Inlet performance characteristics of a generalized 1/5.2-scale aircraft model at transonic and supersonic Mach numbers /

A wind tunnel investigation was conducted at free-stream Mach numbers from 0.55 to 2.0 on a 1/5.2-scale composite inlet model to evaluate configuration factors which affected inlet performance. In addition, flow-field surveys were made at the inlet throat and on the fuselage, forward of the inlet cowl lip. Inlet performance parameters in terms of total-pressure recovery, distortion, and turbulence at the simulated compressor face, as well as compressor face and throat station total-pressure contours and diffuser duct static-pressure distributions are presented for various Mach numbers and angles of attack and sideslip. The basic air induction system consisted of a normal-shock-type inlet with a long splitter plate assembly and was located beneath the wing glove in proximity to the fuselage. The total-pressure recovery for the basic inlet system was slightly better than normal shock recovery for cruise attitudes at supersonic Mach numbers. At Mach numbers approaching 2.0, the basic inlet system was operating close to the buzz limit at design engine airflow. Reductions in total-pressure distortion were achieved by increasing the fuselage-to-inlet standoff distance and by inlet duct boundary-layer blowing. (Author).Photocopy."November 1976.""Final Report: November 24, 1975 -- March 26, 1976."Includes bibliographical references (page 17).A wind tunnel investigation was conducted at free-stream Mach numbers from 0.55 to 2.0 on a 1/5.2-scale composite inlet model to evaluate configuration factors which affected inlet performance. In addition, flow-field surveys were made at the inlet throat and on the fuselage, forward of the inlet cowl lip. Inlet performance parameters in terms of total-pressure recovery, distortion, and turbulence at the simulated compressor face, as well as compressor face and throat station total-pressure contours and diffuser duct static-pressure distributions are presented for various Mach numbers and angles of attack and sideslip. The basic air induction system consisted of a normal-shock-type inlet with a long splitter plate assembly and was located beneath the wing glove in proximity to the fuselage. The total-pressure recovery for the basic inlet system was slightly better than normal shock recovery for cruise attitudes at supersonic Mach numbers. At Mach numbers approaching 2.0, the basic inlet system was operating close to the buzz limit at design engine airflow. Reductions in total-pressure distortion were achieved by increasing the fuselage-to-inlet standoff distance and by inlet duct boundary-layer blowing. (Author).Report supported by the Air Force Systems Command and prepared by ARO, Inc, under Program Element,Mode of access: Internet

Static pressure on sharp and blunt cones in conical and parallel low-density flow /

The static pressure distribution on sharp and blunt 10-deg half-angle cones was measured under rarefied conditions in both a uniform and a source flow field. Angle of attack of the cones was varied up to 10 deg. Pressure distributions are presented for cold-wall conditions at 18.0 = or <Mach number = or <20.3 and 407 <Reynolds number/in. = or <1272. The source flow results are corrected according to the Newtonian pressure distribution, and the results agree with the uniform flow measurements for the longitudinal distribution at zero angle of attack. Comparisons are made with previous experimental and theoretical results.Report was prepared by ARO, Inc., a subsidiary of Sverdrup & Parcel and Associates, Inc."Final Report for Period October 1, 1971 -- March 31, 1972.""November 1974."Includes bibliograqphic references.The static pressure distribution on sharp and blunt 10-deg half-angle cones was measured under rarefied conditions in both a uniform and a source flow field. Angle of attack of the cones was varied up to 10 deg. Pressure distributions are presented for cold-wall conditions at 18.0 = or <Mach number = or <20.3 and 407 <Reynolds number/in. = or <1272. The source flow results are corrected according to the Newtonian pressure distribution, and the results agree with the uniform flow measurements for the longitudinal distribution at zero angle of attack. Comparisons are made with previous experimental and theoretical results.Mode of access: Internet

Unsteady viscous-inviscid interaction procedures for transonic airfoil flows /

"Final Report for Period October 1, 1982--September 30, 1983.""Program Element 65807F.""July 1984."Includes bibliographical references (pages 44-46).Mode of access: Internet

Supersonic wind tunnel tests of the IDS 6208 rocket sled /

"Program Element 74502154.""May 1965."Includes bibliographical references (page 4).Print reproduction.Air Force Contract No.Mode of access: Internet

Force and pressure tests of two elliptic ogive body shapes at Mach numbers 5, 6, and 8 /

Tests were conducted at Mach numbers 5, 6, and 8 to obtain force and pressure distribution data for two elliptic ogives. The results presented show the effect of Mach number and Reynolds number variation on aerodynamic forces for angles of attack from -6 to +6 deg and on pressure distributions for angles of attack from -4 to +4 deg. Shadowgraphs obtained at various angles of attack show shock wave and boundary layer phenomena."Research supported by the United States Air Force, and performed by the VKF, ARO, Inc.""ARO Project No. 321114.""Program Area 750A, Project No. 1370."AD0255762 (from http://www.dtic.mil)."April 1961."Includes bibliographical references (page 14).Tests were conducted at Mach numbers 5, 6, and 8 to obtain force and pressure distribution data for two elliptic ogives. The results presented show the effect of Mach number and Reynolds number variation on aerodynamic forces for angles of attack from -6 to +6 deg and on pressure distributions for angles of attack from -4 to +4 deg. Shadowgraphs obtained at various angles of attack show shock wave and boundary layer phenomena.Print reproduction.U.S. Air Force Contract No.Mode of access: Internet

Solution for the transient one-dimensional heat conduction in an infinite slab /

Report prepared by ARO, Inc., a subsidiary of Sverdrup and Parcel, Inc."March 1963."Includes bibliograqphic references (pages 16-18).Mode of access: Internet

The Use of boundary-layer suction for maintaining laminar flow downstream of a reflected incidence shock wave on a flat plate at free-stream Mach numbers 2.5, 3, and 3.5 /

"Program Element 62405334, Project No. 1366.""May 1965."Includes bibliographical references (page 5).Air Force Contract No.Mode of access: Internet

A method of characteristics computer program for three-dimensional supersonic internal flows /

The reference-plane method of characteristics is applied to the computation of the supersonic flow in a three-dimensional (3-D) channel such as a propulsion nozzle. The fluid is assumed to be an inviscid ideal gas and the flow is assumed to be shock-free, although it can be rotational. Both the basic equations and the numerical procedures are described, as is the computer program, which was written in FORTRAN IV language for the IBM 370/165 computer. The validity of the computer program was established by computing, in various ways, an axisymmetric nozzle flow as a three-dimensional flow; the numerical results are in good agreement with the results from a well-established computer program for axisymmetric flow. (Author)."January 1979.""Final Report: October 1977 -- August 1978."Includes bibliograpic references (page 22).The reference-plane method of characteristics is applied to the computation of the supersonic flow in a three-dimensional (3-D) channel such as a propulsion nozzle. The fluid is assumed to be an inviscid ideal gas and the flow is assumed to be shock-free, although it can be rotational. Both the basic equations and the numerical procedures are described, as is the computer program, which was written in FORTRAN IV language for the IBM 370/165 computer. The validity of the computer program was established by computing, in various ways, an axisymmetric nozzle flow as a three-dimensional flow; the numerical results are in good agreement with the results from a well-established computer program for axisymmetric flow. (Author).Report supported by the Air Force Systems Command and prepared by ARO, Inc,, a Sverdrup Corporation Company, under Program Element,Mode of access: Internet

Comparison of two methods used to measure aerodynamic loads acting on captive store models in wind tunnel tests /

In a series of wind tunnel tests, measurements were made of the aerodynamic loads acting on eight different store configurations supported in and near the external captive position on a 1/20-scale model of the F-4C aircraft. Store models included blunt and contoured afterbody shapes, stable and unstable configurations, and large (pylon-mounted) and small (rack-mounted) configurations. Six components of forces and moments were measured using two methods of supporting the store models: an internal bracket support technique (IBS), and a dual sting support technique. Some characteristics of the interference flow field of the aircraft were also inferred."September 1976.""Final Report: April 1973 -- May 1976."Includes bibliographical references (page 28).In a series of wind tunnel tests, measurements were made of the aerodynamic loads acting on eight different store configurations supported in and near the external captive position on a 1/20-scale model of the F-4C aircraft. Store models included blunt and contoured afterbody shapes, stable and unstable configurations, and large (pylon-mounted) and small (rack-mounted) configurations. Six components of forces and moments were measured using two methods of supporting the store models: an internal bracket support technique (IBS), and a dual sting support technique. Some characteristics of the interference flow field of the aircraft were also inferred.Report supported by the Air Force Systems Command and prepared by ARO, Inc, under Program Element,Mode of access: Internet

Wind tunnel investigation of aerodynamic loads on weapons separated from carriage under the wing of a tactical fighter aircraft at supersonic speeds /

Exploratory static-stability tests were made on models of the MK-84 weapon and a modular weapon configuration during simulated weapon separation from the wing of the F-4C Tactical Fighter Aircraft at freestream Mach numbers 1.76, 2.0, and 2.5. Data were obtained with the parent aircraft at zero angle of attack and the weapon at the carriage incidence angle of -1 deg. Supplementary data were obtained for weapon angles of attack of + or - deg. Unit Reynolds number ranged from 4.6 million per ft at M free stream mach number = 1.76 to 6.2 million per ft at M free stream mach number = 2.5. Selected results are presented which show the effects on the weapon aerodynamic loads of free-stream Mach number, weapon angle of attack, weapon axial position, and the presence of other stores. Representative shadowgraph and schlieren photographs of the interference flow field are also presented. (Author).Photocopy.Report was prepared by ARO, Inc., a subsidiary of Sverdrup & Parcel and Associates, Inc."Final Report for Period January 23 to February 20, 1973.""June 1973."Includes bibliograqphic references (page 10).Exploratory static-stability tests were made on models of the MK-84 weapon and a modular weapon configuration during simulated weapon separation from the wing of the F-4C Tactical Fighter Aircraft at freestream Mach numbers 1.76, 2.0, and 2.5. Data were obtained with the parent aircraft at zero angle of attack and the weapon at the carriage incidence angle of -1 deg. Supplementary data were obtained for weapon angles of attack of + or - deg. Unit Reynolds number ranged from 4.6 million per ft at M free stream mach number = 1.76 to 6.2 million per ft at M free stream mach number = 2.5. Selected results are presented which show the effects on the weapon aerodynamic loads of free-stream Mach number, weapon angle of attack, weapon axial position, and the presence of other stores. Representative shadowgraph and schlieren photographs of the interference flow field are also presented. (Author).Mode of access: Internet