chamber, a nozzle, a diffuser, and an exhaust tank are considered. The fore-chamber is separated from the flow duct by a thin breakable diaphragm. Before the wind tunnel starting, the gas contained in the exhaust tank is pumped out down to a very small pressure, and then the high-pressured working gas is fed into the fore-chamber. Upon reaching some value of this pressure, the diaphragm “instantaneously” ruptures, and the working gas starts exhausting through the nozzle: a rapid unsteady process of the wind tunnel starting arises. The numerical simulation of the flows forming at the impulse starting of the simplest experimental gas-dynamic facility has been carried out; the facility is arranged with a nozzle forming at its exit a two-dimensional supersonic flow with the Mach number of 2.9, and with replaceable diffusers having different relative areas of the throat. The numerical computations of two-dimensional unsteady flows have been carried out using the Reynolds-averaged Navier—Stokes equations and the SST k-ω turbulence model. The flow patterns computed numerically are compared with the data of the optical visualization of the flow obtained in the experimental gas-dynamic facility in the process of its starting.

Unsteady flow characteristics of a normal shock wave, a lambda foot, and a separated turbulent boundary layer are investigated within a unique test section with supersonic inlet flow. The supersonic wind tunnel facility, containing this test section, provides a Mach number of approximately 1.54 at the test section entrance. Digitized shadowgraph flow visualization data are employed to visualize shock wave structure within the test section. These data are analyzed to determine shock wave unsteadiness characteristics, including grayscale spectral energy variations with frequency, as well as time and space correlations, which give coherence and time lag properties associated with perturbations associated with different flow regions. Results illustrate the complexity and unsteadiness of shock-wave-boundary-layer-interactions, including event frequencies from grayscale spectral energy distributions determined using a Lagrangian approach applied to shock wave location, and by grayscale spectral energy distributions determined using ensemble-averaging applied to multiple closely-located stationary pixel locations. Auto-correlation function results and two-point correlation functions (in the form of magnitude squared coherence) quantify the time-scales of periodic events, as well as the coherence of flow perturbations associated with different locations, over a range of frequencies. Associated time lag data provide information on the originating location of perturbation events, as well as the propagation direction and event sequence associated with different flow locations. Additional insight into spatial variations of time lag and flow coherence is provided by application of magnitude squared coherence analysis to multiple locations, relative to a single location associated with the normal shock wave.

A deformable shape-memory alloy (SMA) bump is introduced to control the succeeding cowl-shock/boundary-layer interactions in a supersonic inlet with an operating Mach range of 2.0–3.8. The deformat...

An increased number of requirements for supersonic wind tunnel tests at high angles of attack (AoA) opened the problem of the validation of a high-AoA wind tunnel setup in the Experimental Aerodynamics Laboratory of VTI (Military Technical Institute) in Belgrade. Tests of the supersonic-hypersonic HB-2 models were performed at Mach numbers 1.5 to 4 at AoA up to +30°. A new bent sting was used as the model support, shifting the AoA range of the wind tunnel model support. Obtained data were correlated with results of previous tests performed in the same wind tunnel in the lower-angle-of-attack range and with available test results from other wind tunnel facilities. The results are found to be useful as supersonic test cases for both experimental and numerical researchers

The flow structures of the inlet in continuously variable Mach number inflow are studied experimentally to evaluate accurately the performance of scramjet inlet during acceleration/deceleration in ground experiments. To break the limitation in which majority of supersonic wind tunnels operate at a fixed Mach number condition, a continuously variable Mach number wind tunnel has been built. The running Mach number of the wind tunnel varies continuously from 2.0 to 3.0 under suction mode. The experiments of the inlet at fixed and variable Mach number inflows are conducted successively in the continuously variable Mach number wind tunnel. In the fixed Mach number experiment, the typical flow structures of the inlet under the unstart/start conditions are captured. The pressure distributions on the lower wall of the inlet, which were obtained from the experiments, are consistent with numerical ones. Thus, the effectiveness and accuracy of the experiments and numerical simulations were verified. In the continuously variable Mach number wind tunnel experiment, we found the inlet suffers from four flow structures. With the exception of typical unstart/start flow structures, the inlet undergoes pseudo unstart/start conditions under the influence of the wind tunnel unstart. By analyzing the pressure and Schlieren result, variations in the flow structures are observed in detail, which can provide useful references for the subsequent experiments in continuously variable Mach number wind tunnels. In particular, pseudo unstart/start conditions must be distinguished from real ones in the case wrong experimental data are obtained in the following experiments.

Schlieren and shadowgraph visualizations of flowfields in a typical configuration of supersonic film cooling in a backward-facing slot were conducted in a Mach 2.95 continuous-suction wind tunnel, ...

This study examined the ignition and flameholding effects of the quasi-direct-current discharge on a directly injected hydrocarbon fuel (ethylene) in a supersonic combustion chamber without mechani...

The mechanism of oblique shock wave control by surface arc discharge plasma is explored through a combined numerical and experimental study. The experiments are conducted in a supersonic wind tunne...

An experimental and numerical study was conducted to examine the effects of internal surface curvature and leading-edge angle on the shock waves and steady flow fields produced by axisymmetric ring wedges. Test models with leading-edge-radius-normalised internal radii of curvature of $R_{c}=\{1,1.5,2\}$ and leading-edge angles of $\unicode[STIX]{x1D6FC}=\{0^{\circ },4^{\circ },8^{\circ }\}$ were manufactured and tested. Experimental shadowgraph and schlieren results were obtained for Mach numbers ranging from 2.8 to 3.6 using a blowdown supersonic wind tunnel with accompanying numerical results for additional insight. The higher the internal surface curvature and leading-edge angle, the greater the flow fields were impacted. As a result, steeper compression waves were formed, thus curving the shock wave more noticeably. The internal surface curvature and leading-edge angle were both found to have an effect on the trailing-edge expansion fans. This altered the shape of downstream shock wave structures. The highest curvature models produced steady double reflection patterns due to the imposed internal surface curvature. The effects of conical and curved internal surfaces were explored for the presence of flow-normal curvature and the curving of the attached shock waves.

This investigation concerns the design, build, and testing of direct-measuring skin friction sensors capable of performing in sustained hypersonic flow and detecting transition. A multistep approac...

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.

A Schlieren flow visualization system has been developed for the study of the injection of jets in a supersonic cross-flow, such as used in supersonic-combustion ramjets. Both single-jet in-jec¬tion and cascaded, dual, tandem, injection have been considered in a M∞ supersonic cross-flow in a continuous supersonic wind tunnel. The study included the development of a high-performance, but cheap, light source for the Schlieren system. An earlier study has demonstrated that LEDs are suitable light sources for Schlieren imaging of supersonic flow, though LEDs have to be pushed to their limits in terms of requirements of small pulse width and high optical output. Recent developments in power Ver¬tical Cavity Surface Emitting Lasers (VCSELs) have resulted in a far more efficient light source. Small emission angles, high power and nanosecond pulses enable Schlieren imaging of un¬steady flows in the high-Mach-number regime. Though the design of drive electronics for VCSELs and LEDS are very similar, in order to obtain (much) smaller pulse widths, the de-sign needs an upgrade. The present paper demonstrates Schlieren imaging using an in-house de¬veloped driver emitting pulses of (effective) width down to 12.9 ns, which is at least 10 times faster than required for Mach 1.6 flows. The Schlieren set-up has been used for the investigation of the injection of dual, tandem, un-der-expanded sonic jets into a supersonic cross-flow. The effects of tandem dual injection on the mixing is considered as function of the jet-to-free-stream momentum ratio J and the di-men¬sionless distance S between the two jets. The Schlieren images have been used to extract the location of the upper boundary of the jet shear layer. In order to compare mixing cha¬rac-ter¬istics a three-parameter power-law least-squares fit has been developed to describe the time-averaged location of the upper boundary of the jet shear layer. The fit is able to describe the upper shear layer; however, some data strings show a dip in the penetration height, which sug¬gests that an alternative description should be defined. The calculated averages of the penetration height of the upper boundary of the jet shear layer, de¬fined as a measure for the mixing performance of the tandem-jet system, show that, for each value of the momentum ratio, there is an optimal distance between the injection orifices. It is ob¬served that for a higher momentum ratio the optimal distance between the jets increases. Spe¬cific spacings can improve the average penetration height of the jet shear layer by 30% com¬pared to that of single injection at the same momentum ratio.

This paper aims to design a small-scale, compact supersonic wind tunnel nozzle using the method of characteristics and perform computational fluid dynamics (CFD) simulations to validate the design. The wind tunnel is designed to have a test section of 7cm × 7cm and a test section Mach number of 2.0. Simulations were performed for both 3D inviscid and viscous cases and it shows that for both cases, the obtained Mach number conforms with the design requirements. A further study on the effects of different pressure ratios, towards the position of shock waves, shows that an increase in the ratio would result the shock waves to shift downstream.

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.

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.

In supersonic flows, the separation in streamwise corners is a significant and widely encountered problem which can not be reliably predicted with the numerical methods commonly used in industry. The few previous studies on this topic have suggested conflicting corner flow topologies. Experiments of supersonic flow are typically conducted in wind tunnels with rectangular cross-sections, which use either a symmetric (full) or asymmetric (half-liner) nozzle configuration. However, the effect of the nozzle arrangement on the corner flow itself is not known. This paper examines the influence of nozzle geometry on the corner regions of a Mach 2.5 flow using a joint experimental-computational approach. The full setup and half-liner configuration are shown to produce different corner flow structures. The corner regions of the full setup and top corners of the half-liner exhibit thin sidewall boundary layers and a single primary vortex on the floor or ceiling. Meanwhile, the bottom corners of the half-liner configuration contain thick sidewall boundary layers and a counter-rotating vortex pair. Considerable vertical velocities are measured within the sidewall boundary layers. These are directed towards the tunnel centre-height for the full setup and downwards with the half-liner. The differences in sidewall cross flows between the two nozzle arrangements are likely due to distinct pressure distributions in the nozzle, where the secondary flows are set up. Measurements suggest that these nozzle-dependent transverse flows are responsible for the differences in corner flowfield between the two configurations. The proposed mechanism also explains observed differences in corner flow topology between previous studies in the literature; nozzle geometry therefore appears to be the dominant influence on corner flows in supersonic wind tunnels.

Design of two-dimensional supersonic diffusers as a part of thewind tunnelis investigated in this paper. A methodology based on the mixture of try-and-error method and optimization algorithm is developed to handle thedesign problem. In the first design step, using try-and-error approach, the main parameters related to the geometry of diffuser such as length, angle andarea ratiobetween the throat and the outlet are determined assuming a diffuser with linear walls. Thedesign criterionin this step is the fact that the shock wave should be created near the diffuser throat in order to benefit from the maximum efficiency of diffuser. In the second design step, considering the optimization methodology, it is tried to improve theoptimum designobtained in the first step by modifying the wall point locations and keeping the rest of the geometry fixed. Hence, an optimization problemis defined to find the best curve for the diffuser wall instead of the a linear one used the in the first step. Theobjective functionof this problem is to minimize the output Mach number usingGenetic Algorithm(GA). Thefluid flowis evaluated using theEuler equationsin the conservative form where the Streamline-Upwind/Petrov–Galerkin (SUPG) finite element scheme is used to discretize theflow equations. In order to capture the flow solution around the shock waves accurately, theadaptive mesh refinementtechnique is coupled to the flow solution. The demonstrated results show the efficiency of the proposed method for designing supersonic diffusers

The nonlinear response of a compressible boundary layer to unsteady free-stream vortical fluctuations of the convected-gust type is investigated theoretically and numerically. The free-stream Mach number is assumed to be of O(1) and the effects of compressibility, including aerodynamic heating and heat transfer at the wall, are taken into account. Attention is focused on low-frequency perturbations, which induce strong streamwise-elongated components of the boundary-layer disturbances, known as streaks or Klebanoff modes. The amplitude of the disturbances is intense enough for nonlinear interactions to occur within the boundary layer. The generation and nonlinear evolution of the streaks, which acquire an O(1) magnitude, are described on a self-consistent and first-principle basis using the mathematical framework of the nonlinear unsteady compressible boundary-region equations, which are derived herein for the first time. The free-stream flow is studied by including the boundary-layer displacement effect and the solution is matched asymptotically with the boundary-layer flow. The nonlinear interactions inside the boundary layer drive an unsteady two-dimensional flow of acoustic nature in the outer inviscid region through the displacement effect. A close analogy with the flow over a thin oscillating airfoil is exploited to find analytical solutions. This analogy has been widely employed to investigate steady flows over boundary layers, but is considered herein for the first time for unsteady boundary layers. In the subsonic regime the perturbation is felt from the plate in all directions, while at supersonic speeds the disturbance only propagates within the dihedron defined by the Mach line. Numerical computations are performed for carefully chosen parameters that characterize three practical applications: turbomachinery systems, supersonic flight conditions and wind tunnel experiments. The results show that nonlinearity plays a marked stabilizing role on the velocity and temperature streaks, and this is found to be the case for low-disturbance environments such as flight conditions. Increasing the free-stream Mach number inhibits the kinematic fluctuations but enhances the thermal streaks, relative to the free-stream velocity and temperature respectively, and the overall effect of nonlinearity becomes weaker. An abrupt deviation of the nonlinear solution from the linear one is observed in the case pertaining to a supersonic wind tunnel. Large-amplitude thermal streaks and the strong abrupt stabilizing effect of nonlinearity are two new features of supersonic flows. The present study provides an accurate signature of nonlinear streaks in compressible boundary layers, which is indispensable for the secondary instability analysis of unsteady streaky boundary-layer flows.\ud

A combined experimental and computational study is conducted of heat-shield ablation from a scaled model of NASA’s Orion Multi-Purpose Crew Vehicle in a Mach 5 wind tunnel. The ablating heat shield...

Intake buzz was initiated on a typical two-dimensional supersonic air-intake model at various supersonic Mach numbers up to 3 by gradually changing the back pressure from supercritical to subcritical operating condition in a wind tunnel. Schlieren pictures from a still camera and a high-speed camera were recorded. Analyses of individual high-speed images of the unvented intake were carried out to locate the time-dependent positions and velocities of the ramp shock around the cowl lip. The displacements of the shock indicate sinusoidal oscillations with dominant frequency of 102.4 Hz, close to that obtained from unsteady pressure measurements. Phase trajectories of shock position based on image analyses indicate that the shock oscillations have limit cycle type oscillation, typical of nonlinear dynamic systems. Natural ventilation of the intake was found to be extremely effective in increasing the total pressure recovery, suppress buzz oscillations and in delaying the onset of buzz by preventing the upstream propagation of disturbances through passive bleeding of the internal boundary layer. Effectiveness of natural ventilation in control of air-intake buzz at Mach 3.0

A high-quality model validation experiment was performed in the NASA Langley Research Center Unitary Plan Wind Tunnel to assess the predictive accuracy of computational-fluid-dynamics models for a blunt-body supersonic retropropulsion configuration at freestream Mach numbers from 2.4 to 4.6. Static and fluctuating surface pressure data were acquired on a 5-in.-diam (127-mm-diam) test article with a forebody composed of a spherically blunted, 70 deg half-angle cone and a cylindrical aft body. One unpowered configuration with a smooth outer mold line was tested as well as three powered, forward-firing nozzle configurations: a centerline nozzle, three nozzles equally spaced around the forebody, and a combination with all four nozzles. A key objective of the experiment was the determination of experimental uncertainties from a range of sources such as random measurement variation, flowfield nonuniformity, and model/instrumentation uncertainties. This paper discusses 1) the design of the experiment to best cap...

The aerodynamic design of a rigid-flexible coupling profile is the decisive factor for the flow-field quality of a supersonic free jet wind tunnel nozzle, and its mechanic dynamic features are the key for engineering implementation of continuous Mach number regulations. To fulfill the requirements of a free jet inlet/engine compatibility test within a wide simulation envelop, both uniform flow-fields of continuous acceleration and deceleration are necessary. In this paper, the aerodynamic design methods of an expansion wall and machinery implementation plan for the half-flexible single jack nozzle were researched. The profile control in nozzle flexible plate design was studied with a rigid-flexible coupling method. Design and calculations were performed with the help of numerical simulation. The technique of axial free stretching of the flexible plate was used to improve the matching performance between the designed elasticity profile and the theoretical one, and the rigid-flexible coupling structure was calibrated by wind tunnel tests. Results indicate that the flexible plate aerodynamic design method used here is effective and feasible. Via rigid-flexible coupling design, the flexible plate agrees with the rigid body very well, and continuous Mach number changes can be achieved during the tests. The nozzle’s exit flow-field uniformity meets the requirements of China Military Standard (GJB).

The roll damping and Magnus dynamic stability coefficients, as well as total force and moment coefficients, were numerically computed at high angles of attack up to 90 deg for two generic fin-stabilized flight munitions at a supersonic Mach number of 2.49. The aerodynamic coefficients were computed via time-accurate Reynolds-averaged Navier–Stokes numerical methods, and they were compared with archival wind-tunnel data. Fair to excellent comparisons with the experiment were obtained for the coefficients for the full angle-of-attack range, except for the axial force, which was overpredicted at higher angles of attack. Numerical modeling parameter studies were conducted to include dependence on spin rate, grid resolution, temporal resolution, and other numerical convergence parameters.

The Aerodynamic Characterization Facility is a unique facility located at the University of Florida’s Research and Engineering Education Facility designed with the intent to study complex low Reynolds number, unsteady aerodynamic phenomena. This facility includes an open jet, open return wind tunnel specifically designed to operate in the low Reynolds number regime where current research efforts are being tasked for Micro Air Vehicle flight platforms. Specifically, the wind tunnel operates with freestream velocities ranging from nominally 0 to 22 m/s with 0.1 m/s resolution provided by a variable frequency drive utilized to control the tunnel. The test section entrance is 1.07 m2with a length of 4.6 m. Flow uniformity investigations at free stream velocities of 2 and 4 m/s demonstrate a uniform flow core throughout the test section of at least 50% of the 1.14 m2contraction exit. Hot wire anemometry investigations present turbulence intensities less than 0.22% for free stream velocities greater than 1 m/s. The facility is also equipped with a dynamic motion rig and active turbulence generator. The dynamic motion rig is utilized to investigate unsteady aerodynamics over dynamic kinematic motions similar to what one might find in small birds and insects. The active turbulence generator provides a means to introduce atmospheric-like turbulence into the wind tunnel to understand the coupling between turbulence and unsteady flow phenomena associated with bird and insect flight.

A theory based upon linearized governing equations is developed that describes the operation principles and design parameters for low-speed wind tunnels with longitudinal freestream oscillations. Existing measurements made in unsteady wind tunnels are shown to be consistent with the theory and targeted validation experiments performed in a variable-geometry blowdown-type wind tunnel revealed excellent correspondence with the theoretical results. In particular, the tunnel frequency bandwidth is proportional to the mean tunnel freestream velocity and inversely proportional to the test-section length and the square of the exit area to test-section area ratio. The acoustics equations reveal a “Helmholtz damping ratio” that is not only dependent on the tunnel geometry and exit area but also proportional to the freestream Mach number. At appreciable reduced frequencies, dynamic stall on the louver vanes increases pressure losses, thereby reducing the mean flow speeds. Varying the exit area results in louver-van...

The transient dynamic behaviour of supersonic disk-gap-band parachutes in a Mars entry environment involving fluid structure interactions is studied. Based on the multi-material Arbitrary Lagrange–Euler method, the coupling dynamic model between a viscous compressible fluid and a flexible large deformation structure of the parachute is solved. The inflation performance of a parachute with a fixed forebody under different flow conditions is analysed. The decelerating parameters of the parachute, including drag area, opening loads, and coefficients, are obtained from the supersonic wind tunnel test data from NASA. Meanwhile, the evolution of the three-dimensional shape of the disk-gap-band parachute during supersonic inflation is presented, and the structural dynamic behaviour of the parachute is predicted. Then, the influence of the presence of the capsule on the flow field of the parachute is investigated, and the wake of unsteady fluid and the distribution of shock wave around the supersonic parachute are presented. Finally, the structural dynamic response of the canopy fabric under high-pressure conditions is comparatively analysed. The results show that the disk-gap-band parachute is well inflated without serious collapse. As the Mach numbers increase from 2.0 to 2.5, the drag coefficients gradually decrease, along with a small decrease in inflation time, which corresponds with test results, and proves the validity of the method proposed in this paper.

Experimental and computational methods were used to investigate the characteristics of a scaled, generically shaped weapons internal carriage and separation bay with a length-to-depth ratio of 4.5 at multiple Mach numbers and stagnation pressures. Three new nozzles were designed, manufactured, and characterized for the U.S. Air Force Institute of Technology small supersonic tunnel, yielding freestream Mach numbers of 1.43, 1.84, and 2.22. In addition, a control valve was reconfigured to achieve stagnation pressures as low as 1.0 psia, allowing more realistic scaling. These nozzles were used in conjunction with piezoresistive pressure transducers and high-speed schlieren photography to capture the time history of the pressure and the acoustic spectra of the cavity. The nominal Mach 2.3 nozzle was used in free-drop testing of a 1:20 scaled sphere and compared with computational simulations. The computational solution was obtained using the OVERFLOW solver with incorporated six-degree-of-freedom motion and t...