Experimental testing at transonic speeds
01 Jan 1982-Vol. 81, pp 189-238
TL;DR: In this paper, the authors defined the normal force and pitch moment coefficients of a wing span and its chord, and the slope of a section normal force vs angle-of-a attack curve.
Abstract: Nomenclature b = wing span c = wing chord c , = section drag coefficient d c = section lift coefficient c = section normal force coefficient n c = slope of section normal force vs angle-ofa attack curve c = section pitching moment coefficient C = aircraft drag coefficient C = aircraft lift coefficient LI C = pressure coefficient Cju = blowing coefficient M = Mach number P = total pressure q , q = freestream dynamic pressure
Citations
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TL;DR: In this article, the transonic flow around a low-aspect-ratio infinite wing in a wind tunnel is investigated by means of numerical simulations, with a focus on shock-induced separated flows.
19 citations
01 Jan 1984
TL;DR: In this article, an investigation was carried out on two CAST 10-2 airfoil models with chords of 3 in. and 6 in. The tests were conducted in the Langley 0.3m Transonic Cryogenic Tunnel two-dimensional test section equipped with an upstream sidewall boundary layer removal system.
Abstract: An investigation was carried out on two CAST 10-2 airfoil models with chords of 3 in. and 6 in. To evaluate the extent of sidewall influence on airfoil tests at transonic Mach numbers. The tests were conducted in the Langley 0.3-m Transonic Cryogenic Tunnel two-dimensional test section equipped with an upstream sidewall boundary layer removal system which reduces the boundary layer displacement thickness to about 1 percent of model halfspan from an initial 2 percent without boundary layer removal. Test results have shown the changes in the location of the shock on the upper surface of the airfoil to be about the same for both models with and without sidewall boundary layer removal. Even though large differences were noted in the high lift characteristics of the two models, the sidewall boundary layer removal had little effect on the differences. These tests also served to validate the boundary layer removal technique and the associated Mach number correction required with upstream boundary layer removal.
11 citations
15 Jan 1979
1 citations
22 Oct 2019
TL;DR: This revision of the model representation is based on high fidelity free-air CFD solutions for the NASA Common Research Model, validated by experimental data sets obtained from a semispan version of the CRM tested at the NRC 5ft Trisonic Wind Tunnel.
Abstract: The National Research Council 5ft Trisonic Wind Tunnel enables testing of half-span models at a high Reynolds number cost effectively, however there is the possibility of experiencing relatively larger wall-interference effects compared to full span tests. Subsonic wall-interference effects due to the partially open boundaries at the NRC 5ft wind tunnel are corrected by the “one-variable” method, resulting in corrections in Mach number and angle of attack for the measured quantities. This dissertation introduces the need for adequately correcting for wall interference in subsonic half-model wind-tunnel testing in both pre-stall and post-stall conditions. The one-variable method is described and several aspects that require validation and improvement in this method are pointed out based on specific experimental and numerical results obtained for different test articles.
An initial assessment of the influence of the different possible tunnel configurations and computational parameters is defined based on experimental data from a test performed on a scaled model of the Bombardier Global 6000 business jet. This initial study unveils some of the weaknesses of the one-variable method. While this method provides accurate wall corrections during tests in pre-stall conditions; it is unable to generate reliable corrections in stall. Partially, this is due to the development of flow separation on a model tested at subsonic flow conditions reaching stall, not currently accounted for by the potential theory model representation used by the wall-correction methodology at the NRC 5ft TWT. Possible improvements to this singularity representation, paramount to the one-variable method, are then investigated. This revision of the model representation is based on high fidelity free-air CFD solutions for the NASA Common Research Model, validated by experimental data sets obtained from a semispan version of the CRM tested at the NRC 5ft Trisonic Wind Tunnel. The improved potential representation is tested on solid-wall experimental wind-tunnel data to present more reliable behavior of the wall-interference correction estimates when substantial wing stalling is encountered.
01 Jan 1987
TL;DR: In this paper, a simplified fourwall interference assessment method has been described, and a computer program developed to facilitate correction of the airfoil data obtained in the Langley 0.3m Transonic Cryogenic Tunnel (TCT).
Abstract: A simplified fourwall interference assessment method has been described, and a computer program developed to facilitate correction of the airfoil data obtained in the Langley 0.3-m Transonic Cryogenic Tunnel (TCT). The procedure adopted is to first apply a blockage correction due to sidewall boundary-layer effects by various methods. The sidewall boundary-layer corrected data are then used to calculate the top and bottom wall interference effects by the method of Capallier, Chevallier and Bouinol, using the measured wall pressure distribution and the model force coefficients. The interference corrections obtained by the present method have been compared with other methods and found to give good agreement for the experimental data obtained in the TCT with slotted top and bottom walls.
References
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01 Jan 1975
10 citations
TL;DR: In this article, a technique has been devised for sensing the instantaneous location of a normal shock wave as it oscillates about a mean position based on hot-film anemometer technology.
Abstract: A technique has been devised for sensing the instantaneous location of a normal shock wave as it oscillates about a mean position Based on hot-film anemometer technology, the method involves a probe that is operated by a standard constant-temperature anemometer unit Heat transfer from the probe's heated metal film is affected by the presence of a normal shock, which causes boundary-layer transition As a consequence, the output voltage of the anemometer unit varies according to the location of the shock along the metal film Sensitivity of the shock-position-sensing probe is shown to be in agreement with a simple two-dimensional flow model The shock probe was employed during a transonic airfoil experiment to provide amplitude and frequency information on shock motions, and also to determine the relationship between shock motions and lift fluctuations in buffeting These results have shown that the hot-film shock-position-sensing probe is a simple, effective instrument for studying the behavior of normal shock waves in unsteady flow
6 citations
TL;DR: Prospects in aeronautics research and development, noting socio-economic impact neglect and air transportation effectiveness lag lag, are discussed in this paper, where the authors highlight the socioeconomic impact of air transportation.
Abstract: Prospects in aeronautics research and development, noting socio-economic impact neglect and air transportation effectiveness lag
5 citations
01 Jan 1978
TL;DR: A survey of research relative to scale effects on supercritical airfoils has been conducted in this paper, which indicated that Reynolds number scale effects have a significant impact on airfoil design and performance.
Abstract: A survey of research relative to scale effects on supercritical airfoils has been conducted. The results of this survey indicated that Reynolds number scale effects have a significant impact on airfoil design and performance. Further, this impact is greater for supercritical airfoils than for conventional airfoils. It was found that low Reynolds number drag data could be extrapolated to high Reynolds number conditions provided the flow was attached and the pressure distribution shape did not change appreciably. Airfoil lift and pitching-moment data obtained at low Reynolds numbers cannot be extrapolated to full-scale values. Viscous theoretical transonic analysis methods currently under development will significantly improve the ability of the designer to account for scale effects. Boundary-layer manipulation in low Reynolds number facilities using natural transition or aft located transition strips to simulate high Reynolds number conditions was shown to be an uncertain test procedure and reliance should be made on high Reynolds number facilities if available.
2 citations