Measurements of Pressure and Local Heat Transfer on a 20° Cone at Angles of Attack Up to 20° for a Mach Number of 4.95 PDF Download
Are you looking for read ebook online? Search for your book and save it on your Kindle device, PC, phones or tablets. Download Measurements of Pressure and Local Heat Transfer on a 20° Cone at Angles of Attack Up to 20° for a Mach Number of 4.95 PDF full book. Access full book title Measurements of Pressure and Local Heat Transfer on a 20° Cone at Angles of Attack Up to 20° for a Mach Number of 4.95 by Jerome D. Julius. Download full books in PDF and EPUB format.
Author: Jerome D. Julius
Publisher:
ISBN:
Category : Heat
Languages : en
Pages : 32
Get Book Here
Book Description
Author: Jerome D. Julius
Publisher:
ISBN:
Category : Heat
Languages : en
Pages : 32
Get Book Here
Book Description
Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 32
Get Book Here
Book Description
Author: Jerome D. Julius
Publisher:
ISBN:
Category :
Languages : en
Pages :
Get Book Here
Book Description
Author: Jerome D. Julius
Publisher:
ISBN:
Category :
Languages : en
Pages : 28
Get Book Here
Book Description
Author: Raul Jorge Conti
Publisher:
ISBN:
Category : Heat
Languages : en
Pages : 38
Get Book Here
Book Description
Two circulation conical configurations having 15° half-angles were tested in laminar boundary layer at a Mach number of 6 and angles of attack up to 90°. One cone had a sharp nose and a fineness ratio of 1.87 and the other had a spherically blunted nose with a bluntness ratio of 0.1428 and a fineness ratio of 1.66. Pressure measurements and schlieren pictures of the flow showed that near-conical flow existed above 70° high pressure areas were present near the base and the bow shock wave was considerably curved.
Author: Stallings, Jr. (Robert L.)
Publisher:
ISBN:
Category : Aerodynamic measurements
Languages : en
Pages : 21
Get Book Here
Book Description
Author: John O. Reller
Publisher:
ISBN:
Category : Aerodynamics
Languages : en
Pages : 64
Get Book Here
Book Description
Abstract: Temperature recovery factors were determined for a slender, thin-walled cone-cylinder, having a 12° vertex angle and a 1.25-inch-diameter cylinder, at Mach numbers from 3.02 to 6.30. The angle-of-attack range was 0° to 45° at Mach numbers up to 3.50, and about 0° to 20° at Mach numbers from 4.23 to 6.30. A transverse cylinder of the same diameter was also tested at Mach number 3.02. Free-stream Reynolds numbers varied from 1.8 to 11.0 million per foot. Flow visualization studies of boundary-layer transition and flow separation were made and the results correlated with recovery-factor measurements.
Author: Dennis M. Bushnell
Publisher:
ISBN:
Category : Aerodynamics
Languages : en
Pages : 40
Get Book Here
Book Description
Author: Charles B. Rumsey
Publisher:
ISBN:
Category : Aerodynamic heating
Languages : en
Pages : 26
Get Book Here
Book Description
Aerodynamic-heat-transfer measurements have been made at a station on the 10 degree total angle conical nose of a rocket-propelled model at flight Mach numbers of 1.4 to 3.9. The corresponding values of local Reynolds number varied from 18,000,000 to 46,000,000 and the ratio of skin temperature to local static temperature varied from 1.2 to 2.4. The experimental data, reduced to Stanton number, were in fair agreement with values predicted by Van Driest's theory for heat transfer on a cone with turbulent flow from the nose tip.
Author: Charles B. Rumsey
Publisher:
ISBN:
Category : Aerodynamics
Languages : en
Pages : 42
Get Book Here
Book Description
Abstract: Aerodynamic heat-transfer measurements were at six stations on the 40-inch-long 10° total-angle conical nose of a rocket-propelled model which was flight tested at Mach numbers up to 5.9. The range of local Reynolds number was from 6.6 x 106 to 55.2 x 106. Laminar, transitional, and turbulent heat-transfer coefficients were measured, and, in general, the laminar and turbulent measurements were in good agreement with theory for cones. Experimental transition Reynolds numbers varied from less than 8.5 x 106 to 19.4 x 106. At a relatively constant ratio of wall temperature to local static temperature near 1.2, the transition Reynolds number increased from 9.2 x 106 to 19.4 x 106 as Mach number increased from 1.57 to 3.38. At Mach numbers near 3.7, the transition Reynolds number decreased as the skin temperature increased toward adiabatic wall temperatures.
