scispace - formally typeset
Search or ask a question
Topic

Freestream

About: Freestream is a research topic. Over the lifetime, 3428 publications have been published within this topic receiving 56147 citations.


Papers
More filters
Proceedings ArticleDOI
07 Jun 1999
TL;DR: In this article, pressure sensitive paint (PSP) was used to measure film cooling effectiveness on a turbine nozzle surface in a high speed wind tunnel, and the effects of secondary flow and freestream Mach number and Reynolds number on turbine nozzle suction surface film cooling were discussed.
Abstract: The use of pressure sensitive paint (PSP) to measure film cooling effectiveness on a turbine nozzle surface was demonstrated in a high speed wind tunnel. Film cooling effectiveness was measured from a single row of holes located on a turbine vane suction surface with a shaped exit. Nitrogen gas was used to simulate film cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Three blowing ratios were studied for each of the five freestream conditions: a reference condition, a reduced and an increased Reynolds number condition, and a reduced and an increased Mach number condition. The freestream turbulence intensity was kept at 12.0% for all the tests. The PSP was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the airfoil surface. The film effectiveness increased with blowing ratio for all the freestream conditions. The effects of secondary flow and freestream Mach number and Reynolds number on turbine nozzle suction surface film cooling are also discussed.Copyright © 1999 by ASME

37 citations

Journal ArticleDOI
TL;DR: In this paper, numerical simulations were carried out to determine the sensitivity of results to a variety of geometric and flow parameters commonly employed in high-speed transverse jet-interaction calculations, including turbulence model, freestream turbulence intensity, turbulent Schmidt number, and several injector-pipe configurations.
Abstract: Numerical simulations were carried out to determine the sensitivity of results to a variety of geometric and flow parameters commonly employed in high-speed transverse jet-interaction calculations. The configuration consisted of a single circular, flush-wall porthole injector inclined at 30 deg to the freestream in a Mach 4.0 crossflow. Injection was sonic with a jet-to-freestream momentum flux ratio of 2.1. The primary modeling parameters investigated include turbulence model, freestream turbulence intensity, turbulent Schmidt number, and several injector-pipe configurations. The simulations were conducted using the multispecies Reynolds-averaged Navier–Stokes equations with a number of popular turbulence models including the one-equation Spalart–Allmaras model and the two-equation Menter shear stress transport, two-equation realizable k-e, and two-equation nonlinear (cubic) k-e models. The results were found to be very sensitive to both the choice of turbulence model and value of the turbulent Schmidt n...

37 citations

Proceedings ArticleDOI
01 Jun 1992
TL;DR: In this article, a full-scale F/A-18 was tested in the National Full-Scale Aerodynamic Complex at NASA Ames Research Center at Moffett Field, California, at an angle of attack range of 18 to 50 degrees and at wind speeds of up to 100 knots.
Abstract: This paper presents an overview of high angle-of-attack tests of a full-scale F/A-18 in the 80- by 120-Foot Wind Tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center at Moffett Field, California. A production aircraft was tested over an angle-of-attack range of 18 to 50 deg and at wind speeds of up to 100 knots. These tests had three primary test objectives. Pneumatic and mechanical forebody flow control devices were tested at full-scale and shown to produce significant yawing moments for lateral control of the aircraft at high angles of attack. Mass flow requirements for the pneumatic system were found to scale with freestream density and speed rather than freestream dynamic pressure. Detailed measurements of the pressures buffeting the vertical tail were made and spatial variations in the buffeting frequency were found. The LEX fence was found to have a significant effect on the frequency distribution on the outboard surface of the vertical fin. In addition to the above measurements, an extensive set of data was acquired for the validation of computational fluid dynamics codes and for comparison with flight test and small-scale wind tunnel test results.

37 citations

Journal ArticleDOI
TL;DR: In this article, a shape optimization methodology for reducing the initial shock pressure rise on the ground of a supersonic aircraft is presented, which combines elements from the linearized aerodynamic theory, such as the Whitham F function, with elements from nonlinear aerodynamic theories such as prediction of lift distribution by an Euler or a Navier-Stokes flow solver.
Abstract: A shape optimization methodology for reducing the initial shock pressure rise on the ground of a supersonic aircraft is presented. This methodology combines elements from the linearized aerodynamic theory, such as the Whitham F function, with elements from the nonlinear aerodynamic theory, such as the prediction of lift distribution by an Euler or a Navier-Stokes flow solver. It is applied to the optimization of two different airplane concepts developed by Reno Aeronautical and Lockheed Martin, respectively, for the Defense Advanced Research Projects Agency's Quiet Supersonic Platform program. For Reno Aeronautical's laminar-flow supersonic aircraft, the initial shock pressure rise on the ground is reduced by a factor close to 2, from 1.224 psf (58.605 N/m 2 ) at a freestream Mach number of 1.5 to 0.671 psf (32.127 N/m 2 ), while maintaining constant lift For Lockheed Martin's point of departure aircraft, a tenfold reduction of the initial shock pressure rise on the ground is demonstrated, from 1.623 psf (77.71 N/m 2 ) at a freestream Mach number of 1.5 to 0.152 psf (7.278 N/m 2 ), also while maintaining constant lift.

36 citations

Proceedings ArticleDOI
08 May 2000
TL;DR: In this article, a transonic cascade wind tunnel was used to investigate the film effectiveness and heat transfer coefficient on the suction side of a high-turning turbine rotor blade.
Abstract: Experiments were performed in a transonic cascade wind tunnel to investigate the film effectiveness and heat transfer coefficient on the suction side of a high-turning turbine rotor blade. The coolant scheme consisted of six rows of staggered, discrete cooling holes on and near the leading edge of the blade in a showerhead configuration. Air was cooled in order to match the density ratios found under engine conditions. Six high-frequency heat flux gauges were installed downstream of the cooling holes on the suction side of the blade. Experiments were performed with and without film and the coolant to freestream total pressure ratio was varied from 1.02 to 1.19. In order to simulate real engine flow conditions, the exit Mach number was set to 1.2 and the exit Reynolds number was set to 5×106. The freestream turbulence was approximately 1%. The heat transfer coefficient was found to increase with the addition of film cooling an average of 14% overall and to a maximum of 26% at the first gauge location. The average film cooling effectiveness over the gauge locations was 25%. Both the heat transfer coefficient and the film cooling effectiveness were found to have only a weak dependence upon the coolant to freestream total pressure ratio at the gauge locations used in this study.© 2000 ASME

36 citations


Network Information
Related Topics (5)
Reynolds number
68.4K papers, 1.6M citations
87% related
Boundary layer
64.9K papers, 1.4M citations
84% related
Turbulence
112.1K papers, 2.7M citations
81% related
Laminar flow
56K papers, 1.2M citations
81% related
Nozzle
158.6K papers, 893K citations
79% related
Performance
Metrics
No. of papers in the topic in previous years
YearPapers
2023195
2022350
2021108
2020113
201986
2018118