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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1998

Impact of Fuselage Cross Section on the Stability of a Generic Fighter

Many traditional data bases, which involved smooth-sided forebodies, are no longer relevant for designing advanced aircraft. The current work provides data on the impact of chined-shaped fuselage cross section on the stability of a generic fighter configuration. Two different chined-shaped fuselages were tested upright and inverted. It was found that a fuselage with a 30" included chine angle resulted in significantly higher values of fuselage with a 100" included chine angle. This difference was attributed to the more beneficial vortical interaction between the stronger forebody vortices coming off of the sharper chine edges and the wing vortices. The longitudinal stability of the configuration with the sharper chine angle was also better because, based on pressures and flow visualization, the vortex burst over the wing was delayed until significantly higher values of a. Unstable rolling moment derivatives were also delayed to higher values of a for the sharper chine angle cross section. Furthermore, it was found that directional stability of both of the upright configurations, which had larger lofts in cross section above the chine lines than below the chine lines. was better than for the inverted configurations.

Hall, Robert M.↗
NASA NTRS1990

Transonic wind tunnel test of a 14 percent thick oblique wing

An experimental investigation was conducted at the ARC 11- by 11-Foot Transonic Wind Tunnel as part of the Oblique Wing Research Aircraft Program to study the aerodynamic performance and stability characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing designed by Rockwell International. The 10.3 aspect ratio, straight-tapered wing of 0.14 thickness/chord ratio was tested at two different mounting heights above the fuselage. Additional tests were conducted to assess low-speed behavior with and without flaps, aileron effectiveness at representative flight conditions, and transonic drag divergence with 0 degree wing sweep. Longitudinal stability data were obtained at sweep angles of 0, 30, 45, 60, and 65 degrees, at Mach numbers ranging from 0.25 to 1.40. Test Reynolds numbers varied from 3.2 to 6.6 x 10 exp 6/ft. and angle of attack ranged from -5 to +18 degrees. Most data were taken at zero sideslip, but a few runs were at sideslip angles of +/- 5 degrees. The raised wing position proved detrimental overall, although side force and yawing moment were reduced at some conditions. Maximum lift coefficient with the flaps deflected was found to fall short of the value predicted in the preliminary design document. The performance and trim characteristics of the present wing are generally inferior to those obtained for a previously tested wing designed at ARC.

Kennelly, Robert A., Jr.↗
NASA NTRS1999

Experimental Aerodynamic Characteristics of an Oblique Wing for the F-8 OWRA

An experimental investigation was conducted during June-July 1987 in the NASA Ames 11-Foot Transonic Wind Tunnel to study the aerodynamic performance and stability and control characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing. This effort was part of the Oblique Wing Research Aircraft (OWRA) program performed in conjunction with Rockwell International. The Ames-designed, aspect ratio 10.47, tapered wing used specially designed supercritical airfoils with 0.14 thickness/chord ratio at the root and 0.12 at the 85% span location. The wing was tested at two different mounting heights above the fuselage. Performance and longitudinal stability data were obtained at sweep angles of 0deg, 30deg, 45deg, 60deg, and 65deg at Mach numbers ranging from 0.30 to 1.40. Reynolds number varied from 3.1 x 10(exp 6)to 5.2 x 10(exp 6), based on the reference chord length. Angle of attack was varied from -5deg to 18deg. The performance of this wing is compared with that of another oblique wing, designed by Rockwell International, which was tested as part of the same development program. Lateral-directional stability data were obtained for a limited combination of sweep angles and Mach numbers. Sideslip angle was varied from -5deg to +5deg. Landing flap performance was studied, as were the effects of cruise flap deflections to achieve roll trim and tailor wing camber for various flight conditions. Roll-control authority of the flaps and ailerons was measured. A novel, deflected wing tip was evaluated for roll-control authority at high sweep angles.

Kennelly, Robert A., Jr.↗
NASA NTRS1947

Data Obtained in the Flight Measurements to Determine the Stability and Control Characteristics of a C-54D Airplane (AAF No. 42-72713) and a Summary of the Test Program

The flight investigation of the C-54D airplane was initiated to determine the necessity of changes or additions to existing handling-qualities requirements to cove the case of instrument approaches with large airplanes. This paper gives a brief synopsis of the results and presents the measured data of tests to determine the stability and control characteristics. It was found that no new requirements were necessary to cover the problems of instrument approaches. The C-54D airplane tested met the Amy and Navy stability and control requirements except for the following items. The control-system friction with autopilot installed vas double that allowed by the requirements. The amount of friction was found to impair the controllability of the airplane in precision flying. The lateral and directional characteristics were good except that the maximum pb/2V was slightly below the minimum required, and the elevator-control forces to obtain the maximum pb/2V at low speeds were above the Army and Navy requirements. The longitudinal stability and control characteristics were good except that the elevator-control forces exceeded the limits of the Army and Navy requirements in turns and in landings. The stalling characteristics were considered good in all conditions with the stall warning in the form of tail buffeting occurring at speeds approximately 5 miles per hour above the stall.

Talmage, Donald B.↗
NASA NTRS1947

Force Tests of a 1/5-Scale Model of the McDonnell XP-85 Airplane with Conventional Tail Assembly in the Langley Free-Flight Tunnel

At the request of the Air Materiel Command, Army Air Forces an investigation of the low-speed, power-off stability and control characteristics of the McDonnell XP-85 airplane is being conducted in the Langley free-flight tunnel. The XP-85 airplane is a parasite fighter carried in a bomb bay of the B-36 airplane. As a part of the investigation a few force tests were made of a 1/5 scale model of the XP-85 with a conventional tail assembly installed in place of the original design five-unit tail assembly. The total area of the conventional assembly was approximately 80 percent of the area of the five-unit assembly. The results of this investigation showed that the conventional tail assembly gave about the same longitudinal stability characteristics as the original configuration and improved the directional and lateral stability.

Paulson, John W.↗
NASA NTRS1945

Wind-Tunnel Tests of a 1/5-Scale Semispan Model of the Republic XF-12 Horizontal Tail Surface

Wind-tunnel tests of a 1/5-scale semispan model of the Republic XF-12 horizontal tail surface equipped with an internally balanced elevator were conducted in the 6- by 6-foot test section of the Langley stability tunnel. The tests included measurements of the aerodynamic characteristics of the horizontal tail with and without a beveled trailing edge and also included measurements of the tab characteristics. The variation of the aerodynamic characteristics with boundary-layer conditions and leakage in the internal-balance chambers, measurements of the boundary-layer displacement thickness near the elevator hinge axis, and pressure distributions at the mean geometric chord were also obtained. The results showed that the hinge-moment characteristics of the elevator were critical to boundary-layer conditions and internal-balance leakage. Increasing the boundary-layer displacement thickness by use of roughness strips reduced the rate of change of elevator hinge moments with tab deflection by about 20 percent. The present horizontal tail appears to be unsatisfactory for longitudinal stability with power on, however, an increase in horizontal-tail lift effectiveness should correct this difficulty. The maneuvering stick force per unit acceleration will be extremely critical to minor variations of the elevator hinge moments if the elevator is linked directly to the stick.

Denaci, H. G.↗
NASA NTRS1970

Effects of Nose Bluntness on Aerodynamic Characteristics of Cruciform-Finned Missile Configuration at Mach 1.50 to 2.86

An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the effects of spherical nose bluntness on the aerodynamic characteristics of a typical cruciform-finned short-range missile configuration for roll angles of 0 deg and 45 deg and Mach numbers from 1.50 to 2.86. Also investigated was the effectiveness of a forward protruding nose spike in reducing the axial-force coefficient of a blunt-nose configuration. The results indicated that the effects of nose bluntness and nose spike on the longitudinal stability were small. For roll angles of both 0 deg and 45 deg, the greatest effects of nose bluntness on the aerodynamic-center location near zero angle of attack occurred at a Mach number of 2.86, where for each roll position, a rearward shift of approximately 3.5 percent of the body length occurred as bluntness was increased. Although large magnitudes of nose bluntness were accompanied by high axial-force coefficients, significant reductions in the axial-force coefficient of a blunt-nose configuration were achieved with the use of the nose spike. There were no noteworthy effects of nose bluntness on either the directional stability or side-force parameters. However, for high angles of attack at a Mach number of 1.50 increased nose bluntness generally resulted in decreased effective dihedral.

Jernell, L. S.↗
NASA NTRS1973

Full-scale wind-tunnel investigation of effects of slot spoilers on the aerodynamic characteristics of a light twin-engine airplane

A wind-tunnel investigation has been conducted to determine the effects of slot spoilers on the longitudinal and lateral aerodynamic characteristics of a full-scale mockup of a light twin-engine airplane. The slots were located along the leading edge of the flaps and were used to modulate the flap-induced lift as a possible means of achieving direct lift control. The data showed that the slots were effective in spoiling up to 61 percent of the flap-induced lift, but that an adverse pitching-moment change (nose up) accompanied opening the slots. Opening the slots was found to decrease slightly the downwash angle at the tail and to increase slightly the longitudinal stability of the model.

Verstynen, H. A., Jr.↗
NASA NTRS1974

Results of investigations (OA20C) on an 0.015-scale configuration 140A/B space shuttle vehicle orbiter model (49-0) in the NASA/Langley Research Center Unitary Plan Wind Tunnel

Data obtained from the wind tunnel tests of a scale model of the space shuttle orbiter configuration 140 A/B are presented. The test was conducted at Mach numbers of 2.5, 3.9, and 4.6 with Reynolds numbers from 1.25 million per foot to 5.0 million per foot. Various control surface settings were used ranging from an angle of attack range from minus 4 to plus 42 degrees at zero angle of yaw. Longitudinal stability and control characteristics of the space shuttle configuration were analyzed.

Nichols, M. E.↗
NASA NTRS1974

Wind tunnel tests of an 0.015-scale configuration 140A/B space shuttle orbiter model (67-0) in the NASA/LRC 8-foot TPT to obtain transonic aerodynamic force data (OA106)

These tests were conducted to obtain longitudinal stability and control data for Mach numbers from 0.35 to 1.2. Data were obtained for an alpha range of -2 deg to +22 deg at beta = 0 deg. The effect of speedbrake deflection and body flap deflection was obtained through the Mach range. Boundary layer transition strips were used on the model.

Burrows, R. R.↗
NASA NTRS1976

Effect of hinge-moment parameters on elevator stick forces in rapid maneuvers

The importance of the stick force per unit normal acceleration as a criterion of longitudinal stability and the critical dependence of this gradient on elevator hinge moment parameters are investigated with special reference to transient effects for maneuvers of short duration. The analysis shows that different combinations of elevator parameters, which give the same stick force per unit acceleration in turns, give widely different force variations during the entries into and recoveries from steady turns and during maneuvers of short duration such as abrupt pull-ups. The stick force per unit acceleration is greater for abrupt than for gradual control movements.

Jones, R. T.↗
NASA NTRS1975

Low-speed wind-tunnel tests of a one-tenth-scale model of a blended-arrow advanced supersonic transport

Tests were conducted in the Langley full scale tunnel to determine the low-speed aerodynamic characteristics of a 1/10 scale model of a blended-arrow advanced supersonic transport. Tests were made for the clean configuration and a high-lift configuration with several combinations of leading- and trailing-edge flaps deflected for providing improved lift and longitudinal stability in the landing and takeoff modes. The tests were conducted for a range of angles of attack from about -6 deg to 30 deg, sideslip angles from -5 deg to 10 deg, and for Reynolds numbers from 6.78 x 1,000,000 to 13.85 x 1,000,000 corresponding to test velocities of 41 knots to 85 knots, respectively.

Lemore, H. C.↗
NASA NTRS1977

Low-speed wind tunnel investigation of an advanced supersonic cruise arrow-wing configuration

A preliminary assessment of possible means for improving the low speed aerodynamic characteristics of advanced supersonic cruise arrow wing configurations and to extend the existing data base of such configurations has been made. Principle configuration variables included wing-leading and trailing-edge flap deflection, fuselage nose strakes, and engine exhaust nozzle deflection. Results showed that deflecting the wing leading edge apex flaps downward provided improved longitudinal stability but resulted in reduced directional stability. The model exhibited relatively low values of directional stability over the operational angle of attack range and experienced large asymmetric yawing moments at high angles of attack. The use of nose strakes was found to be effective in increasing the directional stability and eliminating the asymmetric yawing moment.

Coe, P. L., Jr.↗
NASA NTRS1977

Effects of deflected thrust on the longitudinal aerodynamic characteristics of a close-coupled wing-canard configuration

The effects of power on the longitudinal aerodynamic characteristics of a close-coupled wing-canard fighter configuration with partial-span rectangular nozzles at the trailing edge of the wing were investigated. Data were obtained on a basic wing-strake configuration for nozzle and flap deflections from 0 deg to 30 deg and for nominal thrust coefficients from 0 to 0.30. The model was tested over an angle-of-attack range from -2 deg to 40 deg at Mach numbers of 0.15 and 0.18. Results show substantial improvements in lift-curve slope, in maximum lift, and in drag-due-to-lift efficiency when the canard and strakes have been added to the basic wing-fuselage (wing-alone) configuration. Addition of power increased both lift-curve slope and maximum lift, improved longitudinal stability, and reduced drag due to lift on both the wing-canard and wing-canard-strake configurations. These beneficial effects are primarily derived from boundary-layer control due to moderate thrust coefficients which delay flow separation on the nozzle and inboard portion of the wing flaps.

Yip, L. P.↗
NASA NTRS1978

Flight tests of a radio-controlled airplane mode with a free-wing, free-canard configuration

Flight characteristics, controllability, and potential operating problems were investigated in a radio-controlled airplane model in which the wing is so attached to the fuselage that it is free to pivot about a spanwise axis forward of its aerodynamic center and is subject only to aerodynamic pitching moments imposed by lift and drag forces and a control surface. A simple technique of flying the test vehicle in formation with a pickup truck was used to obtain trim data. The test vehicle was flown through a series of maneuvers designed to permit evaluation of certain characteristics by observation. The free-wing free-canard concept was determined to be workable. Stall/spin characteristics were considered to be excellent, and no effect on longitudinal stability was observed when center of gravity changes were made. Several problems were encountered during the early stages of flight testing, such as aerodynamic lockup of the free canard and excessive control sensitivity. Lack of onboard instrumentation precluded any conclusions about gust alleviation or ride qualities.

Gee, S. W.↗
NASA NTRS1978

Wind-tunnel investigation at Mach numbers from 1.90 to 2.86 of a canard-controlled missile with ram-air-jet spoiler roll control

The efficacy of using a ram-air-jet spoiler roll control device on a typical canard-controlled missile configuration was investigated. For roll control comparisons, conventional aileron controls on the tail fins were also tested. The results indicate that the roll control of the ram-air-jet spoiler tail fins at the highest free-stream Mach number compared favorably with that of the conventional 11-70 area-ratio tail fin ailerons, each deflected 10 deg. The roll control of the tail fin ailerons decreased while that of the ram-air-jet spoiler increased with free-stream Mach number. The addition of the ram-air-jet spoiler tail fins or flow-through tip chord nacelles on the tail fins resulted in only small changes in basic missile longitudinal stability. The axial force coefficient of the operating ram-air-jet spoiler is significantly larger than that of conventional ailerons and results primarily from the total pressure behind a normal shock in front of the nacelle inlets.

Blair, A. B., Jr.↗
NASA NTRS1976

Summary of low-speed aerodynamic characteristics of upper-surface-blown jet-flap configurations

The results of recent wind tunnel investigations to provide fundamental information on the upper surface blown (USB) jet flap concept demonstrated that the USB concept provides good high-lift performance. It is shown that the low speed performance is dependent upon the jet turning angle and turning efficiency and on the use of proper leading and trailing edge treatment to prevent premature flow separation. The best means of achieving good turning performance in any particular USB application must be determined from overall operational considerations in which high speed performance, structures and noise, as well as low speed performance, are evaluated. The large diving moments generated at high lift coefficients can be trimmed satisfactorily with a large, conventional horizontal tail; a high tail position is best from longitudinal stability considerations. Large rolling and yawing moments are introduced with the loss of an engine, but these moments can be trimmed satisfactorily through the use of asymmetrical boundary layer control and through the use of spoiler and rudder deflection as needed.

Phelps, A. E., III↗