Engineering topics
Publications and source records attributed to Miller, B. A..
Program computes heating rates and surface friction effects. COLDARC predicts heating rate and surface friction on test article during plasmaarc testing. Uses simplified frozen-flow model to represent dissociated airflow and predict heat flux and surface friction, including effects of retarded atomic recombination from test facility data. COLDARC written in FORTRAN IV.
Correlations of orbiter windward surface entry heating data from the first five flights are presented with emphasis on boundary layer transition and the effects of catalytic recombination. Results show that a single roughness boundary layer transition correlation developed for spherical element trips works well for the orbiter tile system. Also, an engineering approach for predicting heating in nonequilibrium flow conditions shows good agreement with the flight test data in the time period of significant heating. The results of these correlations, when used to predict orbiter heating for a high cross mission, indicate that the thermal protection system on the windward surface will perform successfully in such a mission.
The effect of a combined external load consisting of a normal component and a tangential braking force applied to an inflated torus is addressed. Experimental results obtained by photographic study of the contact area between the torus and load plate are presented as well as measurements of the vertical and horizontal displacement of the torus under load. A numerical procedure for displacement analysis is developed in which the finite element program STAGS (Structural Analysis of General Shells) is used in an iterative manner to produce a flat, horizontal footprint surface under force loading. The redistributed force distribution obtained by the iterative process is displayed along with computed meridional and circumferential stresses. Finally, an extension of the iterative method is introduced which eliminates the need to experimentally determine the footprint area.
Wind tunnel model tests support the hypothesis that a propeller tip vortex may subject a downstream wing surface to greater excitation than would be experienced by the aircraft fuselage side wall exposed to propeller-generated noise, ultimately transmitting this structural response to incident dynamic pressure to the cabin interior. Even if structure-borne excitations are less efficient than airborne excitations in the creation of cabin noise, the higher level of the former could still govern cabin noise levels.
Although the assumption is generally made that cabin noise levels are governed by the transmission of propeller generated noise through the fuselage sidewall, it was postulated that the propeller wake striking the wing, in particular pressure disturbances generated downstream of the propeller by the action of the propeller tip vortex, could be strong enough to excite the aircraft structure and contribute to the cabin noise level. Tests conducted to measure the strength of the propeller tip vortex support this hypothesis. It was found that the propeller tip vortex can produce a fluctuation pressure on a simulated wing surface in the wake of a propeller that exceeds by more than 15 dB the maximum direct noise that would strike the fuselage. Wing surface response to propeller tip vortex induced excitations, and the effectiveness of this response in radiating noise to the cabin interior, must be established to assess the full significance of these results.
NASA's Advanced Turboprop Project is a three-phase effort initiated in 1978 to provide technology readiness for Mach 0.7 to 0.8 turboprop-powered aircraft; with the potential for fuel savings and DOC reductions of up to 30 and 15%, respectively, relative to current in-service aircraft. This paper reviews the status of Phase I in the areas of propeller aeroacoustics, propeller structures, turboprop installed performance, aircraft cabin environment, and turboprop engine and aircraft studies. Current plans to establish large-scale propeller characteristics and to conduct high-speed propeller flight research tests using a modified testbed aircraft are also presented.
Technology readiness for Mach 0.7 to 0.8 turboprop powered aircraft with the potential for fuel savings and DOC reductions of up to 30 and 15 percent respectively relative to current in-service aircraft is addressed. The areas of propeller aeroacoustics, propeller structures, turboprop installed performance, aircraft cabin environment, and turboprop engine and aircraft studies are emphasized. Large scale propeller characteristics and high speed propeller flight research tests using a modified testbed aircraft are also considered.
An inlet suitable for a turbine engine in a STOL, VTOL or CTOL aircraft is described. A circumferentially extended slot is provided in the inner surface of the air inlet at the windward side and downstream of the throat region. The slot communicates with a circumferential plenum chamber formed in the front of the air inlet just behind the lip. Circumferentially extending rows of apertures are provided on the lip establishing two sets of apertures spaced circumferentially away from the slot in opposite directions. The slot removes the boundary layer from the critical portion of the diffuser to minimize or eliminate flow pressure loss or separations resulting from diffusion or tuning. The apertures are in a region of low static pressure on the lip of the inlet and serve as a source of suction to cause air flow into the slot.
Research is reviewed in the following areas: turboprop powered transport aircraft; wind tunnel aerodynamic and acoustics tests of model propellers; turboprop maintenance; and wind tunnel tests on airframe-turboprop interactions. Continued development of the technology for advanced turboprop transport was emphasized.
An experimental investigation was conducted to determine the effect of design changes on the aerodynamic and acoustic performance of translating centerbody sonic inlets. Scale model inlets were tested in the Lewis Research Center's V/STOL wind tunnel. The effects of centerbody position, entry lip contraction ratio, diffuser length, and diffuser area ratio on inlet total pressure recovery, distortion, and noise suppression were investigated at static conditions and at forward velocity and angle of attack. With the centerbody in the takeoff position (retracted), good aerodynamic and acoustic performance was attained at static conditions and at forward velocity. At 0 deg incidence angle with a sound pressure level reduction of 20 dB, the total pressure recovery was 0.986. Pressure recovery at 50 deg was 0.981. With the centerbody in the approach position (extended), diffuser flow separation occurred at an incidence angle of approximately 20 deg. However, good performance was attained at lower angles. With the centerbody in the takeoff position the ability of the inlet to tolerate high incidence angles was improved by increasing the lip contraction ratio. However, at static conditions with the centerbody in the approach position, an optimum lip contraction ratio appears to exist, with both thinner and thicker lips yielding reduced performance.
A self-bleeding method for boundary layer control is described and tested for a subsonic inlet designed to operate in the flowfield generated by high angles of attack. Naturally occurring surface static pressure gradients are used to remove the boundary layer from a separation-prone region of the inlet and to reinject it at a less critical location with a net performance gain. The results suggest that this self-bleeding method for boundary-layer control might be successfully applied to other inlets operating at extreme aerodynamic conditions.
The effect of diffuser wall acoustic treatment on inlet total pressure loss was experimentally determined. Data were obtained by testing an inlet model with 10 different acoustically treated diffusers differing only in the design of the Helmholtz resonator acoustic treatment. Tests were conducted in a wind tunnel at forward velocities to 41 meters per second for inlet throat Mach numbers of .5 to .8 and angles of attack as high as 50 degrees. Results indicate a pressure loss penalty due to acoustic treatment that increases linearly with the porosity of the acoustic facing sheet. For a surface porosity of 14 percent the total pressure loss was 21 percent greater than that for an untreated inlet.
Experimental determination of the effect of diffuser wall acoustic treatment on inlet total pressure loss. Data were obtained by testing an inlet model with 10 different acoustically treated diffusers differing only in the design of the Helmholtz resonator acoustic treatment. Tests were conducted in a wind tunnel at forward velocities to 41 meters per second for inlet throat Mach numbers of .5 to .8 and angles of attack as high as 50 degrees. Results indicate a pressure loss penalty due to acoustic treatment that increases linearly with the porosity of the acoustic facing sheet. For a surface porosity of 14% the total pressure loss was 21% greater than that for an untreated inlet. The penalty resulting from treatment increased for average throat Mach numbers above approximately .7 where local regions of sonic or supersonic surface flow were encountered. Pressure loss over the treatment was not affected by the presence of noise.
Different inlet designs for high angle of attack STOL and VTOL applications were tested in a subsonic wind tunnel. Three removable entry lips having contraction ratios of 1.30, 1.34 and 1.38 were tested with a single diffuser. The internal contour of each entry lip was an ellipse with a major to minor axis of 2.0. Each lip and diffuser assembly was tested to determine its tolerance to angle of attack, first with a conventional centerbody and then with an extended centerbody. Results indicate that a large improvement in separation angle (determined as a function of lip contraction ratio and inlet flow) was obtained for the extended centerbody for all contraction ratios. Improved inlet tolerance to angle of attack was obtained by reducing the adverse pressure gradient downstream of the throat.
Low speed wind tunnel tests were conducted to determine the aeroacoustic performance of several model sonic inlets. The results were analyzed to indicate how inlet aeroacoustic characteristics were affected by inlet design and operating conditions. A system for regulating sonic inlet noise reduction was developed and tested. Results indicate that pressure losses at forward velocity may be substantially less than those at static conditions. This is particularly true for translating centerbody inlets with the centerbody extended in the approach and landing position. Operation to simulated take-off incidence angles of 50 deg was demonstrated with good inlet performance. Results suggest that at takeoff, with 0 deg incidence angle, sonic inlet total pressure losses need not exceed those generated by skin friction for sound pressure level reductions to at least 15 dB. Inlet sound pressure level reduction was regulated to within approximately plus or minus 1 dB by controlling inlet surface static pressure measured at the diffuser exit.
Results of scale model tests of high-throat-Mach-number inlets designed to suppress inlet-emitted engine machinery noise produced in a V/STOL wind tunnel are presented. A vacuum system was used to induce inlet airflow with a siren as a noise source. Inlet mass flow was 11.68 kilograms (25.75 lb. min) per second at a throat Mach number of 0.79. The effect of entry-lip design (contraction ratio and diameter ratio) on inlet total-pressure recovery, steady-state pressure distortion, performance at high incidence angles, and noise suppression was determined. With proper entry-lip design, total-pressure recovery in excess of 0.988 could be obtained statically at an average throat Mach number of 0.79. Total-pressure distortion was 5 percent. The reduction in the siren tone sound pressure level transmitted through the inlet was 10 to 14 db relative to that measured at throat Mach 0.6.
Low speed wind tunnel tests were conducted to determine the aeroacoustic performance of several model sonic inlets. The results were analyzed to indicate how inlet aeroacoustic characteristics were affected by inlet design and operating conditions. A system for regulating sonic inlet noise reduction was developed and tested. Results indicate that pressure losses at forward velocity may be substantially less than those at static conditions. This is particularly true for translating centerbody inlets with the centerbody extended in the approach and landing position. Operation to simulated takeoff incidence angles of 50 degrees was demonstrated with good inlet performance. Inlet sound pressure level reduction was regulation was regulated to within approximately + or - 1 dB by controlling inlet surface static pressure measured at the diffuser exit.
The progress in propulsion system noise reduction is reviewed. The noise technology areas discussed include: fan noise; advances in suppression including conventional acoustic treatment, high Mach number inlets, and wing shielding; engine core noise; flap noise from both under-the-wing and over-the-wing powered-lift systems; supersonic jet noise suppression; and the NASA program in noise prediction.