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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15 Mar 2007
TL;DR: In this article, a transsonique souffleries with a couche couche is described. Reference Record created on 2005-11-18, modified on 2016-08-08
Abstract: Keywords: transsonique ; souffleries ; couche : limite ; ecoulement ; avions ; modeles ; essais ; grille Reference Record created on 2005-11-18, modified on 2016-08-08
87 citations
TL;DR: In this paper, an explicit finite-difference method with time splitting is used to solve the time-dependent equations for compressible turbulent flow, and a nonorthogonal computational mesh of arbitrary configuration facilitates the description of the flow field.
Abstract: A code has been developed for simulating high Reynolds number transonic flow fields of arbitrary configuration. An explicit finite-difference method with time splitting is used to solve the time-dependent equations for compressible turbulent flow. A nonorthogonal computational mesh of arbitrary configuration facilitates the description of the flow field. The code is applied to simulate the flow over an 18 percent thick circular-arc biconvex airfoil at zero angle of attack and free-stream Mach number of 0.775. A simple mixing-length model is used to describe the turbulence and chord Reynolds numbers of 1, 2, 4, and 10 million are considered. The solution describes in sufficient detail both the shock-induced and trailing-edge separation regions, and provides the profile and friction drag.
81 citations
TL;DR: In this article, the transonic small disturbance equation is solved for flow past thin lifting airfoils and slender bodies with M^ < 1, including cases with imbedded shock waves.
Abstract: Solutions of the transonic small disturbance equation are presented for flow past thin lifting airfoils and slender bodies with M^ < 1, including cases with imbedded shock waves. The results are obtained numerically using a mixed finite-difference relaxation method previously reported by the authors. Results are presented for four lifting airfoils at various angles of attack and are compared with shock free theory and experimental data. For the slender body case, comparisons with experiments are given for five geometries both with and without aft stings. The results are also compared with approximate theory. Discussion is given on the treatment of the boundary conditions, computing times and accuracies, and ranges of applicability of the small disturbance theory.
81 citations
69 citations
01 Dec 1975
TL;DR: In this paper, the use of hot-wire anemometers for obtaining fluctuating data in transonic flows has been evaluated from hotwire heat loss correlations based on previous transonic data, the sensitivity coefficients for velocity, density and total temperature fluctuations have been calculated for a wide range of test conditions and sensor parameters.
Abstract: The use of hot-wire anemometry for obtaining fluctuating data in transonic flows has been evaluated From hot-wire heat loss correlations based on previous transonic data, the sensitivity coefficients for velocity, density, and total temperature fluctuations have been calculated for a wide range of test conditions and sensor parameters For sensor Reynolds numbers greater than 20 and high sensor overheat ratios, the velocity sensitivity remains independent of Mach number and equal to the density sensitivity These conclusions were verified by comparisons of predicted sensitivities with those from recent direct calibrations in transonic flows Based on these results, techniques are presented to obtain meaningful measurements of fluctuating velocity, density, and Reynolds shear stress using hot-wire and hot-film anemometers Examples of these measurements are presented for two transonic boundary layers
69 citations