Supersonic Wind Tunnel Tests Research Articles

Fiber Bragg Grating-based sensor for measuring supersonic inlet terminal shock at different angles of attack

This study presents a multipoint pressure measurement system using fiber Bragg grating sensors to detect terminal shock positions. The system consisted of a fiber optic string encapsulated with multiple diaphragm-based fiber Bragg grating sensors, capable of detecting flow field characteristics at fourteen different axial positions of the inlet. The calibration results indicated a sensitivity of 46.47 pm/kPa, a linear correlation coefficient of 0.9997, and a maximum repeatability error below 1.44%. In supersonic wind tunnel tests, the system effectively measured pressure and located the terminal shock across varying angles of attack and throttling degrees. The maximum difference compared to conventional pressure sensors was 4.7%, and the terminal shock positioning error was limited to 1.2%. Overall, the system has robust multipoint measurement capabilities and is priced at only one-tenth of traditional sensors, showing great potential in the field of inlet measurements.

Design of a Wind Tunnel Nozzle Section Trolley Control System Based on Backstepping Method

Abstract To accommodate different types of supersonic wind tunnel tests, frequent replacements of nozzle sections and test sections are required, involving multiple complex technical processes. To optimize operational procedures and improve efficiency, a control system for the wind tunnel nozzle section trolley based on the backstepping method was designed. The system uses radio frequency identification (RFID) signals for station positioning and intelligently adjusts the speed and direction of the control motor, thus driving the trolley to move back and forth automatically. Through the simulation of the motor motion control model, the system effectively reduces random disturbances in the motion system caused by load changes, achieves precise positioning of wind tunnel plugins and nozzle assemblies, enhances system operating accuracy, reduces production cycle times, and realizes a one-button positioning function for all components, playing a significant role in improving wind tunnel test efficiency and operational quality.

Tehnika smanjenja opterećenja na unutrašnjim aerovagama tokom supersoničnih prelaznih pojava

Design of slender supersonic missiles requires a comprehensive experimental support in the form of wind tunnel data for a wide range of flight parameters (angle of attack and Mach number). However, in supersonic wind tunnel testing, the problem exists of transient loads at the times of the starting and stopping of the supersonic flow. In the environments of pronounced transient loads, characteristic of blowdowin wind tunnels like the T-38 of the Military technical Institute in Belgrade, it is necessary to provide control of the use of internal wind tunnel balances in the permitted design load ranges. The presented technique is related to the definition and implementation of a methodology for reducing the transient loads on wind tunnel balances in supersonic wind tunnel tests. By limiting the clearance between the model and its tail support (sting) to a magnitude which permits normal tests, but results in model-support contact during the excessive loads, part of the loads is transferred to the support sting, relieving the balance. The technique improves control over the wind tunnel test process, improves measurement accuracy and prevents damage to sensitive instrumentation (wind tunnel balances).

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.

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.

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.

Response of a plate with piezoelectric elements to turbulent pressure fluctuation in supersonic flow

The aeroelastic response of a plate with supersonic freestream flow on one side and a shallow cavity on the other to turbulent pressure fluctuations is investigated computationally and experimentally. An empirical model is developed for the pressure fluctuations in a turbulent boundary layer that accounts for spatial and spectral variations in the pressure field. Supersonic wind tunnel tests were conducted in a Mach 2.5 flow with and without an impinging shock at the plate surface. In both cases the boundary layer was turbulent. The impinging shock creates shock-wave boundary-layer interaction, which alters the characteristics of the pressure fluctuations. Pressure-sensitive paint was used to measure the unsteady pressure on the surface of a rigid plate and characterize the pressure field (local mean, rms, and the spatial coherence length) and piezoelectric patches were used as sensors to measure the response of an elastic plate. The extracted pressure parameters were used to simulate the fluid–structure response and correlate with experiments. The computed pressure perturbation due to plate motion is found to be small relative to the natural pressure fluctuation for the fluid/structural configuration studied. Computed and measured power spectra of the piezoelectric element voltage show good agreement over a wide range of structural natural frequencies. Aeroelastic response sensitivity to pressure fluctuation coherence length was also investigated computationally. It is found that with small fluid elements, which represent small-scale uncorrelated noise, the structural response is relatively small because the excitation is filtered by the plate dynamics. Experimental results suggest that the effective excitation spatial scale is on the order of the boundary layer thickness.

Unsteady test technique of jet control with multi-parameter coupling

In this paper, the jet control effect of air vehicle is studied, and a new multi parameter coupling unsteady test technique is established, which can ensure that the test model is consistent with the real air vehicle, and the jet is carried out under the condition of complete freedom and no support constraints, so that the wind tunnel test can better simulate the jet control effect of real air vehicle. The new technique overcomes the shortcoming that the previous jet test can only carry out unsteady flow field or unsteady aerodynamic force research, and can ensure that in the whole process of jet flow, the interference flow field at the nozzle, the aerodynamic force of the air vehicle, especially the movement of the air vehicle, are coupled with each other all the time, so as to achieve the same effect as the real air vehicle. The establishment of the new technique overcomes many problems, including the total length of the model is 167 mm, no support for the model, storage of small volume/high-pressure gas in free state, on-time unlocking of sealed gas source in free state of the model, and connection of gas source and nozzle in the model. The test results show that when no jet, the revolution body is head up, which affected by the cavity shock wave. When jet on, the jet force continues to act on the revolution body, resulting in the revolution body lowering its head until it diverges. The new technique can achieve more advanced technical indicators, including: ensuring jet Mach number >1; Ensure that the static pressure ratio of the jet is more than 10 (the static pressure ratio can be greater at hypersonic speed). After the establishment of the new technique, the supersonic wind tunnel test is carried out, and the ideal test results are obtained, which shows that the new technique is reliable.

Experimental study of high-speed air intake performance by side clearance

Japan Aerospace Exploration Agency (JAXA) and universities plan the “High Mach Integrated Control Experiment” (HIMICO), which is a hypersonic flight test by JAXA's S-520 sounding rocket. The experimental aircraft for HIMICO is powered by a ramjet engine, whose intake is a rectangular mixed compression type with a design Mach number of 5.In this research, the effects of the side clearance (SC) on the intake performance are investigated at Mach 3.4 by supersonic wind tunnel tests. The side clearances are thin slits between the 2nd and 3rd ramps and sidewalls, which is provided for driving the variable geometry intake. Three types of models with different SC widths were compared. As the results, it is found that the 2nd ramp SC, which is located upstream of the throat, causes the performance deterioration and the 3rd ramp SC, which is located downstream of the throat can expand the starting range. Hybrid SC which has only the 3rd ramp SC obtains sufficient Mass Capture Ratio (MCR) and maintains the start condition over a wider nozzle throat height range due to the air bleed from the 3rd ramp SC. However, the performance of Hybrid SC deteriorates when adding the sideslip angle, β=5°, because of a leakage from the bleed slit located between the 2nd and 3rd ramps.

Experimental investigation of the effects of leading edge bluntness on supersonic flow over a double compression ramp

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.

Classification of supersonic flow over a rigid parachute model with suspension lines

The supersonic wind tunnel tests using a rigid parachute model were conducted for Mach numbers ranging from 2.0 to 3.0. The rigid model consists of suspension lines and a riser between the forebody and rear body, and the effect of these suspension lines was investigated. Three characteristic flow modes, namely, high-pressure, low-pressure, and attachment modes, were observed. These three modes were classified on the basis of the shape and position of the apex of the shock wave. In addition, the non-dimensional pressure observed at the center of the canopy varied according to the pressure mode produced. The emergence of each pressure mode depended on the parameters, namely, freestream Mach number, trailing distance, and distance between the bundling point and rear body. Pressure-mode transitions were observed when specific conditions and configuration of the parameters were achieved.

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.

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.

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.

Supersonic test cases at high angles of attack

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

Photogrammetry for accurate model deformation measurement in a supersonic wind tunnel

The interest in adaptive devices for high-speed applications leads to the need for an accurate and reliable technique to obtain model deformation measurements during experiments. Point-tracking photogrammetry has been applied to supersonic wind tunnel testing, using four Phantom high-speed cameras placed on either side of the working section, where coded targets were applied to the surface of interest. Calibration experiments on a solid plate beneath a {M=1.4} normal shock and a {M=2} oblique shock allowed the quantification of the sources of optical distortion, namely the wind tunnel glass windows and aerodynamic effects (the lower pressure in the working section and the interaction between shock waves and the boundary layer). A correction matrix was applied to account for the optical distortion due to the glass, and the root-mean-square error due to aerodynamic effects ({<,0.03} mm) is believed to be negligible for applications with significant displacements (of the order of 1 mm). The application of photogrammetry to a flexible shock control bump has shown that the bump shape can be detected accurately, while disclosing some complex 3D effects that could not have been revealed by spanwise-averaged techniques such as schlieren photography.Graphical abstract3D reconstruction of an adaptive shock control bump from photogrammetry with the corresponding shadowgraph picture showing the lambda-shock structure (dimensions in mm)

Study of Pressure Oscillations in Supersonic Parachute

Supersonic parachutes are a critical element of planetary mission whose simple structure, light-weight characteristics together with high ratio of aerodynamic drag makes them the most suitable aerodynamic decelerators. The use of parachute in supersonic flow produces complex shock/shock and wake/shock interaction giving rise to dynamic pressure oscillations. The study of supersonic parachute is difficult, because parachute has very flexible structure which makes obtaining experimental pressure data difficult. In this study, a supersonic wind tunnel test using two rigid bodies is done. The wind tunnel test was done at Mach number 3 by varying the distance between the front and rear objects, and the distance of a bundle point which divides suspension lines and a riser. The analysis of Schlieren movies revealed shock wave oscillation which was repetitive and had large pressure variation. The pressure variation differed in each case of change in distance between the front and rear objects, and the change in distance between riser and the rear object. The causes of pressure oscillation are: interaction of wake caused by front object with the shock wave, fundamental harmonic vibration of suspension lines, interference between shock waves, and the boundary layer of suspension lines.

Numerical modelling of Mars supersonic disk-gap-band parachute inflation

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.

Improvement in spatial resolution of background-oriented schlieren technique by introducing a telecentric optical system and its application to supersonic flow

A telecentric optical system is applied to the background-oriented schlieren (BOS) technique to improve accuracy, overcoming the drawbacks of conventional diverging light observation. This paper describes the optical arrangement and formula for telecentric BOS measurement and presents measurement results obtained by the colored-grid background-oriented schlieren technique to confirm the theoretical prediction. The application of the new approach for a large-scale supersonic wind tunnel test is reported.

CFD modeling of supersonic airflow generated by 2D nozzle with and without an obstacle at the exit section

Computational modeling of complex supersonic airflow patterns is one of the greatest challenges in the domain of CFD analyses. The paper presents initial steps in numerical analysis of such flow, generated by convergentdivergent nozzle with Mach number M = 2.6 at nozzle exit. The aim was to achieve good agreements with available experimental data, obtained during supersonic wind tunnel tests at VTI Žarkovo institute, where nozzle thrust vectoring possibilities had been investigated using air as test fluid, by placing different types of obstacles at the exit section. Paper is focussed on free exit flow, and flow with one selected obstacle type. Using structured mesh for both cases, the RANS equations with k-w SST turbulent model have been applied. After quantitative and qualitative comparisons with available experimental data, good agreements have been obtained, where CFD was also able to provide additional flowfield data, not measuted during experiments.