Supersonic Wind Tunnel Research Articles

Correlation of supersonic wind tunnel measurements with a nonlinear aeroelastic theoretical/computational model of a thin plate

Supersonic wind tunnel test data is compared with a nonlinear aeroelastic computational model, considering a freestream flow with Mach nearly 2, coupled cavity, and no-shock impingement. The measurements include Limit Cycle Oscillations (LCO) for the given set of flow and structural parameters. The effect of a static pressure differential, temperature differential, and the type of LCO is also studied: periodic or chaotic. A fully coupled aero-thermal–acoustic–elastic analysis was carried out, where the strong dependency of the dynamic instability upon the in-plane boundary stiffness was discovered. Associated with aerodynamic and acoustic formulations, the nonlinearities from the structure are found to be the strongest factor determining the occurrence and character of LCO. At the same time, the cavity effect was found to be negligible for this experimental configuration. Additionally, a comparison analysis was performed on the aerodynamic model for this flow condition between the Linear Piston Theory and the Full Potential Aerodynamics. The calculation of the critical flutter boundary was also performed in terms of the static pressure differential between the cavity pressure and the panel surface pressure, presenting regions where the panel is in static deformation due to the pressure differential versus LCO response. The comparison with the analytical model showed good agreement with the measured data, achieving a similar panel displacement at three-quarters of the panel length as well as the expected behavior of periodicity or chaos for specific values of temperature and static pressure differential.

Thermal Behavior of Carbon-Phenolic/Silica Phenolic Dual-Layer Ablator Specimens through Arc-Jet Tests.

We studied the behavioral characteristics of a newly developed dual-layer ablator, which uses carbon-phenolic as a recession layer and silica-phenolic as an insulating layer. The ablator specimens were tested in a 0.4 MW supersonic arc-jet plasma wind tunnel, employing two different shapes (flat-faced and hemispherical-faced) and varying thicknesses of the carbon-phenolic recession layer. The specimens underwent two test conditions, namely, stationary tests (7.5 MW/m2, ~40 s) and transient tests simulating an interplanetary spacecraft re-entry heat flux trajectory (6.25↔9.4 MW/m2, ~108 s). During the stationary tests, stagnation point temperatures of the specimens were measured. Additionally, internal temperatures of the specimens were measured at three locations for both stationary and transient tests: inside the carbon-phenolic recession layer, inside the silica-phenolic insulating layer, and at the recession layer-insulating layer intersection. The hemispherical-faced specimen surface temperatures were about 3000 K, which is about 350 K higher than those of flat-faced specimens, resulting in higher internal temperatures. The recession layer internal temperatures rose more exponentially when moved closer to the specimen stagnation point. Layer interaction and insulating layer internal temperatures were found to be dependent on both the recession layer thickness and the exposed surface shape. The change in exposed surface shape increased mass loss and recession, with hemispherical-faced specimens showing ~1.4-fold higher values than the flat-faced specimens.

Study of Self-Excited Pitch Oscillations of Conical-Spherical Body at Mach Number M = 1.75 and Two Moments of Inertia

Tests of a conical-spherical body in a supersonic wind tunnel were carried out on a free oscillation installation along the pitch angle at a Mach number M = 1.75 and two moments of inertia of the body relative to the axis of rotation. For a body with a large moment of inertia, self-oscillations with the amplitude of ca. 10 deg were obtained. For the first time, the phenomenon of degeneration of self-oscillations for a body with a smaller moment of inertia has been discovered. An attempt was made to model the phenomenon using an ordinary linear differential equation of the first order

Robustness and performance evaluation of TDLAS sensor for scramjet intake

A single-body typed TDLAS sensor was developed, and its robustness and performance were experimentally evaluated under harsh thermal and mechanical environments which could be formed in scramjet intake. To confirm integrity and operability of the TDLAS sensor in the harsh thermal environments, high-thermal gradient, long-duration heating, and aero-heating tests were conducted. The high-thermal gradient, long-duration heating, and aero-heating tests were performed using high-temperature torch, furnace, and supersonic wind tunnel, respectively. In addition, vibration, mechanical shock, and acceleration tests were performed to analyze the durability and operating performance of the TDLAS sensor in the harsh mechanical environments. The vibration/mechanical shock, and acceleration tests were carried out using shaker and centrifuge, respectively. The TDLAS sensor did not experience any damage and had a stability of signal measurement under these severe thermal and mechanical environments. The experimental test results showed that the single-body TDLAS sensor in this paper could be used in the extreme thermal and mechanical conditions which could be formed in the scramjet intake.

Research on time series prediction of the flow field in supersonic combustor based on deep learning

The detection of supersonic combustion flow field provides valuable reference information for the efficient combustion of hypersonic vehicles. From the perspectives of scramjet operation during flight as well as experiment where the measured information is limited, flow field prediction using deep learning technology is a promising method that can provide future flow field evolution in scramjet combustors. This study proposes a convolutional neural network model of multi-temporal pressure information fusion to predict combustion flow field based on wall pressure. In addition, it analyzes the effect of various wall pressures on the flow field prediction. An experimental dataset of supersonic ground-pulse wind tunnel tests with different injection pressures is built through the combustion chamber wall under self-ignition conditions of a hydrogen-fueled scramjet. The results showed that the predicted flow field of the combustor agrees with the ground test, average correlation coefficient reached more than 90% and when the wall pressure number is 32, the average peak signal to noise ratio reaches about 23. and the main wave structure and turbulent boundary layer region are restored even if there is a shock boundary layer interaction.

Arc-Jet Tests of Carbon-Phenolic-Based Ablative Materials for Spacecraft Heat Shield Applications.

We developed and tested two carbon-phenolic-based ablators for future Korean spacecraft heat shield applications. The ablators are developed with two layers: an outer recession layer, fabricated from carbon-phenolic material, and an inner insulating layer, fabricated either from cork or silica-phenolic material. The ablator specimens were tested in a 0.4 MW supersonic arc-jet plasma wind tunnel at heat flux conditions ranging from 6.25 MW/m2 to 9.4 MW/m2, with either specimen being stationary or transient. Stationary tests were conducted for 50 s each as a preliminary investigation, and the transient tests were conducted for ~110 s each to stimulate a spacecraft's atmospheric re-entry heat flux trajectory. During the tests, each specimen's internal temperatures were measured at three locations: 25 mm, 35 mm, and 45 mm from the specimen stagnation point. During the stationary tests, a two-color pyrometer was used to measure specimen stagnation-point temperatures. During the preliminary stationary tests, the silica-phenolic-insulated specimen's reaction was normal compared to the cork-insulated specimen; hence, only the silica-phenolic-insulated specimens were further subjected to the transient tests. During the transient tests, the silica-phenolic-insulated specimens were stable, and the internal temperatures were lower than 450 K (~180 °C), achieving the main objective of this study.

Experimental investigation of mini supersonic wind tunnel for flow visualization

The present study proposes to design, fabricate and investigate a miniature supersonic wind tunnel at an affordable cost, safe to operate, and reliable for education purposes. The configuration consists of a high pressure reservoir, and mini tunnel which will be used for providing the supersonic flow over aerodynamic models. To design the mini supersonic wind tunnel, the method of characteristics is used to ensure that any shocks that arise in contact with the surface do not impact the uniformity and velocity of the flow through the duct. To achieve the supersonic flow, the nozzle contour of the supersonic portion of the wind tunnel is built in such a way that the shocks nullify each other and have zero effect on the velocity of the flow. For compressible flow visualization, an affordable Schlieren setup consisting of a torch, mirrors, and a mobile camera is used for flow visualization. Finally, the visualization results of the present study is in good relation with the theoretical calculations. The present facility is also calibrated by measuring the oblique shock angle in the test section, in the range of Mach number 1.6 to 1.8.

Measurement and Identification of Supersonic Stationary Crossflow Waves Based on Sublimation Method

Aiming at the crossflow transition problem of supersonic swept wings, an experimental study on the measurement and identification of boundary-layer stationary crossflow waves was carried out at Mach number 2.0. A 65 deg straight-swept-wing model with simple airfoil geometry was designed to capture the crossflow transition phenomenon. The sublimation method for a large-scale cold supersonic wind tunnel was developed to acquire a transition pattern. The stationary crossflow wave identification technology is based on image processing, a vertical and horizontal decomposition method is established, and the quantitative characteristic information (such as the vortex axis spacing and direction) can be identified from high-resolution sublimation images. The experimental results show that the transition position caused by crossflow instability is closed to the leading edge for supersonic wings with large sweep angles at high Reynolds numbers, and the transition position moves upstream with the increase of the unit Reynolds number. The identified vortex axis spacing and direction are basically consistent with the linear stability theory. The stationary disturbance wave information at different Reynolds numbers is compared and analyzed. It is found that the unit Reynolds number plays an important role in the selection and amplification of the unstable stationary modes with different wavelengths in the transition process.

Investigation of Three-Dimensional Shock Control Bumps for Transonic Buffet Alleviation

This experimental study investigates the use of shock control bumps (SCBs) for controlling transonic buffet. Three-dimensional SCBs have been applied on the suction side of an OAT15A supercritical airfoil with the experiments conducted in the transonic–supersonic wind tunnel of Delft University of Technology at fully developed buffet conditions (, and ). The effectiveness of the SCBs for different spanwise array spacings (ranging from 20 to ) was verified using two optical techniques: schlieren visualization and particle image velocimetry. Both techniques confirmed the potential of controlling buffet using such devices, resulting in a reduction of the flow unsteadiness in terms of both shock oscillation and pulsation of the separated area. A dedicated particle image velocimetry investigation in a spanwise–chordwise measurement plane was conducted in order to characterize the effect of the spatial distribution of the bumps, focusing on the interaction of the shock-wave structures along the span. The configuration with a spacing of was demonstrated to be the most efficient in reducing the transonic buffet oscillations and was able to reduce the reverse flow region size as compared to the clean configuration.

Density field measurement deflectometry for supersonic wind tunnels.

A method for the quantitative measurement of two-dimensional density field distributions based on deflectometry is proposed. With this method, from the point of view of the inverse Hartmann test, the light rays emitted from the camera reach the screen after being disturbed by the shock-wave flow field. After the coordinates of the point source are obtained by using the phase information, the deflection angle of the light ray can be calculated, and then the distribution of the density field can be determined. The principle of density field measurement deflectometry (DFMD) is described in detail. In the experiment, the density fields in wedge-shaped models with three different wedge angles are measured in supersonic wind tunnels, the experimental results from the proposed method are compared with the theoretical results, and the measurement error is found to be around 2.76 × 10-3kg/m3. This method has the advantages of fast measurement, a simple device, and low cost. It provides a new approach, to the best of our knowledge, to measuring the density field of a shock-wave flow field.

Experimental Investigation of Flameholding in Scramjet Combustor by Pylon with Plasma Actuator Based on Q-DC Discharge

This paper presents the results of testing and optimization of a plasma-assisted combustion scheme based on a pylon for fuel injection equipped with a plasma actuator. Electrodes were installed behind the stern of the pylon for the creation of Q-DC discharge with voltage U = 200–2500 V and current I = 3–7.5 A. The experiments were performed in the PWT-50 supersonic wind tunnel of the JIHT RAS under the following conditions: Mach number M = 2, static pressure ~200 Torr, stagnation temperature T0 = 300 K. Gaseous fuel ethylene was used and the fuel mass flow rate was 0.5–4 g/s. The pylon had a streamlined shape that prevented the formation of a stagnant zone; plasma-assisted combustion was performed under more difficult conditions compared to plasma-assisted combustion on a flat wall, where separated flows near the wall are easily formed by discharge. In this work, two new geometries of pylon equipped with electrodes were proposed and experimentally tested. A second version providing a longer discharge length demonstrates stable ignition and intense combustion in a fully discussed fuel mass flow rate. The process of ignition in a supersonic flow and flame front pulsations was described. A reduction in the energy input in comparison with the previously considered configurations of plasma-assisted combustion was also demonstrated.

Instantaneous Ablation Behavior of Laminated CFRP by High-Power Continuous-Wave Laser Irradiation in Supersonic Wind Tunnel.

Experimental and numerical investigations of the instantaneous ablation behavior of laminated carbon fiber-reinforced polymer (CFRP) exposed to an intense continuous-wave (CW) laser in a supersonic wind tunnel are reported. We establish an in situ observation measurement in the experiments to examine the instantaneous ablation behavior. The surface recession depth is calculated by using the Particle Image Velocimetry (PIV) method, taking the ply angle of laminated CFRP as a reference. A coupled thermal-fluid-ablation numerical model incorporating mechanisms of oxidation, sublimation, and thermomechanical erosion is developed to solve the ablation-through problem of multilayer materials. The results show that the laser ablation depth is related to the laser power density, airflow velocity and airflow mode. Thermomechanical erosion is the primary ablation mechanism when the surface temperature is relatively low and the cavity flow mode is a closed cavity flow. When the surface temperature reaches the sublimation of carbon and the airflow mode is transformed to open cavity flow, sublimation plays a dominant role and the ablation rate of thermomechanical erosion gradually decreases.

High pitch sweep rates in supersonic wind tunnel tests

A typical test in a high-Reynolds-number blowdown wind tunnel, such as the VTI T-38 facility in Belgrade, involves a model sweep in a range of angles of attack. Increasing pitch sweep rates can bring two-fold benefit. The angle-of-attack range can be extended in tests at high stagnation pressures with limited run time, or test time can be shortened for the desired angle-of-attack range in other tests, in both cases bringing significant energy savings. The potential effect of the increased pitch rates on measurement data in the VTI T-38 wind tunnel was investigated in a recent supersonic wind tunnel test campaign of a hypervelocity standard model. The Mach 2 tests were performed at pitch rates of 2, 3, 4, 8 and 12° per second. While the wind tunnel balance measurement data were not seemed to be affected, the base pressure data revealed a high dependency on changing pitch sweep rates. It was found that the high lag in pressure piping affected the measured base pressure and, consequently, the forebody axial force, which is an important fact to know when dealing with ‘base drag corrected’ supersonic data. The conclusions derived from the Mach 2 tests have been verified in the additional Mach 4 tests. The verification tests confirmed that, for the currently employed base pressure measurement system, reliable measurement results can be obtained in wind tunnel tests by applying pitch rates up to 3° per second. The recommendations for modifications of the base pressure measurement system have been given.

A large-deflection deformation reconstruction method for semi-flexible plate nozzle based on strain-moment relationship

To enhance the health detection capability of the semi-flexible plate nozzle in a supersonic wind tunnel, a large-deflection deformation reconstruction algorithm for a semi-flexible plate nozzle based on the strain-moment relationship is developed. The multi-jack semi-flexible plate nozzle in the supersonic wind tunnel is studied; its mechanical model of contour assembly is established. The differential equation of flexible plate contour curve is deduced, and the elliptic integral expression of the flexible plate deformation is given. Based on the deformation theory of a rectangular thin plate, the deformation reconstruction algorithm is modified, and its correction effect is simulated and analysed with the finite element method. The maximum deviation after correction is only 10.78% compared with that before correction. The 0.3 m semi-flexible plate nozzle is used as the experimental platform to carry out the experimental study. The experimental results show that under the 6 Mach number, the large-deflection deformation reconstruction results based on strain-moment relationship are in good agreement with the experimental results. The coordinate deviation of the nozzle throat position is about 6.12% compared with the structural deformation, thus verifying that the deformation reconstruction algorithm in this paper is effective and accurate.

Nozzle geometry effects on supersonic wind tunnel studies of shock–boundary-layer interactions

Many supersonic wind tunnel experiments investigate shock–boundary-layer interactions by measuring the response of tunnel wall boundary layers to an incident shock wave. To generate the supersonic flow, these facilities typically use two-dimensional contoured converging–diverging nozzles which can be arranged in two different ways. One configuration is symmetric about the centre height, whereas this symmetry plane defines the tunnel floor in the other asymmetric arrangement. In order to determine whether these nozzle configurations, which are widely thought to be equivalent, can influence experiments on shock–boundary-layer interactions, two different nozzle geometries are compared with one another in a single facility with rectangular cross section. For each setup, a full-span 8-degree wedge introduces an oblique shock to a Mach 2.5 flow. The two setups exhibit quite dissimilar behaviour, both in the corner regions and on the tunnel’s centre span, with a difference in central separation length of as much as 35% suggesting that nozzle geometry can have a profound impact on these interactions. The observed behaviour is caused by known secondary flows in the sidewall boundary layers which are driven by vertical pressure gradients in the nozzle region. The subsequent impact on the response of the floor boundary layer is consistent with expectations based on local flow momentum affecting corner separation size and on the displacement effect of this corner separation influencing the wider flow.Graphical abstract

Data Reduction Analysis on UTM-LST External Balance

Abstract: The effective use of wind-tunnel testing in determining aerodynamic properties of a body is very much dependent upon the reliability and speed with which wind-tunnel data can be reduced. The operating efficiency of the wind tunnels is substantially improved by the capability of providing lower aerodynamic coefficients in real time, or online, which decreases the expensive wind-tunnel time necessary for each test. This paper describes a system for presenting reduced wind-tunnel data in real time for UTM-LST. The requirements for data-handling equipment and data reduction procedures for wind tunnels are quite diverse, and depend upon the wind tunnel design and the type of tests for which they are used. The supersonic wind tunnels mentioned in this description have a variety of force-balance systems and are used for force tests, pressure tests, and other research projects. Consequently, the problems associated with in order to solve this diversity we build a computerized program where we can find the transformation of axis and aerodynamic characteristics at ease. This program can find the values of different aerodynamic coefficients with certain angle of attacks.

Suppression of flow response hysteresis in the throttling/unthrottling process for supersonic inlet

The flow response hysteresis in a mixed-compression supersonic inlet's throttling/unthrottling process is unconducive to its practical engineering applications. To suppress such hysteresis, herein, a flow control method based on a distributed bleed system in the inlet's internal contraction part was proposed and applied to a rectangular supersonic inlet with a designed Mach number of 4.0 and an internal contraction ratio of 1.5. All tests were conducted in a supersonic wind tunnel at a freestream Mach number of 2.9, and the relevant test data were obtained using the high-speed schlieren and dynamic pressure measurement systems. The results indicate that the distributed bleed system improved the uncontrolled inlet's unstart throttling ratio from 43.8% to 47.3% and the restart throttling ratio from 29.4% to 47.1%, and it essentially eliminated the flow response hysteresis during the throttling/unthrottling process. Moreover, the hysteresis mechanism of the uncontrolled inlet, dominated by the separation-induced flow structures in the internal contraction part, was investigated to help elucidate the targeted control mechanism of the bleed system that suppresses the hysteresis. Finally, the specific flow phenomena in the oscillation of the controlled inlet were examined to study the bleed system's impact on the unsteady flowfield, including the coupled flow of the mainstream with the bleed system and complex wave motion structures.

Nozzle design optimization for supersonic wind tunnel by using surrogate-assisted evolutionary algorithms

For high-precision measurement in supersonic wind tunnel experiments, it is of crucial importance to produce a uniform flow in the measurement section downstream of the nozzle. This paper proposes and verifies a new design methodology for supersonic wind tunnel nozzles that can generate highly uniform airstream at a design Mach number at the nozzle exit. Shape design optimization has been conducted by employing surrogate-assisted evolutionary algorithms coupled with computational fluid dynamics. This approach has yielded higher flow uniformity than that of the nozzle designed by using the method of characteristics in the inviscid regime. By applying boundary layer correction to the nozzle contour obtained from the inviscid optimization, nearly uniform core flow of Mach 2.5 has been achieved at the nozzle exit in the presence of viscosity. In the viscous optimization, nozzle shape optimization has been performed by incorporating viscous simulations without using boundary layer correction to evaluate its efficacy in comparison with the former approach. It has been found that the deviation from the design Mach number and the flow deflection from the horizontal direction cannot be minimized simultaneously. This has been attributed to the constraint associated with the nozzle length. It has also been found that the result of the former approach combining the inviscid optimization and boundary layer correction can be regarded as one of the results of the viscous optimization.

Experimental study of effect of a pair of weak shock waves on laminar-turbulent transition in the boundary layer of flat plate at Mach number 2

An experimental study of the effect of longitudinal vortices generated by weak shock waves on the laminar-turbulent transition in the boundary layer of a flat plate at a Mach number 2 is conducted. To amplify the effect of the longitudinal vortices generation, a model with leading edge blunting radius of 0,5 mm was used. The disturbance intensity was measured with a constant temperature hotwire anemometer and a single-wire sensor. To change the Reynolds numbers in the flow, the method of varying the pressure in the pretest section of a supersonic wind tunnel was used. According to the experimental conditions, the level of mass flow perturbations in the test section T325 of ITAM SB RAS did not exceed 0,1 % of the mean flow in the frequency range from 0,2 to 50 kHz. In this work, it is obtained for the first time that stationary streamwise vortices that arise in the boundary layer of a flat plate because of the action of a pair of weak shock waves on its leading edge lead to an early laminar-turbulent transition at different Reynolds numbers. In the case of a homogeneous boundary layer, the transition is achieved at Rex≈ 3,1 × 106, and in measurements in the region of vortices from shock waves at Rex≈ 1,5 × 106and Rex≈ 2 × 106.

Dual Injection in Supersonic Crossflow: Analysis Jet Shear Layer from Schlieren Images

In supersonic-combustion ramjets (scramjets), fuel is injected, which should mix rapidly with the supersonic crossflow to minimize the length of the scramjet. Tandem dual-jet injection has shown improved mixing performance over single-jet injection. However, experiments on tandem dual-jet injection have not addressed the jet shear layer, in which the mixing occurs, yet. The present study investigates the jet shear layer, as well as the bow shocks in front of the jets, in a continuous air-indraft supersonic wind tunnel at Mach number 1.55. A schlieren setup has been used for visualizing the flow features. A largely automated algorithm for processing schlieren images has been developed to determine the location of the upper boundary of the jet shear layer. The penetration of the jet is studied as a function of 1) , the ratio of the momentum of the jet and that of the crossflow, and 2) the dimensionless distance between the dual-jets. An empirical similarity relation has been established for the time-averaged location of the jet upper shear layer as function of and , covering the investigated conditions (, ). This empirical similarity relation provides , the spacing for maximal penetration of the jets as function of .