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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 505 records · Page 28

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NASA NTRS1982

Study of supersonic wings employing the attainable leading-edge thrust concept

A theoretical study was made of supersonic wing geometries at Mach 1.8, using the attainable leading-edge thrust concept. The attainable thrust method offers a powerful means to improve overall aerodynamic efficiency by identifying wing leading-edge geometries that promote attached flow and by defining a local angle-of-attack range over which attached flow may be obtained. The concept applies to flat and to cambered wings, which leads to the consideration of drooped-wing leading edges for attached flow at high lift coefficients.

Middleton, W. D.↗

NASA NTRS1982

Turbofan noise generation. Volume 1: Analysis

Computer programs were developed which calculate the in-duct acoustic modes excited by a fan/stator stae operating at subsonic tip speed. Three noise source mechanisms are included: (1) sound generated by the rotor blades interacting with turbulence ingested into, or generated within, the inlet duct; (2) sound generated by the stator vanes interacting with the turbulent wakes of the rotors blades; and (3) sound generated by the stator vanes interacting with the mean velocity deficit wakes of the rotor blades. The fan/stator stage is modeled as an ensemble of blades and vanes of zero camber and thickness enclosed within an infinite hard-walled annular duct. Turbulence drawn into or generated within the inlet duct is modeled as nonhomogeneous and anisotropic random fluid motion, superimposed upon a uniform axial mean flow, and convected with that flow. Equations for the duct mode amplitudes, or expected values of the amplitudes, are derived.

Ventres, C. S.↗

NASA NTRS1983

Three dimensional flow field inside compressor rotor, including blade boundary layers

The space marching code was modified in order to be able to predict the flow field inside a rotor passage, including the blade and hub wall boundary layers. The basic changes incorporated are modifications of the equations so that the code can handle three dimensional configurations with changes in the radial direction (for example changes in stagger angle, blade camber and thickness), extensions and modifications in order to implement a physically realistic turbulence model such as a k sigma model and an algebraic Reynolds stress model.

Pouagare, M.↗

NASA NTRS1982

The status of two- and three-dimensional testing in the University of Southampton transonic self-streamlining wind tunnel

An automated test section was used to develop a flexible walled testing technique which eliminates some sources of uncertainty in boundary interference effects which exist in conventional transonic test sections. The flexible floor and ceiling of the test section were adjusted to contours which produce a constant Mach number distribution along each wall with no model present. These aerodynamically straight contours form the basis for all streamlining. The wall data are to contain information on the models performance and on lift. Two dimensional validation testing has continued with a cambered NPL 9510 section. Lift data up to Mach 0.87 are compared with reference data. Drag information on a NACA 0012-64 section is presented to indicate the powerful effects of streamlining. Preliminary three dimensional testing in the two dimensional test section has demonstrated that model and support blockage can be relieved by wall contouring.

Wolf, S. W. D.↗

NASA NTRS1983

Loads and performance data from a wind-tunnel test of model articulated helicopter rotors with 2 different blade torsional stiffnesses

A passive means of tailoring helicopter rotor blades to improve performance and reduce loads was evaluated. The parameters investigated were blade torsional stiffness, blade section camber, and distance between blade structural elastic axis and blade tip aerodynamic center. This offset was accomplished by sweeping the tip. The investigation was conducted at advance ratios of 0.20, 0.30, and 0.40. Data are presented without analysis; however, cross referencing of performance data and harmonic loads data may be useful to the analyst for validating aeroelastic theories and design methodologies as well as for evaluating passive aeroelastic tailoring or rotor blade parameters.

Yeager, W. T., Jr.↗

NASA NTRS1983

Trends of Reynolds number effects on two-dimensional airfoil characteristics for helicopter rotor analyses

The primary effects of Reynolds number on two dimensional airfoil characteristics are discussed. Results from an extensive literature search reveal the manner in which the minimum drag and maximum lift are affected by the Reynolds number. C sub d sub min and C sub l sub max are plotted versus Reynolds number for airfoils of various thickness and camber. From the trends observed in the airfoil data, universal scaling laws and easily implemented methods are developed to account for Reynolds number effects in helicopter rotor analyses.

Yamauchi, G. K.↗

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 NTRS1983

Small transport aircraft technology propeller study

A study to define potential benefits of advanced technology propeller for 1985-1990 STAT commuter airplanes was completed. Two baselines, a Convair, 30 passenger, 0.47 Mach number airplane and a Lockheed, 50 passenger, 0.70 Mach number airplane, were selected from NASA-Ames sponsored airframe contracts. Parametric performance, noise level, weight and cost trends for propellers with varying number of blades, activity factor, camber and diameter incorporating blade sweep, tip proplets, advanced composite materials, advanced airfoils, advanced prevision synchrophasing and counter-rotation are presented. The resulting DOC, fuel burned, empty weight and acquisition cost benefits are presented for resizings of the two baseline airplanes. Six-bladed propeller having advanced composite blades, advanced airfoils, tip proplets and advanced prevision synchrophasers provided the maximum DOC improvements for both airplanes. DOC and fuel burned were reduced by 8.3% and 17.0% respectively for the Convair airplane and by 24.9% and 41.2% respectively for the Lockheed airplane. The larger reductions arose from a baseline definition with very heavy fuselage acoustic treatment. An alternate baseline, with a cabin noise 13dB in excess of the objective, was also studied.

Black, B. M.↗

NASA NTRS1981

Aerodynamic evaluation of circulation control propellers

The aerodynamic compatibility of fixed-pitch circulation-control (CC) propellers with the performance characteristics and requirements of a 1600 kg (3600 lb) single-engine variable-pitch propeller general aviation airplane was evaluated. The initial results with elliptical-shaped CC airfoils indicated that the feasibility of application to fixed-pitch propellers was doubtful for the class of airplane investigated (cruise speed of about 300 km/hr) and improbable for higher-speed airplanes. Supplemental data for a cambered circulation-control supercritical airfoil, which became available after completion of the initial evaluation, were subsequently analyzed and are included in the Addendum to this report. With the much superior aerodynamic characteristics of the CC-supercritical airfoil, elimination of variable pitch appears aerodynamically feasible for low-speed airplanes through the use of a moderate amount of upper-surface blowing near the trailing edge during cruise. Overall feasibility depends upon results of structural and systems-type analyses.

Braslow, A. L.↗

NASA NTRS1983

Longitudinal aerodynamic characteristics of a wing-winglet model designed at M = 0.8, C sub L = 0.4 using linear aerodynamic theory

Wind tunnel test results have been presented herein for a subsonic transport type wing fitted with winglets. Wind planform was chosen to be representative of wings used on current jet transport aircraft, while wing and winglet camber surfaces were designed using two different linear aerodynamic design methods. The purpose of the wind tunnel investigation was to determine the effectiveness of these linear aerodynamic design computer codes in designing a non-planar transport configuration which would cruise efficiently. The design lift coefficient was chosen to be 0.4, at a design Mach number of 0.8. Force and limited pressure data were obtained for the basic wing, and for the wing fitted with the two different winglet designs, at Mach numbers of 0.60, 0.70, 0.75 and 0.80 over an angle of attack range of -2 to +6 degrees, at zero sideslip. The data have been presented without analysis to expedite publication.

Kuhlman, J. M.↗

NASA NTRS1983

Effects of various empennage parameters on the aerodynamic characteristics of a twin-engine afterbody model

An experimental investigation to determine the effects of various empennage parameters on aft-end drag and tail interference drag characteristics of a twin-engine, axisymmetric nozzle, afterbody model has been conducted in the Langley 16-Foot Transonic Tunnel. Model variables included: horizontal tail location; vertical tail location, toe angle, cant angle, configuration (single and twin) and airfoil (Symmetrical and cambered); and tail booms. Tests were conducted over a Mach number range from 0.6 to 1.2 and over an angle-of-attack range from -2 to 10 deg. Nozzle pressure ratio was varied from jet off to approximately 10.0.

Leavitt, L. D.↗

NASA NTRS1983

Low-speed investigation of the maneuver capability of supersonic fighter wings

A low-speed investigation of wings suitable for supersonic fighter application was undertaken to understand the flow regions developed by these wings through flow visualization and to evaluate areas in which low-speed testing could provide information for extrapolation to higher Mach numbers and Reynolds numbers. This study involved cranked wings of highly-swept inboard panels and outboard panel of various sweeps with leading- and trailing-edge devices. Both theoretical and experimental results with flow visualization are presented. Flow visualization results show that a mixed flow condition (vortex flow on the highly-swept inboard section and attached flow on the cranked outboard section) is possible for wings with low sweep (20 deg) outboard cranked sections. Lift and drag measurements on the planar (no flap deflection) wings show good agreement with theoretical results for leading-edge vortex flow. There is little or no change in planar wing performance at maneuver conditions due to outboard sweep effects. Results also show that for cambered wings (flap deflected) only small improvements in performance over flat wings were obtained at high-lift maneuver conditions.

Hom, K. W.↗

NASA NTRS1983

Empennage/afterbody integration for single and twin-engine fighter aircraft

An extensive experimental program to determine the effects of empennage surfaces on single and twin-engine afterbody/nozzle drag has been conducted by the Propulsion Aerodynamics Branch at the NASA Langley Research Center. Empennage interference drag was obtained by using experimental values of afterbody/nozzle drag and computed values of empennage drag. The effects of tail location, span, number (single versus twin), toe angle, cant angle, camber and root chord length are discussed. The magnitude of empennage interference drag on single and twin engine configurations is examined.

Berrier, B. L.↗

NASA NTRS1983

SC3 - A wing concept for supersonic maneuvering

A supersonic wing design concept that overcomes the problems associated with nonlinear inviscid pressure distribution is presented together with a summary of the supporting theoretical and experimental development program. The key design feature is a conical panel supporting a controlled supercritical conical camber (SC3) crossflow expansion and recompression which permits lift on the upper surface to be obtained without producing an adverse pressure gradient or crossflow shock wave strong enough to separate the boundary layer. The role of aerodynamics in the design implementation is discussed, as are the concept development, the initial design validation on a fabricated and tested SC3 design, and the body and canard interaction effect. The design and testing of an isolated wing with a planform and thickness distribution representative of a practical application are described. Wind-tunnel results show that significant performance gains can be obtained for fighter aircraft using this concept.

Mason, W. H.↗

NASA NTRS1983

Family of airfoil shapes for rotating blades

An airfoil which has particular application to the blade or blades of rotor aircraft such as helicopters and aircraft propellers is described. The airfoil thickness distribution and camber are shaped to maintain a near zero pitching moment coefficient over a wide range of lift coefficients and provide a zero pitching moment coefficient at section Mach numbers near 0.80 and to increase the drag divergence Mach number resulting in superior aircraft performance.

Noonan, K. W.↗

NASA NTRS1983

A computer program for estimating the aerodynamic characteristics of NACA 16-series airfoils

A computer program written in a table ""look-up'' format, is presented which provides a comprehensive data base on NACA 16-series airfoils. The geometry covered is limited to cambers for a design-lift coefficient from 0.0 to 0.7 and thickness ratios from 4 to 21%. The data include Mach numbers from 0.3 to 1.6, angles of attack from -4 to 8 degrees, and lift coefficients from 0.0 to 0.8. Extrapolation is used to obtain data from Mach numbers, angles of attack, and lift coefficients beyond those for which data are available. A routine to adjust the lift and drag coefficients beyond stall is included. The uses and limitations of the program are also discussed.

Maksymiuk, C. M.↗

NASA NTRS1983

A wing concept for supersonic maneuvering

A theoretical and experimental program in which a wing concept for supersonic maneuvering was developed and then demonstrated experimentally in a series of wind tunnel tests is described. For the typical fighter wing, the problem of obtaining efficient lift at supersonic maneuvering C sub 's occurs due to development of a strong crossflow shock, and boundary layer separation. A natural means of achieving efficient supersonic maneuvering is based on controlling the non-linear inviscid crossflow on the wing in a manner analogous to the supercritical aerodynamic methods developed for transonic speeds. The application of supercritical aerodynamics to supersonic speeds is carried out using Supercritical Conical Camber (SC3). This report provides an aerodynamic analysis of the effort, with emphasis on wing design using non-linear aerodynamics. The substantial experimental data base is described in three separate wind tunnel reports, while two of the computer programs used in the work are also described in a separate report. Based on the development program it appears that a controlled supercritical crossflow can be obtained reliably on fighter-type wing planforms, with an associated drag due to lift reduction of about 20% projected using this concept.

Mason, W. H.↗

NASA NTRS1983

Flow mechanism and experimental investigation of a rotating stall in transonic compressors

The flow characteristics of the rotating stall in compressors is studied, and a flow model is developed along with a theoretical calculation method based on vortex theory. A detailed theoretical calculation is completed for a two dimensional flow field in a transonic rotor in a rotating stall, and the result is in good agreement with experimental findings. The oscillograms of time-varying stall characteristic parameters recorded for the onset, growth, and cessation processes of rotating stall are analyzed, and some new flow phenomena deserving of further investigation are discovered. These include serious separation of individual blades, often preceding the onset of rotating stall in compressors with very small blade-camber angles, and periodical variation of the circumferential width of the stall cell with time, accompanied by periodical oscillation of the width of the stall cell in the radial direction of the blade. The circumferential and radial oscillation frequencies are the same.

Yajun, L.↗