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At least 253 records · Page 14

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NASA NTRS1972

High-loading, 1800 ft/sec tip speed transonic compressor fan stage. 1: Aerodynamic and mechanical design

A single stage fan with a tip speed of 1800 ft/sec (548.6m/sec) and hub/tip ratio of 0.5 was designed to produce a pressure ratio of 2.285:1 with an adiabatic efficiency of 84.0%. The design flow per inlet annulus area is 38.7 lbm/sq ft-sec (188.9KG/sqm-sec). Rotor blades have modified multiple-circular-arc and precompression airfoil sections. The stator vanes have multiple-circular-arc airfoil sections.

Morris, A. L.↗

NASA NTRS1979

A flight investigation of performance and loads for a helicopter with NLR-1T main-rotor blade sections

Data on performance and rotor loads for a teetering-rotor, AH-1G helicopter flown with a main rotor that had the NLR-1T airfoil as the blade-section contour are presented. The test envelope included hover, forward-flight speed sweeps from 35 to 85 m/sec, and collective-fixed maneuvers at about 0.25 tip-speed ratio. The data set for each test point described vehicle flight state, control positions, rotor loads, power requirements, and blade motions. Rotor loads are reviewed primarily in terms of peak-to-peak and harmonic content. Lower frequency components predominated for most loads and generally increased with increased airspeed, but not necessarily with increased maneuver load factor.

Morris, C. E. K., Jr.↗

NASA NTRS1946

An Experimental Investigation of Flow Conditions in the Vicinity of an NACA D(sub S)-type Cowling

Data are presented of the flow conditions in the vicinity of an NACA D sub S -type cowling. Tests were made of a 1/2 scale-nacelle model at inlet-velocity ratios ranging from 0.23 to 1.02 and angles of attack from 6 deg to 10 deg. The velocity and direction of flow in the vertical plane of symmetry of the cowling were determined from orifices and tufts installed on a board aligned with the flow. Diagrams showing velocity ratio contours and lines of constant flow angles are given.

FLAPPED↗

NASA NTRS1980

Performance deterioration of commercial high-bypass ratio turbofan engines

The results of engine performance deterioration investigations based on historical data, special engine tests, and specific tests to define the influence of flight loads and component clearances on performance are presented. The results of analyses of several damage mechanisms that contribute to performance deterioration such as blade tip rubs, airfoil surface roughness and erosion, and thermal distortion are also included. The significance of these damage mechanisms on component and overall engine performance is discussed.

Mehalic, C. M.↗

NASA NTRS1977

Ultra high tip speed (670.6 m/sec) fan stage with composite rotor: Aerodynamic and mechanical design

A highly loaded, single-stage compressor having a tip speed of 670.6 m/sec was designed for the purpose of investigating very high tip speeds and high aerodynamic loadings to obtain high stage pressure ratios at acceptable levels of efficiency. The design pressure ratio is 2.8 at an adiabatic efficiency of 84.4%. Corrected design flow is 83.4 kg/sec; corrected design speed is 15,200 rpm; and rotor inlet tip diameter is 0.853 m. The rotor uses multiple-circular-arc airfoils from 0 to 15% span, precompression airfoils assuming single, strong oblique shocks from 21 to 43% span, and precompression airfoils assuming multiple oblique shocks from 52% span to the tip. Because of the high tip speeds, the rotor blades are designed to be fabricated of composite materials. Two composite materials were investigated: Courtaulds HTS graphite fiber in a Kerimid 601 polyimide matrix and the same fibers in a PMR polyimide matrix. In addition to providing a description of the aerodynamic and mechanical design of the 670.0 m/sec fan, discussion is presented of the results of structural tests of blades fabricated with both types of matrices.

Halle, J. E.↗

NASA NTRS1923

Wind tunnel tests of five strut sections in yaw

In the first series of wind tunnel tests, the drag and cross wing force of all the struts were measured at a wind speed of 30 mph and at angles of yaw from 0 degrees to 20 degrees. To determine the magnitude of the VL effect, each strut was tested at zero yaw and at a series of speeds ranging from 15 to 38 mph. Although designed as fairings for cables, part of these sections gave such high crosswind forces that they seemed to have possibilities as airfoils. Therefore, the lift (identical with the crosswind force) and drag coefficients were recalculated for four sections on the basis of broadside area to make them comparable with wing coefficients. The general conclusion that the best fineness ratio for a strut is a function of the Reynolds number, decreasing steadily as that quality increases, has of course been reached many times, both by theory and experiment. It was confirmed here once more, and the effect of form on sensitiveness to VL is also strikingly shown. It seems probable that this effect of form is largely due to interaction between the nose and tail, and to the influence which the form of the nose exerts over the whole flow around the strut.

PROPELLERS, METAL↗

NASA NTRS1972

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 2: Data and performance for stage A

Stage A, comprised of a conventional rotor and stator, was designed and tested to establish a performance baseline for comparison with the results of subsequent tests planned for two tandem-blade stages. The rotor had an inlet hub/tip ratio of 0.8 and a design tip velocity of 757 ft/sec. At design equivalent rotor speed, rotor A achieved a maximum adiabatic efficiency of 85.1 percent at a pressure ratio of 1.29. The stage maximum adiabatic efficiency was 78.6 percent at a pressure ratio of 1.27.

Brent, J. A.↗

NASA NTRS1979

A Comparison of the Aerodynamic Characteristics of Eight Sailwing Airfoil Sections

The operational characteristics of sailwings are discussed with emphasis placed on the importance of the trailing edge cable tension. The three dimensional aerodynamic characteristics were obtained from wind tunnel tests, and the results compared to determine the magnitude of the aerodynamic penalties paid for various structural simplifications. For the sectional thickness ratios, it is concluded that, while the basic double-membraned sailwing has exceptional aerodynamic performance, even superior for some applications to the conventional hardwing, any notable deviation from this configuration results in an unacceptably large performance penalty.

Maughmer, M. D.↗

NASA NTRS1934

The effects of equal-pressure fixed slots on the characteristics of a Clark Y Airfoil

A type of fixed open slot so arranged that no flow would pass through it at a lift coefficient corresponding to high-speed flight was investigated in the N.A.C.A. 7 by 10 foot wind tunnel to determine the possibilities of such a high-lift device for increasing the speed-range ratio of a wing. The condition of no through flow was achieved by locating the slot openings at points of equal static pressure at the design lift coefficient as determined from the pressure distribution about the plain wing. Two models of Clark Y wings with such equal-pressure slots were tested and the smoke-flow patterns about them observed. The results of this investigation show that the condition of no air flow through the slot at the desired lift coefficient is attainable. The surface discontinuities produced by the slot openings have, however, such a large effect on the drag that such slots show little promise. An appreciable increase is produced in the maximum lift and the speed-range can be as high as for the plain wing.

Harris, Thomas A↗

NASA NTRS1983

Effects of twin-vertical-tail parameters on twin-engine afterbody/nozzle aerodynamic characteristics

The Langley 16-foot transonic tunnel was used to determine the effects of several empennage and afterbody parameters on twin-engine aft-end aerodynamic characteristics. Model variables included twin-vertical-tail cant angle, toe angle, airfoil camber, and root-chord length and afterbody/engine interfairing shape. Tests were conducted over a Mach number range from 0.6 to 1.2 and over an angle-of-attack range from 2 deg to 10 deg. Nozzle pressure ratio was varied from 1.0 (jet off) to approximately 10.0.

Leavitt, L. D.↗

NASA NTRS1990

The aerodynamic design of the oblique flying wing supersonic transport

The aerodynamic design of a supersonic oblique flying wing is strongly influenced by the requirement that passengers must be accommodated inside the wing. It was revealed that thick oblique wings of very high sweep angle can be efficient at supersonic speeds when transonic normal Mach numbers are allowed on the upper surface of the wing. The goals were motivated by the ability to design a maximum thickness, minimum size oblique flying wing. A 2-D Navier-Stokes solver was used to design airfoils up to 16 percent thickness with specified lift, drag and pitching moment. A new method was developed to calculate the required pressure distribution on the wing based on the airfoil loading, normal Mach number distribution and theoretical knowledge of the minimum drag of oblique configurations at supersonic speeds. The wing mean surface for this pressure distribution was calculated using an inverse potential flow solver. The lift to drag ratio of this wing was significantly higher than that of a comparable delta wing for cruise speeds up to Mach 2.

Vandervelden, Alexander J. M.↗

NASA NTRS1992

Hot streaks and phantom cooling in a turbine rotor passage. I - Separate effects

Experimental documentation and analytical correlations demonstrating the effects of hot streak accumulation and phantom cooling on turbine rotor airfoil surface temperature. Test results are shown for a range of controlling variables to identify where streak accumulation and phantom cooling are most likely to be strongest. These variables include streak injection location, streak-to-free stream density ratio and coolant-to-free stream density and velocity ratios.

Roback, Richard J.↗

NASA NTRS1979

An experimental study of a jet issuing from a lifting wing

An experimental program was conducted to determine the behavior of a round turbulent jet issuing from a lifting two-dimensional wing in crossflow. The jet was located at 65% wing chord on an NACA 0021 airfoil fitted with a 30% chord NACA 4415 flap. The flowfield associated with the jet was surveyed extensively with directional pressure probes to determine local velocity vectors and pressures for three different values of lift coefficient at jet effective velocity ratios (square root of the ratio of the jet dynamic pressure to the freestream dynamic pressure) of 4, 6, and 8. Data describing the jet centerline and the path of the contrarotating vortices accompanying the deflected jet are presented and compared with similar data for a round jet issuing from a large flat plate. The spacing and strength of the vortices are calculated using a simple vortex model previously proposed for the flat plate case. The results show that the penetration of the jet and the vortices increases significantly with increasing lift for the range of test parameters covered in the study. The calculated vortex spacing and strength also show an increase with lift.

Mcmahon, H. M.↗

NASA NTRS1999

Design of a Low Speed Fan Stage for Noise Suppression

This report describes the design of a low tip speed, moderate pressure rise fan stage for demonstration of noise reduction concepts. The fan rotor is a fixed-pitch configuration delivering a design pressure ratio of 1.378 at a specific flow of 43.1 lbm/sec/sq ft. Four exit stator configurations were provided to demonstrate the effectiveness of circumferential and axial sweep in reducing rotor-stator interaction tone noise. The fan stage design was combined with an axisymmetric inlet, conical convergent nozzle, and nacelle to form a powered fan-nacelle subscale model. This model has a 22-inch cylindrical flow path and employs a rotor with a 0.30 hub-to-tip radius ratio. The design is fully compatible with an existing NASA force balance and rig drive system. The stage aerodynamic and structural design is described in detail. Three-dimensional (3-D) computational fluid dynamics (CFD) tools were used to define optimum airfoil sections for both the rotor and stators. A fan noise predictive system developed by Pratt & Whitney under contract to NASA was used to determine the acoustic characteristics of the various stator configurations. Parameters varied included rotor-to-stator spacing and vane leading edge sweep. The structural analysis of the rotor and stator are described herein. An integral blade and disk configuration was selected for the rotor. Analysis confirmed adequate low cycle fatigue life, vibratory endurance strength, and aeroelastic suitability. A unique load carrying stator arrangement was selected to minimize generation of tonal noise due to sources other than rotor-stator interaction. Analysis of all static structural components demonstrated adequate strength, fatigue life, and vibratory characteristics.

Dalton, W. N.↗

NASA NTRSDate not supplied

A Study of Ice-Crystal Icing Erosion Using the NASA Icing Research Tunnel and Propulsion Systems Laboratory

In ice-crystal icing, ice erosion can occur simultaneously with ice accretion. This paper presents a proof-of-concept experimental study that attempts to isolate and quantify ice erosion. The experiments took place at the NASA Icing Research Tunnel and Propulsion Systems Laboratory in 2017 and 2018. The test procedure involved first generating an ice accretion on various NACA 0012 airfoil models using a supercooled liquid water or mixed-phase icing cloud and, subsequently, exposing that ice accretion to an ice crystal cloud. The resulting ice mass decay rates were analyzed in terms of an erosion term, defined as the ratio of the eroded mass flux to impinging mass flux. The erosion term generally increased with airspeed with a value of 1.8% at air velocities of 270 knots and increasing to just over 6% at 360 knots. A preliminary correlation for ice erosion is developed for use in ice accretion models subject to the conditions tested. The study revealed several challenges, both experimental and analytical, including spray bar nozzle freeze-out as well as measurement of the cloud melt ratio to ensure freeze-out.

Icing↗

NASA NTRS1953

Aerodynamic Characteristics of a Two-blade NACA 10-(3)(062)-045 Propeller and of a Two-blade NACA 10-(3)(08)-045 Propeller

Characteristics are given for the two-blade NACA 10-(3)(062)-045 propeller and for the two-blade NACA 10-(3)(08)-045 propeller over a range of advance ratio from 0.5 to 3.8, through a blade-angle range from 20 degrees to 55 degrees measured at the 0.75 radius. Maximum efficiencies of the order of 91.5 to 92 percent were obtained for the propellers. The propeller with the thinner airfoil sections over the outboard portion of the blades, the NACA 10-(3)(062)-045 propeller, had lower losses at high tip speeds, the difference amounting to about 5 percent at a helical tip Mach number of 1.10.

BLADE SECTIONS-PROPELLERS↗

NASA NTRS1995

The effect of adding roughness and thickness to a transonic axial compressor rotor

The performance deterioration of a high speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54-3.18 RMS microns (100-125 RMS microinches) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10% at the hub and 20% at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9% reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effect of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254-0.508 RMS microns (10-20 RMS microinches), compared to the bare metal blade surface finish of 0.508 RMS microns (20 RMS microinches). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60%, 80%, and 100% of design speed. The results indicate that thickness/roughness over the first 10% of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70% of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-3D Navier-Stokes flow solver which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that coating the blade causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

Suder, Kenneth L.↗

NASA NTRS2000

Comparison of Analytical and CAA Solution(s): Unsteady Response of a Rectilinear Swept Cascade to an Incident Gust

This problem is aimed at testing the ability of computational aeroacoustic methods to calculate the acoustic pressure field generated as a result of the interaction of a convected harmonic gust with a rectilinear cascade of swept flat plates. All relevant geometric information is the same as the unsteady response of an isolated finite span swept airfoil to an incident gust problem with the exception of the sweep angle alpha, which is now taken to be a variable. Assume a cascade stagger angle of zero (with the x-axis aligned with the chord) and a gap-to-chord ratio of h/c = 1. The mean flow Mach number M(sub 0), gust frequency omega, gust amplitude A, and chordwise wavenumber k(sub x) are the same. But for the cascade problem take k(sub y) = pi and k(sub z) = 0. The appropriate physical scales are also the same, as is the governing equation and boundary conditions. Naturally, the impermeability condition now applies to the entire cascade. For this problem, determine the amplitude of upstream-radiated rms acoustic pressure as a function of the sweep angle at the specified frequency. Specifically, show the variations of rms acoustic pressure amplitude at the upstream location (-5c, 0, l/2) for sweep angle alpha in the range (0.0 deg, 30.0 deg). Use sweep angle increments no larger than 2.5 deg. Express the results in dB using the rms pressure value for alpha = 0.0 deg as the reference level. The appropriate FORTRAN output statement should read "WRITE(IUNIT,*) alpha, 20-log10 (p(sub rms)(alpha)/p(sub rms)(0))".

Envia, Edmane↗