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Morris, O. A.

Publications and source records attributed to Morris, O. A..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1985

Vapor-screen technique for flow visualization in the Langley Unitary Plan Wind Tunnel

The vapor-screen technique for flow visualization, as developed for the Langley Unitary Plan Wind Tunnel, is described with evaluations of light sources and photographic equipment. Test parameters including dew point, pressure, and temperature were varied to determine optimum conditions for obtaining high-quality vapor-screen photographs. The investigation was conducted in the supersonic speed range for Mach numbers from 1.47 to 4.63 at model angles of attack up to 35 deg. Vapor-screen photographs illustrating various flow patterns are presented for several missile and aircraft configurations. Examples of vapor-screen results that have contributed to the understanding of complex flow fields and provided a basis for the development of theoretical codes are presented with reference to other research.

Morris, O. A.↗
NASA NTRS1984

Investigation of light source and scattering medium related to vapor-screen flow visualization in a supersonic wind tunnel

Methods for increasing the radiant in light sheets used for vapor screen set-ups were investigated. Both high-pressure mercury arc lamps and lasers were considered. Pulsed operation of the air-cooled 1-kW lamps increased the light output but decreased reliability. An ellipsoidal mirror improved the output of the air-cooled lamps by concentrating the light but increased the complexity of the housing. Water-cooled-4-kW lamps coupled with high-aperture Fresnel lenses provided reasonable improvements over the air-cooled lamps. Fanned laser beams measurements of scattered light versus dew point made in conjunction with successful attempts to control the fluid injection. A number of smoke generators are described and test results comparing smoke and vapor screens are shown. Finally, one test included a periscope system to relay the image to a camera outside the flow.

Snow, W. L.↗
NASA NTRS1983

Low-speed investigation of the maneuver capability of supersonic fighter wings

A low-speed investigation of wings suitable for supersonic fighter application was undertaken to understand the flow regions developed by these wings through flow visualization and to evaluate areas in which low-speed testing could provide information for extrapolation to higher Mach numbers and Reynolds numbers. This study involved cranked wings of highly-swept inboard panels and outboard panel of various sweeps with leading- and trailing-edge devices. Both theoretical and experimental results with flow visualization are presented. Flow visualization results show that a mixed flow condition (vortex flow on the highly-swept inboard section and attached flow on the cranked outboard section) is possible for wings with low sweep (20 deg) outboard cranked sections. Lift and drag measurements on the planar (no flap deflection) wings show good agreement with theoretical results for leading-edge vortex flow. There is little or no change in planar wing performance at maneuver conditions due to outboard sweep effects. Results also show that for cambered wings (flap deflected) only small improvements in performance over flat wings were obtained at high-lift maneuver conditions.

Hom, K. W.↗
NASA NTRS1981

Aerodynamic tests and analysis of a turbojet-boosted launch vehicle concept (spacejet) over a Mach number range of 1.50 to 2.86

Results from analytical and experimental studies of the aerodynamic characteristics of a turbojet-boosted launch vehicle concept through a Mach number range of 1.50 to 2.86 are presented. The vehicle consists of a winged orbiter utilizing an area-ruled axisymmetric body and two winged turbojet boosters mounted underneath the orbiter wing. Drag characteristics near zero lift were of prime interest. Force measurements and flow visualization techniques were employed. Estimates from wave drag theory, supersonic lifting surface theory, and impact theory are compared with data and indicate the ability of these theories to adequately predict the aerodynamic characteristics of the vehicle. Despite the existence of multiple wings and bodies in close proximity to each other, no large scale effects of boundary layer separation on drag or lift could be discerned. Total drag levels were, however, sensitive to booster locations.

Riebe, G. D.↗
NASA NTRS1978

Aerodynamic characteristics of a fixed arrow-wing supersonic cruise aircraft at Mach numbers of 2.30, 2.70, and 2.95

Tests were conducted in the Langley Unitary Plan wind tunnel at Mach numbers of 2.30. 2.70, and 2.95 to determine the performance, static stability, and control characteristics of a model of a fixed-wing supersonic cruise aircraft with a design Mach Number of 2.70 (SCAT 15-F-9898). The configuration had a 74 deg swept warped wing with a reflexed trailing edge and four engine nacelles mounted below the reflexed portion of the wing. A number of variations in the basic configuration were investigated; they included the effect of wing leading edge radius, the effect of various model components, and the effect of model control deflections.

Morris, O. A.↗
NASA NTRS1977

Subsonic and supersonic aerodynamic characteristics of a supersonic cruise fighter model with a twisted and cambered wing with 74 deg sweep

A wind tunnel investigation has been conducted to determine the longitudinal and lateral aerodynamic characteristics of a model of a supersonic cruise fighter configuration with a design Mach number of 2.60. The configuration is characterized by a highly swept arrow wing twisted and cambered to minimize supersonic drag due to lift, twin wing mounted vertical tails, and an aft mounted integral underslung duel-engine pod. The investigation also included tests of the configuration with larger outboard vertical tails and with small nose strakes.

Morris, O. A.↗
NASA NTRS1976

Experimental effects of fuselage camber on longitudinal aerodynamic characteristics of a series of wing-fuselage configurations at a Mach number of 1.41

An experimental investigation was conducted to evaluate a method for the integration of a fighter-type fuselage with a theoretical wing to preserve desirable wing aerodynamic characteristics for efficient maneuvering. The investigation was conducted by using semispan wing fuselage models mounted on a splitter plate. The models were tested through an angle of attack range at a Mach number of 1.41. The wing had a leading edge sweep angle of 50 deg and an aspect ratio of 2.76; the wing camber surface was designed for minimum drag due to lift and was to be self trimming at a lift coefficient of 0.2 and at a Mach number of 1.40. A series of five fuselages of various camber was tested on the wing.

Dollyhigh, S. M.↗
NASA NTRS1974

Wind Tunnel Test of Low Boom Equivalent Body at Mach 4

A body of revolution, representing the equivalent area distribution of a low boom aircraft design cruising at 24,384 meters at a Mach number of 4, was tested to determine whether its theoretical sonic boom characteristics could be experimentally verified. Results indicate that the essential features of the ground signature are well predicted.

Howard, F. G.↗