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Johns, A. L.

Publications and source records attributed to Johns, A. L..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

STOVL Hot Gas Ingestion control technology

A comprehensive wind tunnel test program was conducted to evaluate control of Hot Gas Ingestion (HGI) on a 9.2 percent scale model of the McDonnell Aircraft Company model 279-3C advanced Short Takeoff and Vertical Landing (STOVL) configuration. The test was conducted in the NASA-Lewis Research Center 9 ft by 15 ft Low Speed Wind Tunnel during the summer of 1987. Initial tests defined baseline HGI levels as determined by engine face temperature rise and temperature distortion. Subsequent testing was conducted to evaluate HGI control parametrically using Lift Improvement Devices (LIDs), forward nozzle splay angle, a combination of LIDs and forward nozzle splay angle, and main inlet blocking. The results from this test program demonstrate that HGI can be effectively controlled and that HGI is not a barrier to STOVL aircraft development.

Amuedo, K. C.↗
NASA NTRS1984

Flow visualization and interpretation of visualization data for deflected thrust V/STOL nozzles

Flow visualization studies were made for four deflected thrust nozzle models at subsonic speeds. Based on topological rules and the assumption that observed streaks constitute continuous vector fields, available visualization pictures are interpreted and flow patterns on interior surfaces of the nozzles are synthesized. In particular, three dimensional flow structure and separations are discussed. From the synthesized patterns, the overall features of the flow field in a given nozzle can be approximately perceived.

Kao, H. C.↗
NASA NTRS1984

Flow visualization and interpretation of visualization data for deflected thrust V/STOL nozzles

Flow visualization studies were made for four deflected thrust nozzle models at subsonic speeds. Based on topological rules and the assumption that observed streaks constitute continuous vector fields, available visualization pictures are interpreted and flow patterns on interior surfaces of the nozzles are synthesized. In particular, three dimensional flow structure and separations are discussed. From the synthesized patterns, the overall features of the flow field in a given nozzle can be approximately perceived. Previously announced in STAR as N84-14147

Kao, H. C.↗
NASA NTRS1983

Experimental results of a deflected thrust V/STOL nozzle research program

Four deflected thrust nozzle concepts, designed to operate at the low pressure ratio typical of high bypass-ratio turbofan engines for medium speed (subsonic) V/STOL aircraft, were studied. Maps of overall performance characteristics and exit velocity distributions are used to highlight similarities and differences between the four concepts. Analytically determined secondary flows at the exit of a 90 deg circular pipe bend are compared with the experimental results from the more complex three dimensional geometries. The relative impact of total-pressure losses and secondary flows on nozzle thrust coefficient is addressed by numerical integration of exit velocity measurements.

Burstadt, P. L.↗
NASA NTRS1983

Experimental results of a deflected thrust V/STOL nozzle research program

Four deflected thrust nozzle concepts, designed to operate at the low pressure ratio typical of high bypass-ratio turbofan engines for medium speed (subsonic) V/STOL aircraft, were studied. Maps of overall performance characteristics and exit velocity distributions are used to highlight similarities and differences between the four concepts. Analytically determined secondary flows at the exit of a 90 deg circular pipe bend are compared with the experimental results from the more complex three dimensional geometries. The relative impact of total-pressure losses and secondary flows on nozzle thrust coefficient is addressed by numerical integration of exit velocity measurements. Previously announced in STAR as N83-25657

Burstadt, P. L.↗
NASA NTRS1980

Wind tunnel investigation of the Titan Forward Skirt compartment vent from a free-stream Mach number of 0.80 to 1.96

A test was conducted to determine the flow characteristics of the Titan forward skirt compartment vent over a free stream Mach number range of 0.80 to 1.96. The vent was mounted in a flat plate and the plate was flush mounted to the tunnel side wall with coinciding center lines. Air was discharged from a duct, located on the tunnel side wall behind the plate, through a canted aft 30 deg honeycomb vent into the free stream. Data for the analysis of the Titan forward skirt compartment venting during ascent through the atmosphere are provided. Full scale simulated flight hardware, such as the honeycomb vent, duct corrugations and field joint ring were used. Boundary layer thicknesses were used to vary boundary height. The highest vent discharge coefficient for any given Mach number and vent pressure ratio generally occurred at the maximum displacement thickness. With no vent flow the static pressure in the vent region was generally less than the free stream static pressure. With vent flow, the static pressures upstream of the vent increased, and those downstream of the vent decreased.

Johns, A. L.↗
NASA NTRS1979

Performance of a V/STOL tilt nacelle inlet with blowing boundary layer control

A scale model of a V/STOL tilt nacelle fitted to a 0.508 m single stage fan was tested in the NASA Lewis low speed wind tunnel to determine the effect of diffuser blowing on the inlet aerodynamics and aeromechanical performance. The test was conducted over a range of freestream speeds (up to 120 knots) and angles-of attack (up to 120 deg). In general, diffuser blowing had a beneficial affect on all performance parameters. The angle-of-attack range for a separation-free flow substantially increased, and the fan face distortion reduced with a corresponding increase in total pressure recovery. Discrete narrow band blade stress peaks which were common to the nonblowing (baseline) configuration were eradicated with diffuser blowing.

Johns, A. L.↗
NASA NTRS1979

Performance of a V/STOL tilt nacelle inlet with blowing boundary layer control

A scale model of a V/STOL tilt nacelle fitted to a 0.508 m single stage fan was tested in the NASA Lewis 9x15 ft low speed wind tunnel to determine the effect of diffuser blowing on the inlet aerodynamics and aeromechanical performance. The test was conducted over a range of freestream speeds (up to 120 knots) and angles of attack (up to 120 deg). Diffuser blowing had a beneficial affect on all performance parameters. The angle of attack range for separation free flow substantially increased, and the fan face distortion significantly reduced with a corresponding increase in total pressure recovery. Discrete narrow band blade stress peaks which were common to the nonblowing (baseline) configuration were eradicated with diffuser blowing.

Johns, A. L.↗
NASA NTRS1975

Wind tunnel investigation of a Centaur standard shroud compartment vent from Mach number of 0.70 to 1.96

An experimental investigation was conducted in the Lewis Research Center 8- by 6-foot supersonic wind tunnel to determine the vent discharge coefficient for the Centaur standard shroud/liquid hydrogen tank compartment vent. The test was conducted from Mach 0.70 to 1.96 with the vent mounted in a flat plate. Full scale simulated flight hardware, such as the vent, corrugations, aft field joint ring and ice bag clip was used. Air was discharged from a plenum chamber, located on the tunnel sidewall behind the plate, through five 6.35 cm diameter vent orifices into the free stream. Boundary layer thickeners, analytically predicted displacement thickness for the vehicle nominal flight trajectory could be simulated over the Mach number range. The highest vent discharge coefficient for any given Mach number and vent pressure ratio generally occurred at the maximum displacement thickness.

Johns, A. L.↗
NASA NTRS1975

Flight investigation of installation effects on a wedge nozzle installed on an underwing nacelle

A flight research program was conducted using a modified F-106B aircraft with an underwing-nacelle - engine installation to investigate installation effects on a wedge nozzle with retracted shrouds from Mach 0.70 to 1.10. These configurations simulated the subsonic flight geometries of a variable-geometry wedge nozzle design for Mach 2.70 operation. The nozzle was tested with a J85-GE-13 turbojet engine, and data were compared with that of 0.34-scale isolated wind tunnel model. Data are also compared with the flight results of a boattail plug nozzle and a cylindrical nacelle plug nozzle. A favorable installation effect occurred from Mach 0.82 to 0.96 with a nominal 4-percent corrected secondary weight-flow ratio. This favorable effect resulted from changes in pressure forces on the nozzle surfaces. The wedge nozzle gross thrust coefficient was about the same as the boattail nacelle plug nozzle and 2.50 to 1.30 percent higher than the cylindrical nacelle plug nozzle over the subsonic Mach number range.

Johns, A. L.↗
NASA NTRS1974

Venting characteristics of gaseous helium and nitrogen discharging into a free stream at Mach numbers from 0.60 to 1.57

An experimental investigation was conducted in the Lewis Research Center 8- by 6-Foot Supersonic Wind Tunnel to compare the effects on discharge coefficient of venting gaseous helium and gaseous nitrogen into a free stream. The test was conducted from Mach 0.06 to 1.57 with the vent mounted in a flat plate. The plate was strut mounted to the tunnel ceiling and at a 0 angle of attack. The gases were discharged from a plenum chamber through a 2.54-centimeter (1.00-in.) diameter vent. The ratio of local static pressure to plenum pressure was varied from 0.51 to 0.975. The ratio of boundary layer thickness to vent diameter varied from a maximum of 1.34 at Mach 0.60 to a minimum of 0.55 at Mach 1.37.

Johns, A. L.↗
NASA NTRS1974

Effects of flanges on pressure distribution on a flat plate and on a corrugated surface at Mach numbers from 0.60 to 1.97

An 8 by 6 foot supersonic wind tunnel was used to obtain the static pressure distribution on a plate in the region of a flange placed normal to the airstream. Tests were conducted on both a flat plate surface and a corrugated surface using flange heights ranging from 10 to 125 percent of the boundary layer height. Data were obtained at a zero degree angle-of-attack and at Mach numbers from 0.60 to 1.97.

Johns, A. L.↗
NASA NTRS1974

Flight velocity effects on jet noise of several variations of a twelve-chute suppressor installed on a plug nozzle

Because of the relatively high takeoff speeds of supersonic transport aircraft, it is important to know whether the flight velocity effects the noise level of suppressor nozzles. To investigate this, a modified F-106B aircraft was used to conduct a series of flyover and static tests on a 12-chute suppressor installed on an uncooled plug nozzle. Comparison of flyover and static spectra indicated that flight velocity adversely affected noise suppressions of the 12-chute configurations.

Burley, R. R.↗
NASA NTRS1973

Internal performance of a wedge nozzle for a supersonic cruise aircraft with a multispoke primary for noise suppression

Nozzle performance was obtained with cold primary and secondary flows over a range of nozzle pressure ratio from 2 to 31 and corrected secondary-to primary-flow ratio from 0 to 11 percent. Several fixed shroud positions were tested to simulate a translating outer cylindrical shroud to provide data for both supersonic cruise and takeoff configurations. The addition of 14 spokes in the primary nozzle for noise suppression reduced the thrust efficiency 3 percent at supersonic cruise and 4.5 percent at takeoff.

Johns, A. L.↗