Systematic wind-tunnel investigations of vortex breakdown have been conducted on sharpedged delta and modified delta wings with sweep angles from 45° to 85° at Reynolds numbers of about 1 X 10 6 , utilizing a schlieren system for flow visualization. Vortex breakdown positions are presented as a function of angle of attack and sweep. Lift measurements generally compare favorably with Polhamus' leading-edge suction analogy. Drag because of the lift is predicted quite satisfactorily as the streamwise component of normal force for all models tested. Effects of vortex breakdown on pitching moments are presented and discussed. Tests of the modified delta wings show that vortex breakdown is influenced much more strongly by planform changes near the apex than changes near the trailing edge. Nomenclature BD-vortex core breakdown c = wing local chord C = wing mean geometric chord, C 0 = wing reference chord (usually centerline)wing pitching moment coefficient about I mean aerodynamic chord, moment/#>SC K p , K v = constants in Polhamus' lift equation q = dynamic pressure S -wing planform area TE = trailing edge a = angle of attack

A single-engine light airplane was modified by the installation of a wing with reduced area, Fowler flaps, Kruger flaps, and spoilers. Flight test results show that zero-lift drag was reduced 13.8% and a trimmed maximum lift coefficient of 2.73 was achieved. Gust response was significantly reduced and excellent roll control w a s achieved with spoilers. Several design features employed i n t h e n e w wings have excellent potential f o r i ncorporation i n future light airplanes.

No abstract

The fixed-base flight simulator at the University of Kansas Flight Research Laboratory was used to evaluate wing spoilers for longitudinal flight path control on a modified Cessna 177B aircraft. More than 100 simulated ILS approaches were flown by evaluation pilots using both conventional methods and spoiler controls. Three different spoiler control methods were evaluated. Spoilers provided precise glide path control with constant airspeed and attitude. Control is most effective when the steady-state trimmed airplane CL remains independent of spoiler position.Nomenclature ASEL = airplane, single engine, land CL = airplane lift coefficient C m = pitching moment coefficient Ah = altitude perturbation ILS = instrument landing system M § s = pitch acceleration due to spoiler deflection on one side n = load factor rms = root mean square AV = airspeed perturbation a -angle of attack Aa = angle of attack perturbation d s = symmetrical spoiler deflection A6 = pitch angle perturbation

A glycol-exuding porous leading edge ice protection system was tested in the NASAIcing Research Tunnel at Lewis Research Center. Test results showed that the system was very effective in preventing ice accretion (anti-ice mode) or removing ice from an airfoil. Minimumglycol flow rates required for antiicing are a function of velocity, liquid water content in the air, ambient temperature, and droplet size. Large ice caps were removedin only a few minutes using anti-ice flow rates, with the shed time being a function of the type of ice, size of the ice cap, angle of attack, and glycol flow rate. Wake survey measurementsshowedthat no significant drag penalty was associated with the installation or operation of the system tested.

Nomenclature B'= mass addition parameter h Q = stagnation enthalpy = /(To) h w = enthalpy of air at wall = f(T w ) LE = heat of sublimation of CO 2 LT = specific heat transfer to model interior M = molecular weight m = rate of mass addition per unit area p e = static pressure at edge of boundary layer p v = vapor pressure of CO 2 qb = heat-transfer rate from surface to interior of body qo = heat-transfer rate to surface without mass addition TQ,T W = stagnation and wall temperature, respectively X = distance aft of nose Subscripts 1 = ablating material 2 = f reestream fluid

No abstract

No abstract