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Currently experimental aircraft icing research is severely hampered by limitations of ground icing simulation facilities. Existing icing facilities do not have the size, speed, altitude, and icing environment simulation capabilities to allow accurate studies to be made of icing problems occurring for high speed fixed wing aircraft and rotorcraft. Use of the currently dormant NASA Lewis Altitude Wind Tunnel (AWT), as a proposed high speed propulsion and adverse weather facility, would allow many such problems to be studied. The characteristics of the AWT related to adverse weather simulation and in particular to icing simulation are discussed, and potential icing research programs using the AWT are also included.

Currently experimental aircraft icing research is severly hampered by limitations of ground icing simulation facilities. Existing icing facilities do not have the size, speed, altitude, and icing environment simulation capabilities to allow accurate studies to be made of icing problems occurring for high speed fixed wing aircraft and rotorcraft. Use of the currently dormant NASA Lewis Altitude Wind Tunnel (AWT), as a proposed high speed propulsion and adverse weather facility, would allow many such problems to be studied. The characteristics of the AWT related to adverse weather simulation and in particular to icing simulation are discussed, and potential icing research programs using the AWT are also included.

Pneumatic deicer boots for helicopter rotor blades were tested. The tests were conducted in the 6 by 9 ft icing research tunnel on a stationary section of a UH-IH helicopter main rotor blade. The boots were effective in removing ice and in reducing aerodynamic drag due to ice.

A combined experimental and analytical program was conducted to investigate the effects of boundary layer changes on the flow over high angle boattail nozzles. The tests were run on an isolated axisymmetric sting mounted model. Various boattail geometries were investigated at high subsonic speeds over a range of boundary layer thicknesses. In general, boundary layer effects were small at speeds up to Mach 0.8. However, at higher speeds significant regions of separated flow were present on the boattail. When separation was present large reductions in boattail drag resulted with increasing boundary layer thickness. The analysis predicts both of these trends.

An approximately 1/30th-scale model of a CTOL aircraft designed for efficient cruise between Mach 0.9 and 0.98 was tested in the Lewis Research Center's 8- by 6-Foot Supersonic Wind Tunnel. Aft-fuselage axial force and pressure were measured for a series of locally area-ruled engine installations. Two flow-through nacelles with different exit diameters which simulated inlet mass flow ratios of 0.46 and 0.6 (based on maximum nacelle cross-sectional) area were investigated. Both an NACA-1 cowl and a blunter supercritical cowl were tested with these nacelles. Variations in the amount of local area ruling were investigated and an aft fuselage without nacelles was tested for reference. Results observed with flow-through nacelles indicate that efficient cruise may be obtained on this type of aircraft with aft-fuselage-mounted engine nacelles.

Three sting-mounted winged-body models with tunnel blockages of 0.1, 1.0, and 2.0 percent were tested in the Lewis Research Center's 8- by 6- Foot Supersonic Wind Tunnel. Fuselage pressures were obtained over a Mach number range of 0.6 to 1.0 at angles of attack from 0 deg to 4 deg. Two other types of model support were investigated, which included simulated wing-tip and fuselage support-strut mountings. The effects of tunnel porosity and sidewall geometry were also investigated. Model blockage effects were small up to M sub 0 = 0.95. At higher speeds the major blockage effect observed was a displacement of the local transonic terminal shocks on the model. The effects of the wing-tip type of model support were small up to M sub 0 = 0.95, but disturbances were observed on the fuselage at higher speeds. Changes in local tunnel porosity were effective in reducing the disturbances up to M sub 0 = 0.975, but a change in sidewall geometry was not.

Three eight bladed supersonic helical tip speed propellers were tested in the NASA Lewis Wind Tunnel at through flow Mach numbers of 0.6, 0.7, 0.75, 0.8, and 0.85. Such propellers are being considered as propulsors for a future energy conservative aircraft. The three propellers of equal performance, incorporate different plan forms and different amounts of sweep in order to investigate their noise generation, which could cause a cabin environment problem. The acoustically designed propeller SR-3 had 45 deg of tip sweep and was significantly quieter at Mach 0.8 cruise than the straight bladed propeller, SR-2. An intermediate 30 deg tip sweep propeller, SR-1m exhibited a noise level between the other two. Enhanced pressure-time traces indicated that SR-2 and SR-1m exhibited shock-like pressure traces at Mach-0.8 cruise while the SR-3 did not. Noise trends with varying helical tip Mach number and blade loading were also observed.

For the advanced turboprop to be competitive with proposed advanced turbofan-powered aircraft, it must have high propulsive efficiency at Mach 0.8 cruise above 9.144-km altitude with an acceptable cabin noise environment. Four 8-bladed propeller models are designed employing various concepts to reduce compressibility losses. Wind tunnel tests are conducted at zero model incidence to the free-stream flow. Aerodynamic and acoustic test results are presented and discussed. It is shown that the aeroacoustically designed configuration (SR-3) with 45 deg of tip sweep and an area-ruled spinner yields the highest propulsive efficiency (78.7% at Mach 0.8, 3.06 advance ratio, and 1.7 power coefficient), with an improvement of about 3% over the straight bladed configuration (SR-2, with zero-degree sweep). The phase-interference concept for noise reduction used in SR-3 yields about 5-6 dB reduction as compared to SR-2.

Three supersonic helical tip speed propellers were tested in the NASA Lewis 8- by 6-foot wind tunnel. Noise data were obtained while these propellers were operating at a simulated cruise condition. The walls of this tunnel were not acoustically treated and therefore this was not an ideal location for taking noise data, but it was thought that the differences in noise among the three propellers would be meaningful. The straight bladed propeller which did not incorporate sweep was the noisiest with the aerodynamically swept propeller only slightly quieter. However, the acoustically swept propeller was significantly quieter than the straight propeller, thereby indicating the merit of this design technique.

It is believed that a fuel saving of 14 to 40 percent may be realized by the use of an advanced high-speed turboprop. This turboprop must be capable of high efficiency at Mach 0.8 cruise above 9.144 km altitude if it is to compete with turbofan powered commercial aircraft. Several advanced aerodynamic concepts were investigated in recent wind tunnel tests under NASA sponsorship on two propeller models. These concepts included aerodynamically integrated propeller/nacelles, area-ruling, blade sweep, reduced blade thickness and power (disk) loadings several times higher than conventional designs. The aerodynamic design methodology for these models is discussed in this paper. In addition, some of the preliminary test results are presented which indicate that propeller net efficiencies near 80 percent were obtained for high disk loading propellers operating at Mach 0.8.

A flight and wind tunnel investigation was conducted to determine the effects of Reynolds number on the installed boattail drag of an underwing nacelle. Tests were run on a modified F-106B aircraft and 0.05 and 0.22 scale wind tunnel models. Tests were conducted at Mach numbers of 0.6 and 0.9 and over a 16 to 1 range of Reynolds numbers. Highest drag was obtained at intermediate Reynolds numbers corresponding to about the lowest flight values and that of the 0.22 scale model. Significantly lower drag was obtained at both higher and lower Reynolds numbers.

Auxiliary inlet ejector nozzle designed for supersonic cruise aircraft

Airframe installation effects at transonic speeds on underwing supersonic cruise exhaust nozzles, using flight and wind tunnel tests

Location effect of power plants on aft-mounted supersonic cruise exhaust nozzles at transonic speeds in flight and wind tunnel

Wind tunnel studies of flow field under 60 deg swept wing from Mach 0.6 to 2.6

Wind tunnel installation effects on isolated afterbodies at Mach numbers from 0.56 to 1.5

Wind tunnel tests of airframe installation effects on F-106B scale model with simulated underwing engine nacelles at Mach 0.56 to 1.46

Wind tunnel investigation of fluctuating pressures on Centaur model surface at transonic speed

Aerodynamic hinge moments for scale model boost stage nozzle of Saturn V launch vehicle during cold flow jet simulation

Validity of testing full scale flight components in transonic wind tunnel by evaluating flow deviations from free flight in critical transonic phase of flight

A flight and wind tunnel investigation was conducted to determine the effects of Reynolds number on the installed boattail drag of an underwing nacelle. Tests were run on a modified F-106B aircraft and 0.05 and 0.22 scale wind tunnel models. Tests were conducted at Mach numbers of 0.6 and 0.9 and over a 16 to 1 range of Reynolds numbers. Highest drag was obtained at intermediate Reynolds numbers corresponding to about the lowest flight values and that of the 0.22 scale model. Significantly lower drag was obtained at both higher and lower Reynolds numbers.