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  • 1
    Publication Date: 2019-06-28
    Description: A computational investigation was conducted to support the development of a semi-span model test capability in the NASA LaRC's National Transonic Facility. This capability is required for the testing of high-lift systems at flight Reynolds numbers. A three-dimensional Navier-Stokes solver was used to compute the low-speed flow over both a full-span configuration and a semi-span configuration. The computational results were found to be in good agreement with the experimental data. The computational results indicate that the stand-off height has a strong influence on the flow over a semi-span model. The semi-span model adequately replicates the aerodynamic characteristics of the full-span configuration when a small stand-off height, approximately twice the tunnel empty sidewall boundary layer displacement thickness, is used. Several active sidewall boundary layer control techniques were examined including: upstream blowing, local jet blowing, and sidewall suction. Both upstream tangential blowing, and sidewall suction were found to minimize the separation of the sidewall boundary layer ahead of the semi-span model. The required mass flow rates are found to be practicable for testing in the NTF. For the configuration examined, the active sidewall boundary layer control techniques were found to be necessary only near the maximum lift conditions.
    Keywords: Aerodynamics
    Type: NASA-CR-4709 , NAS 1.26:4709
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  • 2
    Publication Date: 2019-06-28
    Description: A three dimensional Navier-Stokes solver is evaluated for transonic flow over a thin, swept, low-aspect ratio wing. The computational study was undertaken in support of a wind tunnel experimental program. The computational results are compared to experimental surface pressure data obtained in a cryogenic wind tunnel with an adaptive wall test section. The results show favorable agreement over a wide range of conditions, further the numerical results provide additional data of the complex three-dimensional flow field. Differences in the predictions and experiment suggest a need to conduct further experiments to evaluate the adaptive wall testing technique, and to model the tunnel sidewall boundary layer in the computations.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1725
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  • 3
    Publication Date: 2019-06-28
    Description: A research program was conducted at NASA Langley Research Center to build and test a thin, pressure instrumented wing. The wing chosen was the canard of the X-29, which has a maximum thickness of 5 percent of chord. The wing has 90 pressure taps and was built utilizing an advanced laminated metal technique. It was tested in the 0.3-Meter Transonic Cryogenic Tunnel at transonic Mach numbers and over a wide range of Reynolds number. The data are compared with flight data and Navier-Stokes computational results.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 91-1626
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  • 4
    Publication Date: 2019-06-28
    Description: A numerical investigation of the interaction between a wind tunnel sidewall boundary layer and a thin low-aspect-ratio wing has been performed for transonic speeds and flight Reynolds numbers. A three-dimensional Navier-Stokes code was applied to calculate the flowfields. The results indicated that the sidewall boundary layer had a strong influence on the flowfield around the wing. The computed wing pressure distributions showed vast improvements over previous free-air computations, and were in excellent agreement with experimental data. The low momentum of the sidewall boundary layer resulted in higher pressures in the juncture region, which decreased the favorable spanwise pressure gradient. This significantly decreased the spanwise migration of the wing boundary layer. Weak vortices were predicted in both the upper and lower surface juncture regions. These vortices are believed to have been generated by lateral skewing of the streamlines in the approaching boundary layer.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 92-4036
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  • 5
    Publication Date: 2019-06-28
    Description: A semi-span testing technique has been proposed for the NASA Langley Research Center's National Transonic Facility (NTF). Semi-span testing has several advantages including (1) larger model size, giving increased Reynolds number capability; (2) improved model fidelity, allowing ease of flap and slat positioning which ultimately improves data quality; and (3) reduced construction costs compared with a full-span model. In addition, the increased model size inherently allows for increased model strength, reducing aeroelastic effects at the high dynamic pressure levels necessary to simulate flight Reynolds numbers. The Energy Efficient Transport (EET) full-span model has been modified to become the EET semi-span model. The full-span EET model was tested extensively at both NASA LRC and NASA Ames Research Center. The available full-span data will be useful in validating the semi-span test strategy in the NTF. In spite of the advantages discussed above, the use of a semi-span model does introduce additional challenges which must be addressed in the testing procedure. To minimize the influence of the sidewall boundary layer on the flow over the semi-span model, the model must be off-set from the sidewall. The objective is to remove the semi-span model from the sidewall boundary layer by use of a stand-off geometry. When this is done however, the symmetry along the centerline of the full-span model is lost when the semi-span model is mounted on the wind tunnel sidewall. In addition, the large semi-span model will impose a significant pressure loading on the sidewall boundary layer, which may cause separation. Even under flow conditions where the sidewall boundary layer remains attached, the sidewall boundary layer may adversely effect the flow over the semi-span model. Also, the increased model size and sidewall mounting requires a modified wall correction strategy. With these issues in mind, the semi-span model has been well instrumented with surface pressure taps to obtain data on the expected complex flow field in the near wall region. This status report summarizes the progress to date on developing the semi-span geometry definition suitable for generating structured grids for the computational research. In addition, the progress on evaluating three state-of-the-art Navier-Stokes codes is presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-194479 , NAS 1.26:194479
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  • 6
    Publication Date: 2019-07-13
    Description: Computational Fluid Dynamics based design methods are maturing to the point that they are beginning to be used in the aircraft design process. Many design methods however have demonstrated deficiencies in the leading edge region of airfoil sections. The objective of the present research is to develop an efficient inverse design method which is valid in the leading edge region. The new design method is a streamline curvature method, and a new technique is presented for modeling the variation of the streamline curvature normal to the surface. The new design method allows the surface coordinates to move normal to the surface, and has been incorporated into the Constrained Direct Iterative Surface Curvature (CDISC) design method. The accuracy and efficiency of the design method is demonstrated using both two-dimensional and three-dimensional design cases.
    Keywords: Aerodynamics
    Type: AIAA Paper 2000-0780 , 38th Aerospace Sciences Meeting and Exhibit; Jan 10, 2000 - Jan 13, 2000; Reno, NV; United States
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  • 7
    Publication Date: 2019-07-13
    Description: The objectives of the present research are: (1) to develop a computational approach to support semi-span model test techniques in the NTF; and (2) to integrate this approach with the conduct of an experimental test program. To meet these objectives, the following approach is taken. A state-of-the-art three-dimensional Navier-Stokes solver is employed to compute the flow over both a full-span configuration and a semi-span configuration mounted on the sidewall of the tunnel. The computations are validated by making direct comparisons to experimental data for both configurations. Then, the semi-span computational results are compared to the full-span results to document how the flow over the semi-span configuration differs from that over the full-span configuration. The results of this comparative study will be used to provide a conceptual framework within which a semi-spa model test technique may be implemented in the NTF.
    Keywords: AERODYNAMICS
    Type: NASA-CR-199272 , NAS 1.26:199272
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  • 8
    Publication Date: 2019-07-13
    Description: A second wind tunnel test of the FAST-MAC circulation control semi-span model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged flap. The model was configured for transonic testing of the cruise configuration with 0deg flap deflection to determine the potential for drag reduction with the circulation control blowing. Encouraging results from analysis of wing surface pressures suggested that the circulation control blowing was effective in reducing the transonic drag on the configuration, however this could not be quantified until the thrust generated by the blowing slot was correctly removed from the force and moment balance data. This paper will present the thrust removal methodology used for the FAST-MAC circulation control model and describe the experimental measurements and techniques used to develop the methodology. A discussion on the impact to the force and moment data as a result of removing the thrust from the blowing slot will also be presented for the cruise configuration, where at some Mach and Reynolds number conditions, the thrust-removed corrected data showed that a drag reduction was realized as a consequence of the blowing.
    Keywords: Aerodynamics
    Type: AIAA Paper-2014-2402 , NF1676L-17629 , AIAA Applied Aerodynamics Conference; Jun 16, 2014 - Jun 20, 2014; Atlanta, GA.; United States
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  • 9
    Publication Date: 2019-07-13
    Description: A second wind tunnel test of the FAST-MAC circulation control model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 and 60 degrees, along with the transonic cruise configuration with zero degree flap deflection. Testing was again conducted over a wide range of Mach numbers up to 0.88, and Reynolds numbers up to 30 million based on the mean chord. The first wind tunnel test had poor transonic force and moment data repeatability at mild cryogenic conditions due to inadequate thermal conditioning of the balance. The second test demonstrated that an improvement to the balance heating system significantly improved the transonic data repeatability, but also indicated further improvements are still needed. The low-speed highlift performance of the model was improved by testing various blowing slot heights, and the circulation control was again demonstrated to be effective in re-attaching the flow over the wing at off-design transonic conditions. A new tailored spanwise blowing technique was also demonstrated to be effective at transonic conditions with the benefit of reduced mass flow requirements.
    Keywords: Aerodynamics
    Type: NF1676L-15676 , AIAA Fluid Dynamics Conference and Exhibit; Jun 24, 2013 - Jun 27, 2013; San Diego, CA; United States
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  • 10
    Publication Date: 2019-07-13
    Description: A new capability to test active flow control concepts and propulsion simulations at high Reynolds numbers in the National Transonic Facility at the NASA Langley Research Center is being developed. The first active flow control experiment was completed using the new FAST-MAC semi-span model to study Reynolds number scaling effects for several circulation control concepts. Testing was conducted over a wide range of Mach numbers, up to chord Reynolds numbers of 30 million. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. Preliminary analysis of the uncorrected lift data showed that the circulation control increased the low-speed maximum lift coefficient by 33%. At transonic speeds, the circulation control was capable of positively altering the shockwave pattern on the upper wing surface and reducing flow separation. Furthermore, application of the technique to only the outboard portion of the wing demonstrated the feasibility of a pneumatic based roll control capability.
    Keywords: Fluid Mechanics and Thermodynamics
    Type: AIAA Paper 2012-0103 , NF1676L-13961 , 50th AIAA Aerospace Sciences Meeting and Exhibit; Jan 09, 2012 - Jan 12, 2012; Nashville, TN; United States
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