Tests of a mixed compression axisymmetric inlet with large transonic mass flow at Mach numbers 0.6 to 2.65 / by Donald B. Smeltzer and Norman E. Sorensen.

Format:

Book

Government document

Author/Creator:

Smeltzer, Donald B., author.

Sorensen, Norman E., 1930- author.

Contributor:

United States. National Aeronautics and Space Administration, issuing body.

Ames Research Center

Series:

NASA technical note ; 6971.

NASA/TN ; D-6971

Language:

English

Subjects (All):

Jet engines--Air intakes.

Jet engines.

Aerodynamics, Transonic.

Physical Description:

1 online resource (v, 183 pages) : illustrations.

Place of Publication:

Washington, D.C. : National Aeronautics and Space Administration, December 1972.

Summary:

A 38.8-cm (15.28-in.) capture diameter model of a mixed-compression axisymmetric inlet system with a translating cowl was designed and tested. The internal contours, designed for Mach number 2.65, provided a throat area of 59 percent of the capture area when the cowl was retracted for transonic operation. Other model features included a boundary-layer removal system, vortex generators, an engine airflow bypass system, cowl support struts, and rotating rakes at the engine face. All tunnel testing was conducted at a tunnel total pressure of about 1 atm (a unit Reynolds number of about 8.53 million/m at Mach number 2.65) at angles of attack from 0 deg to 4 deg. Results for the following were obtained: total-pressure recovery and distortion at the engine face as a function of bleed mass-flow ratio, the effect of bleed and vortex generator configurations on pressure recovery and distortion, inlet tolerance to unstart due to changes in angle of attack or Mach number, surface pressure distributions, boundary-layer profiles, and transonic additive drag. At Mach number 2.65 and with the best bleed configurations, maximum total pressure recovery at the engine face ranged from 91 to 94.5 percent with bleed mass-flow ratios from 4 to 9 percent, respectively, and total-pressure distortion was less than 10 percent. At off-design supersonic Mach numbers above 1.70, maximum total-pressure recoveries and corresponding bleed mass flows were about the same as at Mach number 2.65, with about 10 to 15 percent distortion. In the transonic Mach number range, total pressure recovery was high (above 96 percent) and distortion was low (less than 15 percent) only when the inlet mass-flow ration was reduced 0.02 to 0.06 from the maximum theoretical value (0.590 at Mach number 1.0).

Notes:

"December 1972."

Includes bibliographical references (page 26).

Description based on online resource, PDF version; title from title page (NASA, viewed on June 23, 2021).

OCLC:

20643001