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Journal articles on the topic 'Transonic tunnel'

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Author: Grafiati

Published: 10 December 2022

Last updated: 29 January 2023

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1

Greenwell, D. I. "Transonic industrial wind tunnel testing in the 2020s." Aeronautical Journal 126, no. 1295 (December 2, 2021): 125–51. http://dx.doi.org/10.1017/aer.2021.107.

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AbstractWind tunnels remain an essential element in the design and development of flight vehicles. However, graduates in aerospace engineering tend to have had little exposure to the demands of industrial experimental work, particularly at high speed, a situation exacerbated by a lack of up-to-date reference material. In an attempt to fill this gap, this paper presents an overview of the current and near-term status and usage of transonic industrial wind tunnels. The review is aimed at recent entrants to the field, with the aim of helping them make the step from research projects in small university facilities to commercial projects in large industrial facilities. In addition, a picture has emerged from the review that contradicts received wisdom that the wind tunnel is in decline. Globally, the industrial transonic wind tunnel is undergoing somewhat of a renaissance. Numbers are increasing, investment levels are rising, capabilities are being enhanced, and facilities are busy.

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2

Tsushima, Natsuki, Kenichi Saitoh, Hitoshi Arizono, and Kazuyuki Nakakita. "Structural and Aeroelastic Studies of Wing Model with Metal Additive Manufacturing for Transonic Wind Tunnel Test by NACA 0008 Example." Aerospace 8, no. 8 (July 25, 2021): 200. http://dx.doi.org/10.3390/aerospace8080200.

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Additive manufacturing (AM) technology has a potential to improve manufacturing costs and may help to achieve high-performance aerospace structures. One of the application candidates would be a wind tunnel wing model. A wing tunnel model requires sophisticated designs and precise fabrications for accurate experiments, which frequently increase manufacturing costs. A flutter wind tunnel testing, especially, requires a significant cost due to strict requirements in terms of structural and aeroelastic characteristics avoiding structural failures and producing a flutter within the wind tunnel test environment. The additive manufacturing technique may help to reduce the expensive testing cost and allows investigation of aeroelastic characteristics of new designs in aerospace structures as needed. In this paper, a metal wing model made with the additive manufacturing technique for a transonic flutter test is studied. Structural/aeroelastic characteristics of an additively manufactured wing model are evaluated numerically and experimentally. The transonic wind tunnel experiment demonstrated the feasibility of the metal AM-based wings in a transonic flutter wind tunnel testing showing the capability to provide reliable experimental data, which was consistent with numerical solutions.

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3

Qian, Wei, and De Guan. "The Design, Manufacture and Wind Tunnel Test of the Full Aircraft Transonic Flutter Model." Advanced Materials Research 487 (March 2012): 267–72. http://dx.doi.org/10.4028/www.scientific.net/amr.487.267.

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This paper discusses the design, manufacture and wind tunnel test of a full aircraft structure similar transonic flutter model in the wind tunnel FL-26. It introduces the mechanics hypothesis, use of materials, and design methods of this model design, in which it uses a technology of dynamic finite element model’s flexibility-mode collaborative correction. In the process of the model, it adopts glass fiber, carbon fiber reinforced plastic and foam for manufacturing of dynamics similar model. After simulation calculation of the model, transonic flutter wind tunnel test of the model is finally accomplished in the wind tunnel FL-26.

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4

Chen, Dan, Xiaosong Yang, Gang Li, Shouchun Guo, and Tianyi Chen. "Relativity Research of Total Pressure and Regulating Valve in Continuous Wind Tunnel and Its Application." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 38, no. 2 (April 2020): 325–32. http://dx.doi.org/10.1051/jnwpu/20203820325.

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As the main adjusting means of the total pressure for the continuous transonic wind tunnel, the characteristics of regulating valve directly affect the flow field performance of the wind tunnel, therefore, it is important to analyze and establish the correlation between the regulating valve and the total pressure, and it is necessary to select the appropriate regulating valve and its combination accordingly. Firstly, in terms of the pressure regulation principle of the wind tunnel pressure regulating system, combining with the flow characteristics of the regulating valve, the correlation between the position control of the regulating valve and the total pressure control of the wind tunnel is established, then the static test is conducted to verify the relationship. In order to shorten the flow field stability time under the negative pressure of 0.6m continuous transonic wind tunnel, based on the established theory, the valve system is optimized and reformed, and the blowing test is carried out. The results show that the time of optimized Mach number polar curve decreases by 40%~50%, which greatly improves the test efficiency, which further proves that the present analysis is correct and effective, and can provide reference for the design of pressure regulating system in continuous transonic wind tunnel.

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5

Kaczyński, P., R. Szwaba, M. Piotrowicz, P. Flaszyński, and P. Doerffer. "Wind tunnel investigations of aircraft airfoil in cruise conditions." Journal of Physics: Conference Series 2367, no. 1 (November 1, 2022): 012019. http://dx.doi.org/10.1088/1742-6596/2367/1/012019.

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Abstract Presented work is focused on analysis of the flow over the 2D model of Airbus A320 airfoil wing in cruise phase of the flight. For this purpose the measurement channel with an airfoil model was designed and assembled in a transonic wind tunnel to obtain a similar flow pattern as in the reference two-dimensional freestream flow. Experimental investigations were conducted in the IMP PAN transonic wind tunnel with a relative narrow test section which is a novel approach in terms of these type of research. The test section was designed using CFD simulations based on 2D freestream flow for the tested wing profile and in the next step the research were continued experimentally in transonic wind tunnel with a measurement chamber width of 100 mm. This paper presents the results of reference investigations on the A320 wing profile which combines experimental tests and CFD calculations. The obtained results show that the approach presented in the paper is appropriate and the obtained flow features in the tunnel do not differ much from the freestream conditions.

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6

Kiock, R., F. Lehthaus, N. C. Baines, and C. H. Sieverding. "The Transonic Flow Through a Plane Turbine Cascade as Measured in Four European Wind Tunnels." Journal of Engineering for Gas Turbines and Power 108, no. 2 (April 1, 1986): 277–84. http://dx.doi.org/10.1115/1.3239900.

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Reliable cascade data are esssential to the development of high-speed turbomachinery, but it has long been suspected that the tunnel environment influences the test results. This has now been investigated by testing one plane gas turbine rotor blade section in four European wind tunnels of different test sections and instrumentation. The Reynolds number of the transonic flow tests was Re2 = 8 × 105 based on exit flow conditions. The turbulence was not increased artificially. A comparison of results from blade pressure distributions and wake traverse measurements reveals the order of magnitude of tunnel effects.

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7

BRUCE, P. J. K., D. M. F. BURTON, N. A. TITCHENER, and H. BABINSKY. "Corner effect and separation in transonic channel flows." Journal of Fluid Mechanics 679 (May 31, 2011): 247–62. http://dx.doi.org/10.1017/jfm.2011.135.

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An investigation into parameters affecting separation in normal shock wave/boundary layer interactions (SBLIs) has been conducted. It has been shown that the effective aspect ratio of an experimental facility (defined as δ*/tunnel width) is a critical factor in determining when shock-induced separation will occur. Experiments examining M∞ = 1.4 and 1.5 normal shock waves in a wind tunnel with a small rectangular cross-section have been performed and show that a link exists between the extent of shock-induced separation on the tunnel centre-line and the size of corner-flow separations. In tests where the corner-flows were modified ahead of the shock (through suction and vortex generators), the extent of separation around the tunnel centre-line was seen to vary significantly. These observations are attributed to the way corner flows modify the three-dimensional shock-structure and the impact this has on the magnitude of the adverse pressure gradient experienced by the tunnel wall boundary layers.

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8

Damljanović, Dijana, Đorđe Vuković, Goran Ocokoljić, and Boško Rašuo. "Convergence of transonic wind tunnel test results of the AGARD-B standard model." FME Transactions 48, no. 4 (2020): 761–69. http://dx.doi.org/10.5937/fme2004761d.

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

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9

Phillips, Pamela S., and Edgar G. Waggoner. "Transonic wind-tunnel wall interference prediction code." Journal of Aircraft 27, no. 11 (November 1990): 915–16. http://dx.doi.org/10.2514/3.45959.

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10

Edwards, John W. "National Transonic Facility Model and Tunnel Vibrations." Journal of Aircraft 46, no. 1 (January 2009): 46–52. http://dx.doi.org/10.2514/1.30080.

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11

Mokry, M., M. Khalid, Y. Mébarki, and A. Rebaine. "Experimental and numerical Investigation of wind tunnel wall interference near Mach one." Aeronautical Journal 105, no. 1052 (October 2001): 589–96. http://dx.doi.org/10.1017/s0001924000012537.

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Surface pressure measurements on the CAST-10-2/DOA 2 aerofoil, conducted in the IAR l.5m wind tunnel and supported by CFD simulations, are used to validate some new theoretical analyses of transonic wall interference. Based on the transonic freeze principle, it is shown that the stream Mach number correction is indeterminate near Mach one.

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12

Chen, Jiming, Shenghao Wu, Zhenhua Chen, Jinlei Lyu, and Haitao Pei. "Experimental Research on Noise Reduction for Continuous Transonic Wind Tunnel Loop." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 38, no. 4 (August 2020): 855–61. http://dx.doi.org/10.1051/jnwpu/20203840855.

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The noise level of wind tunnel test section is respected as one of the most important performance specifications to represent the flow field quality, especially for large scale wind tunnel. According to the acoustic experimental research conducted in the 0.6 m continuous transonic wind tunnel of CARDC, main noise sources in the tunnel loop included the compressor, the high-speed diffuser and the test section. To reduce the noise in the test section, it is necessary to prevent the test section from the compressor noise propagated both forward and backward. In 0.6 m wind tunnel loop, acoustic treatments were installed on both the compressor rear cone and the fourth corner to prevent the noise emitted from the compressor from propagating forward. The vanes in the forth corner were filled with glass fibers and covered with perforated panels. And the compressor rear cone was covered with three layers of micro-perforated panels. With acoustic treatment in the tunnel loop and the second throat throttling, the fluctuation pressure coefficient (ΔCp) is lower than 0.8%, which is close to the international advanced level.

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13

HAN, Jiangxu, Nan LIU, Xiaoming SHI, Jin GUO, Song WANG, and Xianpeng YU. "Flutter test of rudder with real servo actuator in continuous transonic wind tunnel." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 40, no. 2 (April 2022): 401–6. http://dx.doi.org/10.1051/jnwpu/20224020401.

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Flutter wind tunnel test is an important approach of investigation of transonic flutter characteristics of flight vehicle. Comparing with the blow-down wind tunnel, the long-running and low dynamic pressure capabilities of continuous transonic wind tunnel are very suitable for flutter test. The flutter safety protection and analysis of dynamic signals are developed. The safety protection control software, rapid reduction of Mach number and dynamic pressure, model debris catch screen are integrated, which can provide safety protection of test model and wind tunnel. During the test process, the flutter boundary can be achieved by interpolating the reciprocal of spectrum peak. The flutter tests of rudder are conduct through two methods of step and continuous varying dynamic pressure. It is illustrated that the error of flutter dynamic pressure is relatively small, less than 5% between the two methods. Meanwhile, the feedback effect of the real servo actuator on the flutter characteristics is hard to be obtained via numerical simulation. It is demonstrated that the flutter dynamic pressure has been increased by 10% due to the feedback effect.

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14

Xue, Fei, Yuchao Wang, and Peng Bai. "Investigation of aerodynamics of separator delivery from cavity." International Journal of Modern Physics B 34, no. 14n16 (June 2, 2020): 2040104. http://dx.doi.org/10.1142/s0217979220401049.

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The ejection test technology is studied in a sub-transonic supersonic wind tunnel using a single cylinder to provide ejection velocity. The angular velocity adjusting device of ejection mechanism is designed, which can adjust the ejection velocity and angular velocity of the model independently. When the ejection cylinder moves downward, the angular velocity adjusting mechanism works at the same time, so that the model has the preset ejection velocity and angular velocity at the moment of leaving the ejection frame. The ejection velocity error is less than 5%, the angular velocity error is less than 10%, and the repetition rate is more than 95%. The new technology has been verified by wind tunnel tests under complex aerodynamic conditions of sub-transonic supersonic and multi-body interference. All parameters have reached or surpassed the existing technical specifications. It has served for model tests many times and met the needs of wind tunnel test research on ejection of embedded weapons in aircraft.

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15

Fei, XUE, MING Chengdong, WANG Huaqiang, and WANG Yuchao. "Research on High Brightness Light Source Equipment for Wind Tunnel Test." MATEC Web of Conferences 288 (2019): 02008. http://dx.doi.org/10.1051/matecconf/201928802008.

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The experimental technology of high brightness light source was studied in sub-transonic supersonic wind tunnel. The elevation light source should be installed on the smooth wall of the tunnel, and the elevation camera should be installed in the safe area of the lower wall of the wind tunnel. The falling image of the missile model in the test is reflected into the elevation camera through a reflector mounted on a curved knife. The full trajectory images and aerodynamic parameters of projectiles of embedded weapons in aircraft can be obtained by the wind tunnel dual-view angle, high brightness light path system and six-degree-of-freedom image analysis system. The newly developed high brightness light source system makes the image clearer and the accuracy of model angle of attack identification less than 0.2 degrees, which is conducive to the analysis of model trajectory. The optical system is designed reasonably, so that the motion trajectory and six-degree-of-freedom data of the model can be obtained easily by using the dual-view technology. Wind tunnel tests under complex aerodynamic conditions of sub-transonic supersonic and multi-body interference have been completed, and all parameters have reached or surpassed the existing technical indicators, meeting the requirements of wind tunnel test research on ejection of embedded weapons in aircraft.

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16

Mohammed Asadullah, Sher Afghan Khan, Parvathy Rajendran, and Ervin Sulaeman. "Design Intent of Future Tunnels." Journal of Advanced Research in Fluid Mechanics and Thermal Sciences 88, no. 2 (November 1, 2021): 50–63. http://dx.doi.org/10.37934/arfmts.88.2.5063.

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The sound barrier for bullet trains remains a challenge due to the piston effect causing compression waves at the entry and exit of the tunnel. The air ahead of the train nose is compressed, and the wave propagates through the tunnel at the speed of sound and exits with the generation of micro pressure waves. It gives rise to a complex wave pattern comprising compression at the train nose & expansion at the train tail leading to the positive pressure around the nose and suction around the tail. This is intended to provide exhaustive input for the proper design of a futuristic tunnel. The cross-sectional shapes of the tunnel, whether square, rectangular, circular, or semi-circular, will experience pressure compression wave generated by high-speed train but will influence the flow pattern and hence the compression wave. This paper presents the pressure load on the walls of long and short tunnels for subsonic compressible and transonic flows. The experimental investigation is carried out only for length parameters to study short and long tunnels. Further, flow visualization is also provided after the formation of the sonic boom. The results of this investigation can be an essential data source for optimum design of high-speed tunnels so as to suppress or break the sound barriers, thus, resulting in a safer high-speed train network.

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17

Ristić, S. "Disturbance of transonic wind tunnel flow by a slot in the tunnel wall." Experiments in Fluids 11, no. 6 (October 1991): 403–4. http://dx.doi.org/10.1007/bf00211796.

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18

Yeager, William T., Paul H. Mirick, M‐Nabil H. Hamouda, Matthew L. Wilbur, Jeffrey D. Singleton, and W. Keats Wilkie. "Rotorcraft Aeroelastic Testing in the Langley Transonic Dynamics Tunnel." Journal of the American Helicopter Society 38, no. 3 (July 1, 1993): 73–82. http://dx.doi.org/10.4050/jahs.38.3.73.

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19

Yeager, William T., Paul H. Mirick, M.-Nabil H. Hamouda, Matthew L. Wilbur, Jeffrey D. Singleton, and W. Keats Wilkie. "Rotorcraft Aeroelastic Testing in the Langley Transonic Dynamics Tunnel." Journal of the American Helicopter Society 38, no. 3 (July 1, 1993): 73–82. http://dx.doi.org/10.4050/jahs.38.73.

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20

King, L. S., and D. A. Johnson. "Transonic airfoil calculations including wind tunnel wall-interference effects." AIAA Journal 24, no. 8 (August 1986): 1378–80. http://dx.doi.org/10.2514/3.9448.

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21

Reis, M. L. C. C., J. B. P. Falcão Filho, and L. F. G. Moraes. "The TTP Transonic wind tunnel Mach number uniformity analysis." Measurement 51 (May 2014): 356–66. http://dx.doi.org/10.1016/j.measurement.2014.01.021.

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22

Tabatabaei, Narges, Ramis Örlü, Ricardo Vinuesa, and Philipp Schlatter. "Aerodynamic Free-Flight Conditions in Wind Tunnel Modelling through Reduced-Order Wall Inserts." Fluids 6, no. 8 (July 27, 2021): 265. http://dx.doi.org/10.3390/fluids6080265.

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Parallel sidewalls are the standard bounding walls in wind tunnels when making a wind tunnel model for free-flight condition. The consequence of confinement in wind tunnel tests, known as wall-interference, is one of the main sources of uncertainty in experimental aerodynamics, limiting the realizability of free-flight conditions. Although this has been an issue when designing transonic wind tunnels and/or in cases with large blockage ratios, even subsonic wind tunnels at low-blockage-ratios might require wall corrections if a good representation of free-flight conditions is intended. In order to avoid the cumbersome streamlining methods especially for subsonic wind tunnels, a sensitivity analysis is conducted in order to investigate the effect of inclined sidewalls as a reduced-order wall insert in the airfoil plane. This problem is investigated via Reynolds-averaged Navier–Stokes (RANS) simulations, and a NACA4412 wing at the angles of attack between 0 and 11 degrees at a moderate Reynolds number (400 k) is considered. The simulations are validated with well-resolved large-eddy simulation (LES) results and experimental wind tunnel data. Firstly, the wall-interference contribution in aerodynamic forces, as well as the local pressure coefficients, are assessed. Furthermore, the isolated effect of confinement is analyzed independent of the boundary-layer growth. Secondly, wall-alignment is modified as a calibration parameter in order to reduce wall-interference based on the aforementioned assessment. In the outlined method, we propose the use of linear inserts to account for the effect of wind tunnel walls, which are experimentally simple to realize. The use of these inserts in subsonic wind tunnels with moderate blockage ratio leads to very good agreement between free-flight and wind tunnel data, while this approach benefits from simple manufacturing and experimental realization.

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23

Jelínek, Tomáš, Erik Flídr, Martin Němec, and Jan Šimák. "Test Facility for High-Speed Probe Calibration." EPJ Web of Conferences 213 (2019): 02033. http://dx.doi.org/10.1051/epjconf/201921302033.

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A new test facility was built up as a part of a closed-loop transonic wind tunnel in VZLU´s High-speed Aerodynamics Department. The wind tunnel is driven by a twelve stage radial compressor and Mach and Reynolds numbers can be changed by the compressor speed and by the total pressure in the wind tunnel loop by a set of vacuum pumps, respectively. The facility consists of an axisymmetric subsonic nozzle with an exit diameter de = 100 mm. The subsonic nozzle is designed for regimes up to M = 1 at the nozzle outlet. At the nozzle inlet there is a set of a honeycomb and screens to ensure the flow stream laminar at the outlet of the nozzle. The subsonic nozzle can be supplemented with a transonic slotted nozzle or a supersonic rigid nozzle for transonic and supersonic outlet Mach numbers. The probe is fixed in a probe manipulator situated downstream of the nozzle and it ensures a set of two perpendicular angles in a wide range (±90°). The outlet flow field was measured through in several axial distances downstream the subsonic nozzle outlet. The total pressure and static pressure was measured in the centreline and the total pressure distribution in the vertical and horizontal plane was measured as well. Total pressure fluctuations in the nozzle centreline were detected by a FRAP probe. From the initial flow measurement in a wide range of Mach numbers the best location for probe calibration was chosen. The flow field was found to be suitable for probe calibration.

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24

Henderson, W. P. "Airframe/Propulsion Integration at Transonic Speeds." Journal of Engineering for Gas Turbines and Power 113, no. 1 (January 1, 1991): 51–59. http://dx.doi.org/10.1115/1.2906530.

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A significant level of research is ongoing at NASA’s Langley Research center on integrating the propulsion system with the aircraft. This program has included nacelle/pylon/wing integration for turbofan transports, propeller/nacelle/wing integration for turboprop transports, and nozzle/afterbody/empennage integration for high-performance aircraft. The studies included in this paper focus more specifically on pylon shaping and nacelle bypass ratio studies for turbofan transports, nacelle and wing contouring, and propeller location effects for turboprop transports, empennage effects, and thrust vectoring for high-performance aircraft. The studies were primarily conducted in NASA Langley’s 16-Foot Transonic Tunnel at Mach numbers up to 1.20.

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25

Liu, Guang Yuan, Rui Bo Wang, Chang Rong Zhang, Feng Chen, Jiang Yu Xie, and Shang Ma. "Numerical Investigation on Boundary Layer Flow Control with Vortex Generators." Applied Mechanics and Materials 432 (September 2013): 351–57. http://dx.doi.org/10.4028/www.scientific.net/amm.432.351.

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The numerical simulation method was adopted to analyze the effect on boundary layer thickness reduction of various vortex generator parameters. Results show that vortex generators are capable of reducing boundary layer thickness for about 66 percent, and the influence on centerline Mach number distributions is neglectable. Practicable vortex generators for 2.4m transonic wind tunnel half-model test section side wall are founded. Research results can be used for further applications of vortex generator in wind tunnel tests.

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26

Streit, T., and C. Hoffrogge. "DLR transonic inverse design code, extensions and modifications to increase versatility and robustness." Aeronautical Journal 121, no. 1245 (October 11, 2017): 1733–57. http://dx.doi.org/10.1017/aer.2017.101.

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ABSTRACTThe DLR inverse design code computes the wing geometry for a prescribed target pressure distribution. It is based on the numerical solution of the integral inverse transonic small perturbation (TSP) equations. In this work, several extensions and modifications of the inverse design code are described. Results are validated with corresponding redesign test cases. The first modification concerns applications for high transonic Mach numbers or cases with strong shocks. The introduced modifications enable converged design solutions for cases where the original method failed. The second modification is the extension of the code to general non-planar wings. Previously, the design code was restricted to non-planar wing designs with small dihedral or to nacelle design. A third modification concerns aerofoil/wings designed for wind-tunnel design. In order to design a swept wing between two wind-tunnel walls, the solution method was extended to two symmetry planes. The introduced extensions and modifications have increased the robustness and range of applicability of the inverse design code.

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27

Mitchell, D. A., R. K. Cooper, and S. Raghunathan. "Effect of heat transfer on periodic transonic flows." Aeronautical Journal 103, no. 1025 (July 1999): 329–37. http://dx.doi.org/10.1017/s0001924000064708.

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Abstract The effects of the model surface to free stream adiabatic temperature ratio (Tw/Tad) on periodic transonic flow over a 14% thick biconvex aerofoil are evaluated using a computational fluid dynamic approach. The analysis is based on the thin layer Navier Stokes equations with Baldwin-Lomax turbulence model. The results of computations showed that on biconves aerofoils there is a large effect of heat transfer on instantaneous pressure distributions and periodic buffet excitation level confirming some of the available experimental data. The effects observed have an implication in wind tunnel measurement of buffet associated periodic transonic flows.

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28

Zhang, Zheng Yu, Shui Liang Wang, and Yan Sun. "Videogrammetric Measurement for Model Displacement in Wind Tunnel Test." Applied Mechanics and Materials 130-134 (October 2011): 103–7. http://dx.doi.org/10.4028/www.scientific.net/amm.130-134.103.

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It is crucial measuring position and attitude of model to gain the precise and accurate data in wind tunnel tests. The model displacement videogrammetric measurement (MDVM) system and its key techniques such as the exterior orientation with big rotation angles and large-overlap, mark points, image processing and calibration based on the known distances are therefore presented. The practice example in Asia's largest (2.4m) transonic wind tunnel has demonstrated the MDVM system and its key techniques are correct and feasible, and they have application value.

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29

Rona, Aldo, Renato Paciorri, and Marco Geron. "Design and Testing of a Transonic Linear Cascade Tunnel With Optimized Slotted Walls." Journal of Turbomachinery 128, no. 1 (June 23, 2005): 23–34. http://dx.doi.org/10.1115/1.2101856.

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In linear cascade wind tunnel tests, a high level of pitchwise periodicity is desirable to reproduce the azimuthal periodicity in the stage of an axial compressor or turbine. Transonic tests in a cascade wind tunnel with open jet boundaries have been shown to suffer from spurious waves, reflected at the jet boundary, that compromise the flow periodicity in pitch. This problem can be tackled by placing at this boundary a slotted tailboard with a specific wall void ratio s and pitch angle α. The optimal value of the s-α pair depends on the test section geometry and on the tunnel running conditions. An inviscid two-dimensional numerical method has been developed to predict transonic linear cascade flows, with and without a tailboard, and quantify the nonperiodicity in the discharge. This method includes a new computational boundary condition to model the effects of the tailboard slots on the cascade interior flow. This method has been applied to a six-blade turbine nozzle cascade, transonically tested at the University of Leicester. The numerical results identified a specific slotted tailboard geometry, able to minimize the spurious reflected waves and regain some pitchwise flow periodicity. The wind tunnel open jet test section was redesigned accordingly. Pressure measurements at the cascade outlet and synchronous spark schlieren visualization of the test section, with and without the optimized slotted tailboard, have confirmed the gain in pitchwise periodicity predicted by the numerical model.

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30

Underbrink, James R. "Pletharrays for aeroacoustic phased array applications." International Journal of Aeroacoustics 16, no. 4-5 (July 2017): 202–29. http://dx.doi.org/10.1177/1475472x17718884.

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“Pletharrays” are introduced, motivated, and presented for application to aeroacoustic phased array measurements. Pletharrays contain a plethora of arrays composed from a modest to high number of array elements to field a remarkably large number of high element count arrays for use in noise source imaging applications. Pletharrays that have been deployed for closed jet transonic wind tunnel, static engine ground, open jet wind tunnel, and flyover phased array tests are presented. Tremendous array element leverage to provide extensive measurement flexibility and fidelity are demonstrated.

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31

MITSUO, Kazunori. "Application of Binary PSP Measurement to Transonic Wind Tunnel Testing." Journal of the Visualization Society of Japan 28-1, no. 1 (2008): 101. http://dx.doi.org/10.3154/jvs.28.101.

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32

Lombardi, Giovanni, and Mauro Morelli. "Analysis of some interference effects in a transonic wind tunnel." Journal of Aircraft 32, no. 3 (May 1995): 501–9. http://dx.doi.org/10.2514/3.46748.

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33

Goffert, Bruno, Marcos Aurélio Ortega, and João Batista Pessoa Falcão Filho. "Numerical Study of Wall Ventilation in a Transonic Wind Tunnel." Journal of Aerospace Technology and Management 7, no. 1 (February 22, 2015): 81–92. http://dx.doi.org/10.5028/jatm.v7i1.417.

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MARU, Yusuke, Hiroki TAKAYANAGI, Kazuhiko YAMADA, and Kazuhisa FUJITA. "Wind Tunnel Testing of Parachutes at Transonic and Supersonic Speed." Proceedings of Mechanical Engineering Congress, Japan 2016 (2016): S1910104. http://dx.doi.org/10.1299/jsmemecj.2016.s1910104.

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35

Cole, Stanley R., Thomas E. Noll, and Boyd Perry. "Transonic Dynamics Tunnel Aeroelastic Testing in Support of Aircraft Development." Journal of Aircraft 40, no. 5 (September 2003): 820–31. http://dx.doi.org/10.2514/2.6873.

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36

Azzazy, M., D. Modarress, and R. M. Hall. "Optical boundary-layer transition detection in a transonic wind tunnel." AIAA Journal 27, no. 4 (April 1989): 405–10. http://dx.doi.org/10.2514/3.10127.

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37

Fico, Nide G. C. R., and Marcos A. Ortega. "Numerical prediction of flap losses in a transonic wind tunnel." AIAA Journal 31, no. 1 (January 1993): 133–39. http://dx.doi.org/10.2514/3.11329.

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38

Green, John, and Jürgen Quest. "A short history of the European Transonic Wind Tunnel ETW." Progress in Aerospace Sciences 47, no. 5 (July 2011): 319–68. http://dx.doi.org/10.1016/j.paerosci.2011.06.002.

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39

Betyaev, S. K. "Flow in the working section of a transonic wind tunnel." Journal of Engineering Physics and Thermophysics 84, no. 2 (March 2011): 402–7. http://dx.doi.org/10.1007/s10891-011-0485-9.

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40

Rašuo, B. "On Results' Accuracy at Two-Dimensional Transonic Wind Tunnel Testing." PAMM 2, no. 1 (March 2003): 306–7. http://dx.doi.org/10.1002/pamm.200310137.

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41

Tateishi, Atsushi, Toshinori Watanabe, Takehiro Himeno, and Seiji Uzawa. "Numerical method for an assessment of steady and motion-excited flowfields in a transonic cascade wind tunnel." Journal of the Global Power and Propulsion Society 1 (August 25, 2017): QL9XVI. http://dx.doi.org/10.22261/ql9xvi.

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AbstractThis article presents a numerical method and its application for an assessment of the flow field inside a wind tunnel. A structured computational fluid dynamics (CFDs) solver with overset mesh technique is developed in order to simulate geometrically complex configurations. Applying the developed solver, a whole transonic cascade wind tunnel is modeled and simulated by a two-dimensional manner. The upstream and downstream periodicity of the cascade and the effect of the tunnel wall on the unsteady flow field are focused on. From the steady flow simulations, the existence of an optimum throttle position for the best periodicity for each tailboard angle is shown, which provides appropriate aerodynamic characteristics of ideal cascades in the wind tunnel environment. Unsteady simulations with blade oscillation is also conducted, and the difference in the influence coefficients between ideal and wind tunnel configurations becomes large when the pressure amplitude increases on the lower blades.

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42

Srivastava, Ankur, and Andrew J. Meade. "Use of Active Learning to Design Wind Tunnel Runs for Unsteady Cavity Pressure Measurements." International Journal of Aerospace Engineering 2014 (2014): 1–11. http://dx.doi.org/10.1155/2014/218710.

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Wind tunnel tests to measure unsteady cavity flow pressure measurements can be expensive, lengthy, and tedious. In this work, the feasibility of an active machine learning technique to design wind tunnel runs using proxy data is tested. The proposed active learning scheme used scattered data approximation in conjunction with uncertainty sampling (US). We applied the proposed intelligent sampling strategy in characterizing cavity flow classes at subsonic and transonic speeds and demonstrated that the scheme has better classification accuracies, using fewer training points, than a passive Latin Hypercube Sampling (LHS) strategy.

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43

Franzmann, Christian, Friedrich Leopold, and Christian Mundt. "Low-Interference Wind Tunnel Measurement Technique for Pitch Damping Coefficients at Transonic and Low Supersonic Mach Numbers." Aerospace 9, no. 2 (January 20, 2022): 51. http://dx.doi.org/10.3390/aerospace9020051.

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An experimental method for the determination of the pitch damping moment coefficient sum Cmq+Cmα˙ in a wind tunnel at transonic and low supersonic Mach numbers is developed. With support interference being a major issue for dynamic tests at these velocities, a minimum interference wire suspension approach is used. The motion of the wind tunnel model is restricted to a single-degree of freedom pitching oscillation through the geometry of the support system. A statistical evaluation procedure allows the simultaneous evaluation of multiple tests to increase confidence in the results. The influence of the wires as well as nonlinear effects are accounted for. The method is validated in an extensive test series at Mach numbers ranging from 0.6 to 2.0. Two reference missile models—the Basic Finner and the Army-Navy Spinner Rocket (ANSR)—are used. The results agree very well with CFD calculations throughout the transonic range. In comparison to free-flight tests the accuracy is significantly improved and result uncertainties are reduced by an order of magnitude.

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44

Biryukov, V. I., S. A. Glazkov, A. R. Gorbushin, A. I. Ivanov, and A. V. Semenov. "Experimental investigation of the effect of nozzle shape and test section perforation on the stationary and non-stationary characteristics of flow field in the large transonic TsAGI T-128 Wind tunnel." Aeronautical Journal 109, no. 1092 (February 2005): 75–82. http://dx.doi.org/10.1017/s0001924000000579.

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Abstract The results are presented for a cycle of experimental investigations of flow field characteristics (static pressure distribution, static pressure fluctuations, upwash, boundary-layer parameters) in the perforated test section of the transonic TsAGI T-128 Wind Tunnel. The investigations concern the effect of nozzle shape, wall open-area ratio, Mach and Reynolds numbers on the above-outlined flow characteristics. During the tests, the main Wind-tunnel drive power is measured. Optimal parameters of the nozzle shape and test section perforation are obtained to minimise acoustic perturbations in the test section and their non-uniformity in frequency, static pressure field non-uniformity, nozzle and test section drag and, accordingly, required main Wind-tunnel drive power.

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45

Zhang, Yong Cun, Liu Sheng Chen, Jin Kui Shang, Xiao Guang Ma, Xue Yuan Chen, Li Yan, Wen Tao Zhao, Xue Ying Deng, and Hou Mei Cheng. "Application of Two-Component Pressure Sensitive Paint in Transonic Wind Tunnel." Advanced Materials Research 216 (March 2011): 181–87. http://dx.doi.org/10.4028/www.scientific.net/amr.216.181.

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As a new optical pressure sensor technique, Pressure Sensitive Paint (PSP) is one of the important techn iques for model surface p ressure m easurement in wind tunnel test s . With the help of PSP , it is possible to do p ressure m easurement on compl icated or special model surface , which is usually difficult to be measured by pressure tap s . Since PSP technique being introduced into China from TsAGI (Russia) , AVIC ARI has investigated two-component PSP technique in high-speed wind tunnel in cooperati on with ICCAS China . T his report present s the principle of PSP technique, test control system development and the test result comparison s between PSP technique with two-component pressure sensitive paint FOP-2 and classic tap measurement on wing surface of an airplane model . T he results showed that the two-component pressure sensitive paint has better performance and can be used for model pressure measurement.

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46

Lepicovsky, J., R. V. Chima, E. R. McFarland, and J. R. Wood. "On Flowfield Periodicity in the NASA Transonic Flutter Cascade." Journal of Turbomachinery 123, no. 3 (February 1, 2000): 501–9. http://dx.doi.org/10.1115/1.1378300.

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A combined experimental and numerical program was carried out to improve the flow uniformity and periodicity in the NASA transonic flutter cascade. The objectives of the program were to improve the periodicity of the cascade and to resolve discrepancies between measured and computed flow incidence angles and exit pressures. Previous experimental data and some of the discrepancies with computations are discussed. In the present work surface pressure taps, boundary layer probes, shadowgraphs, and pressure-sensitive paints were used to measure the effects of boundary layer bleed and tailboard settings on flowfield periodicity. These measurements are described in detail. Two numerical methods were used to analyze the cascade. A multibody panel code was used to analyze the entire cascade and a quasi-three-dimensional viscous code was used to analyze the isolated blades. The codes are described and the results are compared to the measurements. The measurements and computations both showed that the operation of the cascade was heavily dependent on the endwall configuration. The endwalls were redesigned to approximate the midpassage streamlines predicted using the viscous code, and the measurements were repeated. The results of the program were that: (1) Boundary layer bleed does not improve the cascade flow periodicity. (2) Tunnel endwalls must be shaped like predicted cascade streamlines. (3) The actual flow incidence must be measured for each cascade configuration rather than using the tunnel geometry. (4) The redesigned cascade exhibits excellent periodicity over six of the nine blades.

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Yu, Li, Bin Bin Lv, Hong Tao Guo, Yu Yan, Xing Hua Yang, and Jian Guo Luo. "Research on Transonic Wind Tunnel Flutter Test for a Wing Model." Advanced Materials Research 1006-1007 (August 2014): 26–29. http://dx.doi.org/10.4028/www.scientific.net/amr.1006-1007.26.

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This paper adopts self-designed wing model to conduct flutter test on subsonic and transonic, and obtains flutter characteristic of the model, and the test results are used for calibration and verification of flutter procedures. The sub-critical extrapolation is used to obtain the flutter sub-critical parameters and the direct observation method is used to obtain comparison of results. Error of results obtained by the two approaches does not exceed 5%, and validates reliability of the sub-critical prediction approach in continuous adjusted dynamic pressure flutter test.

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48

Yin, Gang, Hong Jiang, Yan Xie, Ping Li, Li Zhao, and Zhenhua Yang. "Vibration effect correction method of inclinometer in intermittent transonic wind tunnel." Sensors and Actuators A: Physical 331 (November 2021): 112938. http://dx.doi.org/10.1016/j.sna.2021.112938.

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49

Falcão Filho, João Batista Pessoa, Marcos Aurélio Ortega, and Luiz Carlos Sandoval Góes. "Prediction of transients and control reactions in a transonic wind tunnel." Journal of the Brazilian Society of Mechanical Sciences 22, no. 2 (2000): 317–39. http://dx.doi.org/10.1590/s0100-73862000000200014.

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50

Ng, W. F., M. Gundappa, D. O. Griffith, and J. B. Peterson. "Turbulence measurements and noise generation in a transonic cryogenic wind tunnel." AIAA Journal 28, no. 5 (May 1990): 853–58. http://dx.doi.org/10.2514/3.25129.

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