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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 523 records · Page 29
All sourcesNASA NTRSDOE OSTI
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.↗
NASA NTRS1984

A recontoured, upper surface designed to increase the maximum lift coefficient of a modified NACA 65 (0.82) (9.9) airfoil section

A recontoured upper surface was designed to increase the maximum lift coefficient of a modified NACA 65 (0.82)(9.9) airfoil section which was tested at Mach numbers of 0.3 and 0.4 and Reynolds numbers of 2.3x10(6) and 4.3x10(6). The original 6-series section was tested for comparison with the recontoured section. The recontoured profile was found to have a higher maximum lift coefficient at all test conditions than the original airfoil. The recontoured airfoil showed less drag and nearly the same pitching moment characteristics as the original 6-series airfoil at all test conditions. The improvements found for the recontoured airfoil of the present study are similar to those found during previous investigations of recontoured 6-series airfoils with less camber.

Hicks, R. M.↗
NASA NTRS1983

The Concorde and aeronautical research

Theoretical and experimental work carried out in various research centers, and particularly at ONERA, which led to the conception and to the main technical solutions included in the design of Concorde: plane form, twist and camber of the wing, lift augmentation by upper surface vortices, kinetic heating, air intakes and jet exhausts, materials, aeroelasticity. The development of research, and the numerous tests carried out for the benefit of the designers since the beginning of the project, are also outlined.

Salmon, M.↗
NASA NTRS1984

Secondary flow spanwise deviation model for the stators of NASA middle compressor stages

A model of the spanwise variation of deviation for stator blades is presented. Deviation is defined as the difference between the passage mean flow angle and the metal angle at the outlet of a blade element of an axial compressor stage. The variation of deviation is taken as the difference above or below that predicted by blade element, (i.e., two-dimensional) theory at any spanwise location. The variation of deviation is dependent upon the blade camber, solidity and inlet boundary layer thickness at the hub or tip end-wall, and the blade channel aspect ratio. If these parameters are known or can be calculated, the model provides a reasonable approximation of the spanwise variation of deviation for most compressor middle stage stators operating at subsonic inlet Mach numbers.

Roberts, W. B.↗
NASA NTRS1984

An analytical design procedure for the determination of effective leading edge extensions on thick delta wings

An analytical design procedure for leading edge extensions (LEE) was developed for thick delta wings. This LEE device is designed to be mounted to a wing along the pseudo-stagnation stream surface associated with the attached flow design lift coefficient of greater than zero. The intended purpose of this device is to improve the aerodynamic performance of high subsonic and low supersonic aircraft at incidences above that of attached flow design lift coefficient, by using a vortex system emanating along the leading edges of the device. The low pressure associated with these vortices would act on the LEE upper surface and the forward facing area at the wing leading edges, providing an additional lift and effective leading edge thrust recovery. The first application of this technique was to a thick, round edged, twisted and cambered wing of approximately triangular planform having a sweep of 58 deg and aspect ratio of 2.30. The panel aerodynamics and vortex lattice method with suction analogy computer codes were employed to determine the pseudo-stagnation stream surface and an optimized LEE planform shape.

Farhad Ghaffari↗
NASA NTRS1984

The design and operational development of self-streamlining 2-dimensional flexible walled test sections

Self streamlining two dimensional flexible walled test sections eliminate the uncertainties found in data from conventional test sections particularly at transonic speeds. The test section sidewalls are rigid, while the floor and ceiling are flexible and are positioned to streamline shapes by a system of jacks, without reference to the model. The walls are therefore self streamlining. Data are taken from the model when the walls are good streamlines such that the inevitable residual wall induced interference is acceptably small and correctable. Successful two dimensional validation testing at low speeds has led to the development of a new transonic flexible walled test section. Tunnel setting times are minimized by the development of a rapid wall setting strategy coupled with on line computer control of wall shapes using motorized jacks. Two dimensional validation testing using symmetric and cambered aerofoils in the Mach number range up to about 0.85 where the walls are just supercritical, shows good agreement with reference data using small height-chord ratios between 1.5 and unity.

Wolf, S. W. D.↗
NASA NTRS1984

Numerical simulation of the tip vortex off a low-aspect-ratio wing at transonic speed

The viscous transonic flow around a low aspect ratio wing was computed by an implicit, three dimensional, thin-layer Navier-Stokes solver. The grid around the geometry of interest is obtained numerically as a solution to a Dirichlet problem for the cube. A low aspect ratio wing with large sweep, twist, taper, and camber is the chosen geometry. The topology chosen to wrap the mesh around the wing with good tip resolution is a C-O type mesh. The flow around the wing was computed for a free stream Mach number of 0.82 at an angle of attack of 5 deg. At this Mach number, an oblique shock forms on the upper surface of the wing, and a tip vortex and three dimensional flow separation off the wind surface are observed. Particle path lines indicate that the three dimensional flow separation on the wing surface is part of the roots of the tip vortex formation. The lifting of the tip vortex before the wing trailing edge is observed by following the trajectory of particles release around the wing tip.

Mansour, N. N.↗
NASA NTRS1984

Application of an optimization method to high performance propeller designs

The application of an optimization method to determine the propeller blade twist distribution which maximizes propeller efficiency is presented. The optimization employs a previously developed method which has been improved to include the effects of blade drag, camber and thickness. Before the optimization portion of the computer code is used, comparisons of calculated propeller efficiencies and power coefficients are made with experimental data for one NACA propeller at Mach numbers in the range of 0.24 to 0.50 and another NACA propeller at a Mach number of 0.71 to validate the propeller aerodynamic analysis portion of the computer code. Then comparisons of calculated propeller efficiencies for the optimized and the original propellers show the benefits of the optimization method in improving propeller performance. This method can be applied to the aerodynamic design of propellers having straight, swept, or nonplanar propeller blades.

Li, K. C.↗
NASA NTRS1984

An investigation of aerodynamic characteristics of wings having vortex flow using different numerical codes

Three different numerical codes are employed to determine the aerodynamic characteristics of wings with separation induced vortex flows. Both flat as well as cambered wings of various planform shapes are studied. The effects of wing thickness, fuselage, notch ratio and multiple vortex modeling on aerodynamic performance of the wing are also examined. The theoretically predicted results are compared with experimental results to validate the various computer codes used in this study. An analytical procedure for designing aerodynamically effective leading edge extension (LEE) for a thick delta wing is also presented.

Sushil K Chaturvedi↗
NASA NTRS1984

Shapes for rotating airfoils

An airfoil which has particular application to the blade or blades of rotor aircraft and aircraft propellers is presented. The airfoil thickness distribution, camber and leading edge radius are shaped to locate the airfoil crest at a more aft position along the chord, and to increase the freestream Mach number at which sonic flow is attained at the airfoil crest. The reduced slope of the airfoil causes a reduction in velocity at the airfoil crest at lift coefficients from zero to the maximum lift coefficient. The leading edge radius is adjusted so that the maximum local Mach number at 1.25 percent chord and at the designed maximum lift coefficient is limited to about 0.48 when the Mach number normal to the leading edge is approximately 0.20. The lower surface leading edge radius is shaped so that the maximum local Mach number at the leading edge is limited to about 0.29 when the Mach number normal to the leading edge is approximately 0.20. The drag divergence Mach number associated with the airfoil is moved to a higher Mach number over a range of lift coefficients resulting in superior aircraft performance.

Bingham, G. J.↗
NASA NTRS1981

Computer program documentation for a subcritical wing design code using higher order far-field drag minimization

A subsonic, linearized aerodynamic theory, wing design program for one or two planforms was developed which uses a vortex lattice near field model and a higher order panel method in the far field. The theoretical development of the wake model and its implementation in the vortex lattice design code are summarized and sample results are given. Detailed program usage instructions, sample input and output data, and a program listing are presented in the Appendixes. The far field wake model assumes a wake vortex sheet whose strength varies piecewise linearly in the spanwise direction. From this model analytical expressions for lift coefficient, induced drag coefficient, pitching moment coefficient, and bending moment coefficient were developed. From these relationships a direct optimization scheme is used to determine the optimum wake vorticity distribution for minimum induced drag, subject to constraints on lift, and pitching or bending moment. Integration spanwise yields the bound circulation, which is interpolated in the near field vortex lattice to obtain the design camber surface(s).

Kuhlman, J. M.↗
NASA NTRS1982

Production version of the extended NASA-Langley vortex lattice FORTRAN computer program. Volume 2: Source code

The source code for the latest production version, MARK IV, of the NASA-Langley Vortex Lattice Computer Program is presented. All viable subcritical aerodynamic features of previous versions were retained. This version extends the previously documented program capabilities to four planforms, 400 panels, and enables the user to obtain vortex-flow aerodynamics on cambered planforms, flowfield properties off the configuration in attached flow, and planform longitudinal load distributions.

Herbert, H. E.↗