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  • AERODYNAMICS  (415)
  • Life and Medical Sciences  (290)
  • Humans
  • 1975-1979  (705)
  • 1940-1944
  • 1975  (705)
  • 101
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    Transonic aerodynamic characteristics of a wing/body combination incorporating jet flaps (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A 0.25-scale semispan wing/body model with two types of jet flaps was tested in the Ames 11- by 11-Foot Transonic Wind Tunnel. The objective of that testing was to measure the static aerodynamic forces and moments and wing pressure distributions on six configurations differentiated by wing camber, jet flap type, and jet flap angle. Maximum thrust coefficients were limited to 0.12. Angle of attack was varied from -4 deg to 15 deg for Mach numbers between 0.6 and 0.95 at a constant unit Reynolds number of 18.0 million/m (5.5 million/ft). More refined designs and considerably more testing will be required to establish the practicability of the total-exhausting jet flap concept.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62461 , A-6203
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  • 102
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    Experimental investigation of wall shock cancellation and reduction of wall interference in transonic testing (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A series of experiments were performed to evaluate the effectiveness of a three-dimensional land and groove wall geometry and a variable permeability distribution to reduce the interference produced by the porous walls of a supercritical transonic test section. The three-dimensional wall geometry was found to diffuse the pressure perturbations caused by small local mismatches in wall porosity permitting the use of a relatively coarse wall porosity control to reduce or eliminate wall interference effects. The wall porosity distribution required was found to be a sensitive function of Mach number requiring that the Mach number repeatability characteristics of the test apparatus be quite good. The effectiveness of a variable porosity wall is greatest in the upstream region of the test section where the pressure differences across the wall are largest. An effective variable porosity wall in the down stream region of the test section requires the use of a slightly convergent test section geometry.
    Keywords: AERODYNAMICS
    Type: NASA-CR-144979
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  • 103
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    Wind tunnel investigation of the aerodynamic characteristics of five forebody models at high angles of attack at Mach numbers from 0.25 to 2 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Five forebody models of various shapes were tested in the Ames 6- by 6-Foot Wind Tunnel to determine the aerodynamic characteristics at Mach numbers from 0.25 to 2 at a Reynolds number of 800000. At a Mach number of 0.6 the Reynolds number was varied from 0.4 to 1.8 mil. Angle of attack was varied from -2 deg to 88 deg at zero sideslip. The purpose of the investigation was to determine the effect of Mach number of the side force that develops at low speeds and zero sideslip for all of these forebody models when the nose is pointed. Test results show that with increasing Mach number the maximum side forces decrease to zero between Mach numbers of 0.8 and 1.5, depending on the nose angle; the smaller the nose angle of the higher the Mach number at which the side force exists. At a Mach number of 0.6 there is some variation of side force with Reynolds number, the variation being the largest for the more slender tangent ogive.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-73076 , A-6280
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  • 104
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    The effects on propulsion-induced aerodynamic forces of vectoring a partial-span rectangular jet at Mach numbers from 0.40 to 1.20 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was conducted in the Langley 16-foot transonic tunnel to determine the induced lift characteristics of a vectored thrust concept in which a rectangular jet exhaust nozzle was located in the fuselage at the wing trailing edge. The effects of nozzle deflection angles of 0 deg to 45 deg were studied at Mach numbers from 0.4 to 1.2, at angles of attack up to 14 deg, and with thrust coefficients up to 0.35. Separate force balances were used to determine total aerodynamic and thrust forces as well as thrust forces which allowed a direct measurement of jet turning angle at forward speeds. Wing pressure loading and flow characteristics using oil flow techniques were also studied.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8039 , L-10352
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  • 105
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    Viscous/potential flow about multi-element two-dimensional and infinite-span swept wings: Theory and experiment (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The viscous subsonic flow past two-dimensional and infinite-span swept multi-component airfoils is studied theoretically and experimentally. The computerized analysis is based on iteratively coupled boundary layer and potential flow analysis. The method, which is restricted to flows with only slight separation, gives surface pressure distribution, chordwise and spanwise boundary layer characteristics, lift, drag, and pitching moment for airfoil configurations with up to four elements. Merging confluent boundary layers are treated. Theoretical predictions are compared with an exact theoretical potential flow solution and with experimental measures made in the Ames 40- by 80-Foot Wind Tunnel for both two-dimensional and infinite-span swept wing configurations. Section lift characteristics are accurately predicted for zero and moderate sweep angles where flow separation effects are negligible.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62513 , A-6389
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  • 106
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    Aerodynamic results of a separation test (CA20) conducted at the Boeing transonic wind tunnel using 0.030-scale models of the configuration 140A/B (modified) SSV orbiter (model no. 45-0) and the Boeing 747 carrier (model no. AX 1319 I-1), volume 3 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N76-16033.
    Keywords: AERODYNAMICS
    Type: NASA-CR-141846 , DMS-DR-2217-VOL-3
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  • 107
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    Aerodynamic results of a separation test (CA20) conducted at the Boeing transonic wind tunnel using 0.030-scale models of the configuration 140A/B (modified) SSV orbiter (model no. 45-0) and the Boeing 747 carrier (model no. AX 1319 I-1), volume 1 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A Rockwell built 0.030-scale 45-0 modified Space Shuttle Orbiter Configuration 14?A/B model and a Boeing built 0.030-scale 747 carrier model were tested to provide six component force and moment data for each vehicle in proximity to the other at a matrix of relative positions, attitudes and test conditions (angles of attack and sideslip were varied). Orbiter model support system tare effects were determined for corrections to obtain support-free aerodynamics. In addition to the balance force data, pressures were measured. Pressure orifices were located at the base of the Orbiter, on either side of the vertical blade strut, and at the mid-root chord on either side of the vertical tail. Strain gages were installed on the Boeing 747 vertical tail to indicate buffet onset. Photographs of aerodynamic configurations tested are shown.
    Keywords: AERODYNAMICS
    Type: NASA-CR-141844 , DMS-DR-2217-VOL-1
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  • 108
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    Aerodynamic characteristics of a powered, externally blown flap STOL transport model with two engine simulator sizes (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The low-speed aerodynamic characteristics are investigated of a general research model - a swept-wing, jet-powered STOL transport with externally blown flaps. The model was tested with four-engine simulators mounted on pylons under the 9.3-percent-thick supercritical airfoil wing. Two sets of air ejectors were used to provide data with large and small engines. Tests were conducted in the Langley V/STOL tunnel over an angle-of-attack range of -4 deg to 22 deg and a thrust-coefficient range from 0 to approximately 4. The effects are described of power, wing leading-edge slat configuration, T-tail and low horizontal-tail positions, and double-slotted flap deflection. Additional untrimmed and trimmed engine-out data and tail-body data are included.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8057 , L-10129
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  • 109
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    Fully unsteady subsonic and supersonic potential aerodynamics for complex aircraft configurations with applications to flutter (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A general formulation is presented for the analysis of steady and unsteady, subsonic and supersonic aerodynamics for complex aircraft configurations. The theoretical formulation, the numerical procedure, the description of the program SOUSSA (steady, oscillatory and unsteady, subsonic and supersonic aerodynamics) and numerical results are included. In particular, generalized forces for fully unsteady (complex frequency) aerodynamics for a wing-body configuration, AGARD wing-tail interference in both subsonic and supersonic flows as well as flutter analysis results are included. The theoretical formulation is based upon an integral equation, which includes completely arbitrary motion. Steady and oscillatory aerodynamic flows are considered. Here small-amplitude, fully transient response in the time domain is considered. This yields the aerodynamic transfer function (Laplace transform of the fully unsteady operator) for frequency domain analysis. This is particularly convenient for the linear systems analysis of the whole aircraft.
    Keywords: AERODYNAMICS
    Type: NASA-CR-146067
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  • 110
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    A Study of Gust and Control Response of Model Rotor-Propellers in a Wind Tunnel Airstream (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: One-ninth-scale model hingeless and gimballed rotor-propellers were tested at an advance ratio of 0.7 in the Wright Brothers Wind Tunnel at MIT, in the presence of sinusoidal longitudinal and vertical gusts produced by a gust generator of novel design. The gimballed rotor was also subjected to sinusoidal collective and cyclic control inputs. Model test data in terms of blade inplane and out-of-plane bending, longitudinal and lateral gimbal motion, wing vertical and chordwise bending, and blade and wing torsion are presented and compared with theory.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137756 , ASRL-TR-174-4
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  • 111
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    Free-flight model investigation of a vertical-attitude VTOL fighter with twin vertical tails (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Free-flight tests were conducted in the Langley full-scale tunnel to determine the stability and control characteristics of a vertical-attitude VTOL fighter having twin vertical tails and a pivoted fuselage forebody (nose-cockpit) arrangement. The flight tests included hovering flights and transition flights from hover to conventional forward flight. Static force tests were also made to aid in the analysis of the flight tests. The model exhibited satisfactory stability and control characteristics, and the transition from hovering flight to conventional forward flight was relatively smooth and straightforward.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8089 , L-10450
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  • 112
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    Theoretical aerodynamics of upper-surface-blowing jet-wing interaction (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A linear, inviscid subsonic compressible flow theory is formulated to treat the aerodynamic interaction between the wing and an inviscid upper-surface-blowing (USB) thick jet with Mach number nonuniformity. The predicted results show reasonably good agreement with some available lift and induced-drag data. It was also shown that the thin-jet-flap theory is inadequate for the USB configurations with thick jet. Additional theoretical results show that the lift and induced drag were reduced by increasing jet temperature and increased by increasing jet Mach number. Reducing jet aspect ratio, while holding jet area constant, caused reductions in lift, induced drag, and pitching moment at a given angle of attack but with a minimal change in the curve of lift coefficient against induced-drag coefficient. The jet-deflection effect was shown to be beneficial to cruise performance. The aerodynamic center was shifted forward by adding power or jet-deflection angle. Moving the jet away from the wing surface resulted in rapid changes in lift and induced drag. Reducing the wing span of a rectangular wing by half decreased the jet-circulation lift by only 24 percent at a thrust coefficient of 2.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-7936 , L-10037
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  • 113
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    NASA/FRC wake turbulence flight test program: Ride quality aspects (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: To determine how much the ride quality of the Shuttle Carrier (Boeing 747) Aircraft is affected at various spoiler settings, the PEMS II (Portable Environmental Measuring System) was used to measure onboard motion during a test flight October 15, 1975. The PEMS II measures acceleration in the vertical, transverse and longitudinal directions as well as angular rates of pitch, roll, and yaw. The data acquired by this instrument, combined with an airline passenger comfort model, gives an indication of how passengers would react to the motion induced by flying in a vortex alleviation configuration.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145700 , MEMO-403225 , ESS-4032-104-75
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  • 114
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    Pressure data from a 64A010 airfoil at transonic speeds in heavy gas media of ratio of specific heats from 1.67 to 1.12 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A NACA 64A010 pressure-instrumented airfoil was tested at transonic speeds over a range of angle of attack from -1 to 12 degrees at various Reynolds numbers ranging from 2 to 6 million in air, argon, Freon 12, and a mixture of argon and Freon 12 having a ratio of specific heats corresponding to air. Good agreement of results is obtained for conditions where compressibility is not significant and for the air and comparable argon-Freon 12 mixture. Comparison of heavy gas results with air, when adjusted for transonic similarity, show improved, but less than desired agreement.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62468 , A-6225
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  • 115
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    Mach reflection associated with over-expanded nozzle free jet flows (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The occurrence of Mach reflection within the over-expanded nozzle free jet flow has been examined. A flow model emphasizing the interaction between the outer and central core streams has been developed to deal with flow situations where detailed inviscid calculations of the flowfield with Mach reflexion are not possible. The results obtained show reasonably good agreement with the available experimental data. This method has also produced comparable results where detailed calculations of the flowfield are possible.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 13; June 197
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  • 116
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    Application of a transitional boundary-layer theory in the low hypersonic Mach number regime (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: An investigation is made to assess the capability of a finite-difference boundary-layer procedure to predict the mean profile development across a transition from laminar to turbulent flow in the low hypersonic Mach-number regime. The boundary-layer procedure uses an integral form of the turbulence kinetic-energy equation to govern the development of the Reynolds apparent shear stress. The present investigation shows the ability of this procedure to predict Stanton number, velocity profiles, and density profiles through the transition region and, in addition, to predict the effect of wall cooling and Mach number on transition Reynolds number. The contribution of the pressure-dilatation term to the energy balance is examined and it is suggested that transition can be initiated by the direct absorption of acoustic energy even if only a small amount (1 per cent) of the incident acoustic energy is absorbed.
    Keywords: AERODYNAMICS
    Type: International Journal of Heat and Mass Transfer; 18; Nov. 197
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  • 117
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    Structure of turbulent jets and wakes (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The structure of two-dimensional turbulent jets and wakes is studied using two eddy viscosity models. The turbulent energy equation is used together with the mean momentum equations, and the system is closed by introducing eddy coefficients. The fine structure turbulence is obtained by applying proper boundary conditions at the mean turbulent interface position. The results, after a second averaging process, are compared with available experimental data.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 13; May 1975
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  • 118
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    Supercritical wing sections II: A handbook (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The numerical aspect of theoretical work on transonics and supercritical wing sections are compiled. A model of the trailing edge is introduced which eliminates the loss of 15 to 20 percent experienced with heavily aft-loaded models, and it is indicated how drag creep can be reduced at off-design conditions. A rotated finite difference scheme is presented which can handle supersonic as well as subsonic free stream Mach numbers and leads to an effective three-dimensional program for the computation of transonic flow past an oblique wing. In the case of two-dimensional flow, the method is extended to take into account the displacement thickness computed by a semiempirical turbulent boundary layer correction. A series of supercritical wing sections is discussed together with comparisons between experimental and theoretical data. Computer programs and a brief manual for their operation are listed. It is shown that the programs furnish a physically adequate computer simulation of the compressible flows that arise in problems of transonic aerodynamics.
    Keywords: AERODYNAMICS
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  • 119
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    Small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts. Volume 1. Aerodynamic test results (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results and analyses of aerodynamic and acoustic studies conducted on the small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts are presented. Various types of nozzle flap concepts were tested. These are an upper surface blowing concept with a multiple slot arrangement with seven slots (seven slotted nozzle), an upper surface blowing type with a large nozzle exit at approximately mid-chord location in conjunction with a powered trailing edge flap with multiple slots (split flow or partially slotted nozzle). In addition, aerodynamic tests were continued on a similar multi-slotted nozzle flap, but with 14 slots. All three types of nozzle flap concepts tested appear to be about equal in overall aerodynamic performance but with the split flow nozzle somewhat better than the other two nozzle flaps in the landing approach mode. All nozzle flaps can be deflected to a large angle to increase drag without significant loss in lift. The nozzle flap concepts appear to be viable aerodynamic drag modulation devices for landing.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137747
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  • 120
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    Flow visualization of vortices locked by spanwise blowing over wings featuring a unique leading and trailing-edge flap system (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Flow visualization studies were conducted to qualitatively determine the effects of active generation and augmentation of vortex flow over wings by blowing a discrete jet in a spanwise direction in the channel formed by extension of upper surface leading- and trailing-edge flaps. Spanwise blowing from a reflection plane over a rectangular wing was found to generate and lock a dual corotating vortex system within the channel and, at sufficient blowing rates, cause the separated flow off the upper end of the leading-edge flap to reattach to the trailing-edge flap. Test parameters included wing angle of attack, jet momentum coefficient, leading- and trailing-edge flap deflection angle, and jet location above the wing surface. Effects due to removal of the leading- and trailing-edge flap were also investigated.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72788
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  • 121
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    Derivation of equations of motion for multi-blade rotors employing coupled modes and including high twist capability (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The equations of motion are derived for a multiblade rotor. A high twist capability and coupled flatwise-edgewise assumed normal modes are employed instead of uncoupled flatwise - edgewise assumed normal models. The torsion mode is uncoupled. Support system models, consisting of complete helicopters in free flight, or grounded flexible supports, arbitrary rotor-induced inflow, and arbitrary vertical gust models are also used.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137810 , SER-50912
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  • 122
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    Fully unsteady subsonic and supersonic potential aerodynamics for complex aircraft configurations for flutter applications (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A general theory for study, oscillatory or fully unsteady potential compressible aerodynamics around complex configurations is presented. Using the finite-element method to discretize the space problem, one obtains a set of differential-delay equations in time relating the potential to its normal derivative which is expressed in terms of the generalized coordinates of the structure. For oscillatory flow, the motion consists of sinusoidal oscillations around a steady, subsonic or supersonic flow. For fully unsteady flow, the motion is assumed to consist of constant subsonic or supersonic speed for time t or = 0 and of small perturbations around the steady state for time t 0.
    Keywords: AERODYNAMICS
    Type: NASA-CR-146573 , AIAA/ASME/SAE 17th Structures, Structural Dynamics and Materials Conf.
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  • 123
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    Computation of the transonic perturbation flow fields around two- and three-dimensional oscillating wings (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Analytical and empirical studies of a finite difference method for the solution of the transonic flow about an harmonically oscillating wing are presented along with a discussion of the development of a pilot program for three-dimensional flow. In addition, some two- and three-dimensional examples are presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2599 , D6-42536
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  • 124
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    Laminar separation on a blunted cone at high angles of attack (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A new technique is reported for calculating the entire flow field on spherically blunted cones at high angles of attack and high laminar Reynolds numbers. An approximate system of parabolic equations obtained from the steady Navier-Stokes equations by assuming the viscous, streamwise derivative terms are small compared to the viscous normal and circumferential derivatives is the basis of the calculations. These equations are valid for both the inviscid and viscous regions, including the circumferential separation zone that develops on the leeward side at high angles of attack. Two different methods are used to obtain the initial conditions for these equations at the sphere cone tangency plane. For small nose Reynolds numbers, an axisymmetric merged layer solution around a sphere is rotated to provide a three-dimensional initial plane of data. For large nose Reynolds numbers, the nose region is solved using an inviscid, three dimensional time dependent solution combined with a boundary layer solution for the viscous flow. The computed flowfield including the leeward separation region is described and compared with data for a 7 deg half angle cone at 10 deg angle of attack, and a blunt 15 deg half angle cone at 15 deg angle of attack.
    Keywords: AERODYNAMICS
    Type: AGARD Flow Separation; 11 p
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  • 125
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    Flow behavior in inlet guide vanes of radial turbines (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Scroll flow is discussed. Streamline pattern and velocity distribution in the guide vanes are calculated. The blade surface temperature distribution is also determined. The effects of the blade shapes and the nozzle channel width on the velocity profiles at inlet to the guide vanes are investigated.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137632
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  • 126
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    An investigation of airflow over the aft portions of a variable sweep fighter configuration (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation of air flow over the aft portions of a variable sweep fighter aircraft configuration was made. Tests conducted in the unitary plan wind tunnel at Mach number 2.16 included measurements of forces, moments, and local static pressures as well as visual recordings of the air flow. An aerodynamic analytical prediction method was evaluated when used in data comparison at angles of attack of 0, 5, and 15 degrees. The results indicate that in supersonic flow the typical outboard located twin vertical tail arrangement tends to provide a more positive increment in normal-force on the afterbody fuselage and the horizontal tail than a single center-mounted vertical tail of similar planform shape. In addition, the results indicate that a method for aerodynamic analysis of wing-body-tail configurations currently available can provide reasonable estimates of pressure coefficient distributions on configurations in regions of complex supersonic flow. At this time, however, the available analytical method cannot adequately replace experimental wind tunnel tests for determining the supersonic flow environment of a given configuration.
    Keywords: AERODYNAMICS
    Type: NASA-CR-146361
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  • 127
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    Mated aerodynamic characteristics investigation for 0.04-scale model Boeing 747 CAM/external tank (model AX1284 E-5) combination in the University of Washington Aeronautical Laboratory F. K. Kirsten Wind Tunnel (CA11) (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Experimental investigations of the aerodynamic characteristics of a 0.04-scale external tank (ET) force model in combination with a 0.04-scale Boeing 747 force model were conducted. Test purposes were: (1) to determine ET airloads for selected configurations and (2) to determine the effectiveness of ET position, incidence, and support structure and 747 vertical stabilizing surfaces on stability, control, and performance of 747/ET combinations. The 747 was tested alone to establish baseline data and to verify test results. Six-component aerodynamic force and moment data were recorded for the 747 CAM and ET combination. Six-component force and moment data were also recorded for the ET, which was mounted on an internal balance supported by the 747. Data were recorded for angles of attack from -4 deg to +24 deg in 2 deg increments and angles of sideslip of - deg to + or - 20 deg. Testing was conducted at Mach 0.15 with dynamic pressure deg at 36 psf and unit Reynolds number of 1.3 million per foot. Photographs of test configurations are shown.
    Keywords: AERODYNAMICS
    Type: NASA-CR-141835 , DMS-DR-2236
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  • 128
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    Low speed wind tunnel investigation of span load alteration, forward-located spoilers, and splines as trailing-vortex-hazard alleviation devices on a transport aircraft model (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The effectiveness of a forward-located spoiler, a spline, and span load alteration due to a flap configuration change as trailing-vortex-hazard alleviation methods was investigated. For the transport aircraft model in the normal approach configuration, the results indicate that either a forward-located spoiler or a spline is effective in reducing the trailing-vortex hazard. The results also indicate that large changes in span loading, due to retraction of the outboard flap, may be an effective method of reducing the trailing-vortex hazard.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8133 , L-10568
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    Longitudinal aerodynamic characteristics of a deflected-thrust propulsive-lift transport model (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: A wind-tunnel investigation was conducted to determine the effect of deflecting the engine exit of a four-engine double-slotted flap transport to provide STOL performance. Longitudinal aerodynamic data were obtained at various engine exit positions and deflections. The data were obtained at three flap deflections representing cruise, take-off, and landing conditions for a range of angles of attack and various thrust coefficients. Downwash angles at the location of the horizontal tail were measured. The data are presented without analysis or discussion. Photographs of the test configurations are shown.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-3234 , L-10106
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  • 130
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    A review of the NASA V-G/VGH general aviation program (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The V-G and VGH data collected from a wide variety of general aviation airplanes since the inception of the NASA V-G/VGH General Aviation Program in 1962 are presented. These data were analyzed to obtain information on the gust and maneuver loads, on the operating practices, and on the effects of different types of operations on these parameters.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8058 , L-10355
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  • 131
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    A new unified approach to analyze wing-body-tail configurations with control surfaces in steady, oscillatory and fully unsteady, subsonic and supersonic flows (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A general formulation for the analysis of steady and unsteady, subsonic and supersonic potential aerodynamics for arbitrary complex geometries is presented. The theoretical formulation, the numerical procedure, and numerical results are included. In particular, generalized forces for fully unsteady (complex frequency) aerodynamics for an AGARD coplanar wing-tail interfering configuration in both subsonic and supersonic flows are considered.
    Keywords: AERODYNAMICS
    Type: NASA-CR-146073
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  • 132
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    Solutions to Kuessner's integral equation in unsteady flow using local basis functions (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The computational procedure and numerical results are presented for a new method to solve Kuessner's integral equation in the case of subsonic compressible flow about harmonically oscillating planar surfaces with controls. Kuessner's equation is a linear transformation from pressure to normalwash. The unknown pressure is expanded in terms of prescribed basis functions and the unknown basis function coefficients are determined in the usual manner by satisfying the given normalwash distribution either collocationally or in the complex least squares sense. The present method of solution differs from previous ones in that the basis functions are defined in a continuous fashion over a relatively small portion of the aerodynamic surface and are zero elsewhere. This method, termed the local basis function method, combines the smoothness and accuracy of distribution methods with the simplicity and versatility of panel methods. Predictions by the local basis function method for unsteady flow are shown to be in excellent agreement with other methods. Also, potential improvements to the present method and extensions to more general classes of solutions are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137719 , D6-43599
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  • 133
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    Attitude determination of a high altitude balloon system. Part 2: Development of the parameter determination process (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The attitude of the balloon system is determined as a function of time if: (a) a method for simulating the motion of the system is available, and (b) the initial state is known. The initial state is obtained by fitting the system motion (as measured by sensors) to the corresponding output predicted by the mathematical model. In the case of the LACATE experiment the sensors consisted of three orthogonally oriented rate gyros and a magnetometer all mounted on the research platform. The initial state was obtained by fitting the angular velocity components measured with the gyros to the corresponding values obtained from the solution of the math model. A block diagram illustrating the attitude determination process employed for the LACATE experiment is shown. The process consists of three essential parts; a process for simulating the balloon system, an instrumentation system for measuring the output, and a parameter estimation process for systematically and efficiently solving the initial state. Results are presented and discussed.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145958
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  • 134
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    Study of viscous flow about airfoils by the integro-differential method (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An integro-differential method was used for numerically solving unsteady incompressible viscous flow problems. A computer program was prepared to solve the problem of an impulsively started 9% thick symmetric Joukowski airfoil at an angle of attack of 15 deg and a Reynolds number of 1000. Some of the results obtained for this problem were discussed and compared with related work completed previously. Two numerical procedures were used, an Alternating Direction Implicit (ADI) method and a Successive Line Relaxation (SLR) method. Generally, the ADI solution agrees well with the SLR solution and with previous results are stations away from the trailing edge. At the trailing edge station, the ADI solution differs substantially from previous results, while the vorticity profiles obtained from the SLR method there are in good qualitative agreement with previous results.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145693
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  • 135
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    Fluctuating surface pressure measurements on USB wing using two types of transducers (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Measurements of the fluctuating pressures on the wing surface of an upper-surface-blown powered-lift model and a JT15 engine were obtained using two types of pressure transducers. The pressures were measured using overall-fluctuating pressures and power spectral density analyses for various thrust settings and two jet impingement angles. Comparison of the data from the two transducers indicate that similar results are obtained in the lower frequency ranges for both transducers. The data also indicate that for this configuration, the highest pressure levels occur at frequencies below 2000 Hz.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72750
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  • 136
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    Transonic pressure measurements and comparison of theory to experiment for an arrow-wing configuration. Volume 1: Experimental data report, base configuration and effects of wing twist and leading-edge configuration (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A wind tunnel test of an arrow-wing-body configuration consisting of flat and twisted wings, as well as a variety of leading- and trailing-edge control surface deflections, was conducted at Mach numbers from 0.4 to 1.1 to provide an experimental pressure data base for comparison with theoretical methods. Theory-to-experiment comparisons of detailed pressure distributions were made using current state-of-the-art attached and separated flow methods. The purpose of these comparisons was to delineate conditions under which these theories are valid for both flat and twisted wings and to explore the use of empirical methods to correct the theoretical methods where theory is deficient.
    Keywords: AERODYNAMICS
    Type: NASA-CR-132727 , D6-42670-2-VOL-1
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  • 137
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    Asymptotic theory of two-dimensional trailing-edge flows (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: Problems of laminar and turbulent viscous interaction near trailing edges of streamlined bodies are considered. Asymptotic expansions of the Navier-Stokes equations in the limit of large Reynolds numbers are used to describe the local solution near the trailing edge of cusped or nearly cusped airfoils at small angles of attack in compressible flow. A complicated inverse iterative procedure, involving finite-difference solutions of the triple-deck equations coupled with asymptotic solutions of the boundary values, is used to accurately solve the viscous interaction problem. Results are given for the correction to the boundary-layer solution for drag of a finite flat plate at zero angle of attack and for the viscous correction to the lift of an airfoil at incidence. A rational asymptotic theory is developed for treating turbulent interactions near trailing edges and is shown to lead to a multilayer structure of turbulent boundary layers. The flow over most of the boundary layer is described by a Lighthill model of inviscid rotational flow. The main features of the model are discussed and a sample solution for the skin friction is obtained and compared with the data of Schubauer and Klebanoff for a turbulent flow in a moderately large adverse pressure gradient.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Aerodynamic Analysis Requiring Advanced Computers, Pt. 1; p 177-249
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  • 138
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    Aerodynamic characteristics of a tandem wing configuration of a Mach number of 0.30 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was conducted to determine the aerodynamic characteristics of a tandem wing configuration. The configuration had a low forward mounted sweptback wing and a high rear mounted sweptforward wing jointed at the wing tip by an end plate. The investigation was conducted at a Mach number of 0.30 at angles of attack up to 20 deg. A comparison of the experimentally determined drag due to lift characteristics with theoretical estimates is also included.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72779
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  • 139
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    Aerodynamic characteristics of a large-scale hybrid upper surface blown flap model having four engines (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Data are presented from an investigation of the aerodynamic characteristics of large-scale wind tunnel aircraft model that utilized a hybrid-upper surface blown flap to augment lift. The hybrid concept of this investigation used a portion of the turbofan exhaust air for blowing over the trailing edge flap to provide boundary layer control. The model, tested in the Ames 40- by 80-foot Wind Tunnel, had a 27.5 deg swept wing of aspect ratio 8 and 4 turbofan engines mounted on the upper surface of the wing. The lift of the model was augmented by turbofan exhaust impingement on the wind upper-surface and flap system. Results were obtained for three flap deflections, for some variation of engine nozzle configuration and for jet thrust coefficients from 0 to 3.0. Six-component longitudinal and lateral data are presented with four engine operation and with the critical engine out. In addition, a limited number of cross-plots of the data are presented. All of the tests were made with a downwash rake installed instead of a horizontal tail. Some of these downwash data are also presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62460 , A-6202
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  • 140
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    Analytical study of mixing and reacting three-dimensional supersonic combustor flow fields (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: An analytical investigation is presented of mixing and reacting hydrogen jets injected from multiple orifices transverse and parallel to a supersonic airstream. The COMOC computer program, based upon a finite-element solution algorithm, was developed to solve the governing equations for three-dimensional, turbulent, reacting, boundary-region, and confined flow fields. The computational results provide a three-dimensional description of the velocity, temperature, and species-concentration fields downstream of hydrogen injection. Detailed comparisons between cold-flow data and results of the computational analysis have established validity of the turbulent-mixing model based on the elementary mixing-length hypothesis. A method is established to initiate computations for reacting flow fields based upon cold-flow correlations and the appropriate experimental parameters of Mach number, injector spacing, and pressure ratio. Key analytical observations on mixing and combustion efficiency for reacting flows are presented and discussed.
    Keywords: AERODYNAMICS
    Type: Aerodynamic Analyses Requiring Advanced Computers, Pt. 1; p 251-315
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  • 141
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    A numerical method for the prediction of high-speed boundary-layer transition using linear theory (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: A method is described of estimating the location of transition in an arbitrary laminar boundary layer on the basis of linear stability theory. After an examination of experimental evidence for the relation between linear stability theory and transition, a discussion is given of the three essential elements of a transition calculation: (1) the interaction of the external disturbances with the boundary layer; (2) the growth of the disturbances in the boundary layer; and (3) a transition criterion. The computer program which carried out these three calculations is described. The program is first tested by calculating the effect of free-stream turbulence on the transition of the Blasius boundary layer, and is then applied to the problem of transition in a supersonic wind tunnel. The effects of unit Reynolds number and Mach number on the transition of an insulated flat-plate boundary layer are calculated on the basis of experimental data on the intensity and spectrum of free-stream disturbances. Reasonable agreement with experiment is obtained in the Mach number range from 2 to 4.5.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Aerodynamic Analyses Requiring Advanced Computers, Pt. 1; p 101-123
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  • 142
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    Transient airload computer analysis for simulating wind induced impulsive noise conditions of a hovering helicopter rotor (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: A numerical analysis was developed to determine the airloads on helicopter rotors operating under near-hovering flight conditions capable of producing impulsive noise. A computer program was written in which the solutions for the rotor tip vortex geometry, inflow, aeroelastic response, and airloads are solved in a coupled manner at sequential time steps, with or without the influence of an imposed steady ambient wind or transient gust. The program was developed for future applications in which predicted airloads would be incorporated in an acoustics analysis to attempt to predict and analyze impulsive noise (blade slap). The analysis was applied to a hovering full-scale rotor for which impulsive noise was recorded in the presence of ambient wind. The predicted tip vortex coordinates are in reasonable agreement with the test data, and the blade airload solutions converged to a periodic behavior for an imposed steady ambient wind conditions.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137772
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  • 143
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    Experimental investigation of supersonic flow past double-wedge configurations (1975)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: Viscous-inviscid interactions characteristic of those which occur when the fuselage-generated shock wave interacts with the wing-generated shock wave of a shuttle orbiter were studied experimentally. Surface-pressure measurements and schlieren photographs were obtained to define the flowfield generated when a Mach 4.97 stream encounters a double-wedge configuration. The deflection angles for the two wedge surfaces were such that the shock interaction pattern was either a Type-V pattern or a Type-VI pattern, as defined by Edney. The correlation between the present data and the theoretical solution for the Type-VI solution is satisfactory. The correlation between the measured Type-V shock-interaction pattern and the theoretical solution is satisfactory up to the interaction region. Downstream of the interaction the Type-V data depend primarily on the shape of the leading-edge shock wave.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 13; July 197
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  • 144
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    Perturbation solution of the Navier-Stokes equations and its relation to the Lighthill-Curle solution of aerodynamic sound (1975)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: The aerodynamic sound described by the Lighthill-Curle solution is reexamined using a method of matched asymptotic expansions. The governing Navier-Stokes equations written in nondimensional form are expanded for a small Mach number. First- and second-order solutions for the pressure field are obtained, and the singular nature of the expansion at large distances is indicated. The nearfield pressure is governed by the Poisson equation, whereas the farfield equations describe a linear wave system in a dissipative medium. The pseudosound is related to the incompressible Reynolds stresses associated with a solenoidal velocity field, the velocity, the pressure perturbation, and their derivatives on the boundaries. A uniformly valid first-order solution for the pressure is obtained. It is shown that viscosity, thermal conductivity, and entropy in the flow do not contribute to the first-order noise generation, while the viscous stress contributes to noise only from some boundaries. The application of the proposed perturbation method to a subsonically moving surface and a hot jet is discussed.
    Keywords: AERODYNAMICS
    Type: Acoustical Society of America; vol. 58
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  • 145
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    An investigation of several NACA 1-series inlets at Mach numbers from 0.4 to 1.29 for mass flow ratios near 1.0 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation to determine the performance of eight NACA 1-series inlets at massflow ratios near 1.0 was conducted in the Langley 16-foot transonic tunnel. The inlet diameter ratios (ratio of inlet diameter to maximum diameter) were 0.85 and 0.89 for an inlet length ratio (ratio of inlet length to maximum diameter) of 1.0. Inlet lip radius varied from 0.061 cm to 0.251 cm, and internal contraction area ratio (ratio of inlet area to throat area) varied from 1.006 to 1.201. Reynolds number based on model maximum diameter ranged from 3,600,000 at a Mach number of 400,000 to 5,900,000 at a Mach number of 1.29. The results indicate that nearly uniform pressure distributions on a given inlet were obtained over a limited range of mass-flow ratios and Mach numbers. When inlet lip thickness was increased by means of lip radius or contraction ratio, the inlet critical Mach number decreased. Drag-divergence Mach number inferred from forebody pressure integrations was above 0.94 for most of the inlets tested.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-3324 , L-10497
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  • 146
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    A finite element analysis of the exact nonlinear formulation of a lifting surface in steady incompressible flow, with the evaluation of the correct wake geometry (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: The problem of steady incompressible flow for lifting surfaces is considered. An integral equation is solved relating the values of the potential discontinuity on the lifting surface and its wake to the values of the normal derivative of the potential which are known from the boundary conditions. The lifting surface and the wake are divided into small quadrilateral surface elements. The values of the potential discontinuity and the normal derivative of the potential are assumed to be constant within each lifting surface element and equal to their values at the centroids of the lifting surface elements. This yields a set of linear algebraic equations. An iteration procedure is used to obtain the wake geometry: the velocities at the corner points of the wake elements are calculated and the wake streamlines are aligned to be parallel to the velocity vector. The procedure is repeated until convergence is attained.
    Keywords: AERODYNAMICS
    Type: NASA-CR-146074
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  • 147
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    Application of multivariable search techniques to the optimization of airfoils in a low speed nonlinear inviscid flow field (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: Multivariable search techniques are applied to a particular class of airfoil optimization problems. These are the maximization of lift and the minimization of disturbance pressure magnitude in an inviscid nonlinear flow field. A variety of multivariable search techniques contained in an existing nonlinear optimization code, AESOP, are applied to this design problem. These techniques include elementary single parameter perturbation methods, organized search such as steepest-descent, quadratic, and Davidon methods, randomized procedures, and a generalized search acceleration technique. Airfoil design variables are seven in number and define perturbations to the profile of an existing NACA airfoil. The relative efficiency of the techniques are compared. It is shown that elementary one parameter at a time and random techniques compare favorably with organized searches in the class of problems considered. It is also shown that significant reductions in disturbance pressure magnitude can be made while retaining reasonable lift coefficient values at low free stream Mach numbers.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137760 , TN-206
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  • 148
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    Numerical solution of the Navier-Stokes equations for arbitrary two-dimensional airfoils (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: A method for numerical solution of the Navier-Stokes equations for the flow about arbitrary airfoils or other bodies is presented. This method utilizes a numerically generated curvilinear coordinate system having a coordinate line coincident with the body contour. Streamlines, velocity profiles, and pressure and force coefficients for several airfoils and an arbitrary rock are given. Potential flow solutions are also presented. The procedure capable of treating multiple-element airfoils, and potential flow results are presented.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Aerodynamic Analyses Requiring Advanced Computers, Pt. 1; p 469-530
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  • 149
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    Effects of jet exhaust gas properties on exhaust simulation and afterbody drag (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: The effect of varying the jet exhaust's ratio of specific heats, gas constant, and temperature on airplane afterbody drag was investigated. Jet exhaust simulation parameters were evaluated also. Subsonic and transonic tests were made using a single nacelle model with afterbodies having boattail angles of 10 deg and 20 deg. Besides air, three other jet exhaust gases were investigated. The ratios of specific heats, gas constants, and total temperatures of the four exhaust gases ranged from 1.40 to 1.26, 287 to 376 J/kg-K, and 300 to 1013 K, respectively. For steep boattail angles, and transonic speeds and typical turbojet pressure ratios, the current data indicate that the use of air to simulate a dry turbojet exhaust can result in an overprediction of afterbody drag as high as 17 percent of the dry turbojet value.
    Keywords: AERODYNAMICS
    Type: NASA-TR-R-444 , L-10183
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  • 150
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    Poppet valve control of throat stability bypass to increase stable airflow range of a Mach 2.5. inlet with 60 percent internal contraction (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The throat of a Mach 2.5 inlet with a coldpipe termination was fitted with a stability-bypass system. System variations included several stability bypass entrance configurations. Poppet valves controlled the bypass airflow. The inlet stable airflow range achieved with each configuration was determined for both steady state conditions and internal pulse transients. Results are compared with those obtained without a stability bypass system. Transient results were also obtained for the inlet with a choke point at the diffuser exit and for the inlet with large and small stability bypass plenum volumes. Poppet valves at the stability bypass exit provided the inlet with a stable airflow range of 20 percent or greater at all static and transient conditions.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-3297 , E-8382
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  • 151
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    A flying hot wire study of the turbulent near wake of a circular cylinder at Reynolds number of 140,000 (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: The phenomenology was studied of the processes of vortex formation and transport in the near wake, at a Reynolds number sufficiently high to insure a fully turbulent wake, but low enough to insure a laminar separation. The apparatus developed for measuring this flow consisted of X-array hot wire probes mounted on the ends of a pair of whirling arms. A computer controlled data acquisition system was slaved to the position of the rotating arm and managed, monitored, edited, and recorded the vast profusion of data which is continuously poured out by the device. Results are presented which show the instantaneous velocity, intermittency, vorticity, and stress fields as a function of phase for the first six diameters of the near wake. The stresses in the near wake emerge as a concatenation of peaks and valleys, some the result of strong induced motions in the outer flow which cause free stream fluid to move rapidly inward toward the center of the wake, others the result of the random motions of the background turbulence.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145429 , PR-7
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  • 152
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    Center-line velocity decay in coaxial free jets (1975)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: The effects of initial nonuniformities in the exit velocity profiles of incompressible coaxial air jets expelled into free air on the fully developed region of the jets is studied experimentally. The primary and annular jet exit velocities were nonuniform, and the average velocity of the primary jet was greater than that of the annular jet. The center-line velocity of the jet in the fully developed region is found to depend on the primary jet exit velocity profile and on the ratio of the central to annular jet average velocities.
    Keywords: AERODYNAMICS
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  • 153
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    Improved solution for potential flow about arbitrary axisymmetric bodies by the use of a higher-order surface source method (1975)
    In:  Other Sources
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    Publication Date: 2019-06-27
    Description: An investigation is conducted of a case of axisymmetric bodies in which the application of main interest is an inlet, possibly with centerbody and ring vanes. The technique employed makes use of curved surface elements and a source density which varies over the element. Such an approach is designated a higher-order implementation. Questions of surface element geometry are discussed along with the computation of the induced velocity matrices and the organization of the calculation. The calculated results are compared with analytic solutions.
    Keywords: AERODYNAMICS
    Type: Computer Methods in Applied Mechanics and Engineering; 5; May 1975
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  • 154
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    Investigation of rotor blade element airloads for a teetering rotor in the blade stall regime (second wind tunnel test) (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A test was conducted in the NASA-Ames 7 x 10 ft low speed wind tunnel on a seven-foot diameter model of a teetering rotor. The objectives of the test were: (1) acquire pressure data for correlation with laser and flow visualization measurements; (2) explore rotor propulsive force limits by varying the advance ratio at constant lift and propulsive force coefficients; (3) obtain additional data to define the differences between teetering and articulated rotors; and (4) verify the acceleration sensitivity of experimental transducers. Results are presented.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137534 , D210-10792-3
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  • 155
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    SENS-5D trajectory and wind-sensitivity calculations for unguided rockets (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A computational procedure is described which numerically integrates the equations of motion of an unguided rocket. Three translational and two angular (roll discarded) degrees of freedom are integrated through the final burnout; and then, through impact, only three translational motions are considered. Input to the routine is: initial time, altitude and velocity, vehicle characteristics, and other defined options. Input format has a wide range of flexibility for special calculations. Output is geared mainly to the wind-weighting procedure, and includes summary of trajectory at burnout, apogee and impact, summary of spent-stage trajectories, detailed position and vehicle data, unit-wind effects for head, tail and cross winds, coriolis deflections, range derivative, and the sensitivity curves (the so called F(Z) and DF(Z) curves). The numerical integration procedure is a fourth-order, modified Adams-Bashforth Predictor-Corrector method. This method is supplemented by a fourth-order Runge-Kutta method to start the integration at t=0 and whenever error criteria demand a change in step size.
    Keywords: AERODYNAMICS
    Type: NASA-CR-141402
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  • 156
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    Aerodynamic results of a separation test (CA20) conducted at the Boeing transonic wind tunnel using 0.030-scale models of the configuration 140A/B (modified) SSV orbiter (model no. 45-0) and the Boeing 747 carrier (model no. AX 1319 I-1), volume 2 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N76-16033.
    Keywords: AERODYNAMICS
    Type: NASA-CR-141845 , DMS-DR-2217-VOL-2
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  • 157
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    Numerical studies of incompressible viscous flow in a driven cavity (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A series of project papers is presented in computational fluid dynamics. The work was performed during the 1973-74 academic year at Old Dominion University. Each paper briefly examines a numerical method(s) that can be applied to the Navier-Stokes equations governing incompressible flow in a driven cavity. Solutions obtained with a cubic spline procedure are also included.
    Keywords: AERODYNAMICS
    Type: NASA-SP-378
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  • 158
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    Low speed wind tunnel investigation of a four-engine upper surface blown model having swept wing and rectangular and D-shaped exhaust nozzles (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A low speed investigation was conducted in the Langley V/STOL tunnel to determine the power-on static-turning and powered-lift aerodynamic performance of a four engine upper surface blown transport configuration. Initial tests with a D-shaped exhaust nozzle showed relatively poor flow-turning capability, and the D-nozzles were replaced by rectangular nozzles with a width-height ratio of 6.0. The high lift system consisted of a leading edge slat and two different trailing-edge-flap configurations. A double slotted flap with the gaps sealed was investigated and a simple radius flap was also tested. A maximum lift coefficient of approximately 9.3 was obtained for the model with the rectangular exhaust nozzles with both the double slotted flap deflected 50 deg and the radius flap deflected 90 deg.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8061 , L-10173
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  • 159
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    Flight test investigation of the vortex wake characteristics behind a Boeing 727 during two-segment and normal ILS approaches (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A series of flight tests were performed to evaluate the vortex wake characteristics of a Boeing 727 (B727-200) aircraft during conventional and two-segment ILS approaches. Flights of the B727, equipped with smoke generators for vortex marking, were flown wherein its vortex wake was intentionally encountered by a Lear Jet model 23 (LR-23) or a Piper Twin Comanche (Pa-30); and its vortex location during landing approach was measured using a system of photo-theodolites. The tests showed that at a given separation distance there were no differences in the upsets resulting from deliberate vortex encounters during the two types of approaches. Timed mappings of the position of the landing configuration vortices showed that they tended to descend approximately 91 meters (300 feet) below the flight path of the B727. The flaps of the B727 have a dominant effect on the character of the trailed wake vortex. The clean wing produces a strong, concentrated vortex. As the flaps are lowered, the vortex system becomes more diffuse. Pilot opinion and roll acceleration data indicate that 4.5 nautical miles would be a minimum separation distance at which roll control could be maintained during parallel encounters of the B727's landing configuration wake by small aircraft.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72908 , FAA-NA-75-151 , AD-A018366
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  • 160
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    Analysis of measured airfoil pressure distributions (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A method for evaluating the Glauert coefficients from airfoil pressure distributions is investigated. The linear operating range of the airfoils in steady-state and periodic operating conditions are considered. A rational method for quantitatively characterizing airfoil pressure distributions relative to their geometry and aerodynamic operating environment is developed. The characteristics of the airfoil operating environment is determined from its measured pressure distribution.
    Keywords: AERODYNAMICS
    Type: NASA-CR-144901 , VIZEX-CR-75-6
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  • 161
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    Force and pressure measurements on an airfoil oscillating through stall, part 2 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Details of force, moment, and pressure distributions on a two dimensional, four foot chord, NACA 0012 airfoil, oscillating in pitch through stall, in a 7 ft. x 10 ft. low speed wind tunnel, are presented. Tests were run with the airfoil in a closed test section and also in a test section having four longitudinal slots in each sidewall set to provide minimum tunnel interference on the wing in steady flow. In unsteady flow, differences between the results for the closed and 2% open case are small. The dynamic stall process is not triggered by the bursting of a laminar separation bubble but rather by the separation of the turbulent boundary layer downstream of the bubble.
    Keywords: AERODYNAMICS
    Type: NASA-CR-145877 , TEES-3018-75-01A
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  • 162
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    Study to define unsteady flow fields and their statistical characteristics (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Preliminary estimates of space shuttle fluctuating pressure environments were made based on analyses of wind tunnel data, and empirical prediction techniques. Particular emphasis was given to the external tank and solid rocket boosters for the transonic speed regime during launch of a parallel-burn space shuttle configuration. Predicted environments are presented as space-averaged zonal profiles with progressive shading from zone to zone to illustrate spatial variations in the magnitude of the fluctuating pressure coefficient over the surfaces of the external tank and solid rocket boosters. Predictions are provided for the transonic Mach number range from 0.8 equal to or less than M sub infinity equal to or less than 1.5, and for supersonic Mach numbers of 2.0 and 3.0.
    Keywords: AERODYNAMICS
    Type: NASA-CR-144092 , WR-75-1
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  • 163
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    Computed lateral rate and acceleration power spectral response of conventional and STOL airplanes to atmospheric turbulence (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Power-spectral-density calculations were made of the lateral responses to atmospheric turbulence for several conventional and short take-off and landing (STOL) airplanes. The turbulence was modeled as three orthogonal velocity components, which were uncorrelated, and each was represented with a one-dimensional power spectrum. Power spectral densities were computed for displacements, rates, and accelerations in roll, yaw, and sideslip. In addition, the power spectral density of the transverse acceleration was computed. Evaluation of ride quality based on a specific ride quality criterion was also made. The results show that the STOL airplanes generally had larger values for the rate and acceleration power spectra (and, consequently, larger corresponding root-mean-square values) than the conventional airplanes. The ride quality criterion gave poorer ratings to the STOL airplanes than to the conventional airplanes.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8022 , L-10018
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  • 164
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    A preliminary study of the effects of vortex diffusers (winglets) on wing flutter (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Some experimental flutter results are presented for a simple, flat-plate wing model and for the same wing model equipped with two different upper surface vortex diffusers over the Mach number range from about 0.70 to 0.95. Both vortex diffusers had the same planform, but one weighed about 0.3 percent of the basic wing weight, whereas the other weighed about 1.8 percent of the wing weight. The addition of the lighter vortex diffuser reduced the flutter dynamic pressure by about 3 percent; the heavier vortex diffuser reduced the flutter dynamic pressure by about 12 percent. The experimental flutter results are compared at a Mach number of 0.80 with analytical flutter results obtained by using doublet lattice and lifting surface (Kernel function) unsteady aerodynamic theories.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72799
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  • 165
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    Vector thrust induced lift effects for several ejector exhaust locations on a V/STOL wind tunnel model at forward speed (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results and analysis of aerodynamic force data obtained from a small scale model of a V/STOL research vehicle in a low speed wind tunnel are presented. The analysis of the data includes the evaluation of aerodynamic-propulsive lift performance when operating twin ejector nozzles with thrust deflected. Three different types of thrust deflector systems were examined: 90 deg downward deflected nozzle, 90 deg slotted nozzle with boundary layer control, and an externally blown flap configuration. Several nozzle locations were tested, including over and underwing positions. The interference lift of the nacelle and model due to jet exhaust thrust is compared and results show that 90 deg turned nozzles located over the wing (near the trailing edge) produce the largest interference lift increment for an untrimmed aircraft, and that the slotted nozzle located under the wing near the trailing edge (in conjunction with a BLC flap) gives a comparable interference lift in the trimmed condition. The externally blown flap nozzle produced the least interference lift and significantly less total lift due to jet thrust effects.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137733
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  • 166
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    An experimental investigation of the NASA space shuttle external tank at hypersonic Mach numbers (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Pressure and heat transfer tests were conducted simulating flight conditions which the space shuttle external tank will experience prior to break-up. The tests were conducted in the Calspan 48-inch Hypersonic Shock Tunnel and simulated entry conditions for nominal, abort-once-around (AOA), and return to launch site (RTLS) launch occurrences. Surface pressure and heat-transfer-rate distributions were obtained with and without various protuberences (or exterior hardware) on the model at Mach numbers from 15.2 to 17.7 at angles of attack from -15 deg to -180 deg and at several roll angles. The tests were conducted over a Reynolds number range from 1300 to 58,000, based on model length.
    Keywords: AERODYNAMICS
    Type: NASA-CR-144090 , CALSPAN-ZE-5716-A-1
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  • 167
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    Adaptation of the Theodorsen theory to the representation of an airfoil as a combination of a lifting line and a thickness distribution (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The theory provides a direct method for resolving an airfoil into a lifting line and a thickness distribution as well as a means of synthesizing thickness and lift components into a resultant airfoil and computing its aerodynamic characteristics. Specific applications of the technique are discussed.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8117 , L-10476
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  • 168
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    Flutter of laminated plates in supersonic flow (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A solution procedure was developed using linear small deflection theory for the flutter of simply supported laminated plates. For such plates, the bending and extensional governing equations are coupled and have cross-stiffness terms which do not appear in classical plate theory. An extended Galerkin method is used to obtain approximate solutions to the governing equations, and the aerodynamic pressure loading used in the analysis is that given by linear piston theory with flow at arbitrary cross-flow angle. A limited parametric study was conducted for typical laminated composite plates. The calculations show that both the bending-extensional coupling and the cross-stiffness terms have a large destabilizing effect on flutter. Since classical plate theory does not consider bending-extensional coupling and cross stiffness terms, it usually gives inaccurate and nonconservative flutter boundaries for laminated plates.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72800
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  • 169
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    Development and flight tests of vortex-attenuating splines (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The ground tests and full-scale flight tests conducted during development of the vortex-attenuating spline are described. The flight tests were conducted using a vortex generating aircraft with and without splines; a second aircraft was used to probe the vortices generated in both cases. The results showed that splines significantly reduced the vortex effects, but resulted in some noise and climb performance penalties on the generating aircraft.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8083 , L-10442
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  • 170
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    Evaluation of viscous drag reduction schemes for subsonic transports (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results are described of a theoretical study of viscous drag reduction schemes for potential application to the fuselage of a long-haul subsonic transport aircraft. The schemes which were examined included tangential slot injection on the fuselage and various synergetic combinations of tangential slot injection and distributed suction applied to wing and fuselage surfaces. Both passive and mechanical (utilizing turbo-machinery) systems were examined. Overall performance of the selected systems was determined at a fixed subsonic cruise condition corresponding to a flight Mach number of free stream M = 0.8 and an altitude of 11,000 m. The nominal aircraft to which most of the performance data was referenced was a wide-body transport of the Boeing 747 category. Some of the performance results obtained with wing suction are referenced to a Lockheed C-141 Star Lifter wing section. Alternate designs investigated involved combinations of boundary layer suction on the wing surfaces and injection on the fuselage, and suction and injection combinations applied to the fuselage only.
    Keywords: AERODYNAMICS
    Type: NASA-CR-132718 , ATL-TR-216
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  • 171
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    Pressure Distribution at Subsonic Speeds over the Forepart of Two Blunt Circular Cylinders (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A wind tunnel investigation was made at subsonic speeds to determine the pressure distribution over the forward part of a circular cylinder. The cylinder was equipped with interchangeable faces, one having a flat face and one having a dome shaped face. The investigation was made over angle of attack range from -1 deg to 26 deg and a Mach number range from 0.30 to 0.89. Pressure coefficients are presented in tabular form and plotted data are presented for some selected angles of attack about the surface of the cylinder.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72784 , PAPER-645
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  • 172
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    Normal- and oblique-shock flow parameters in equilibrium air including attached-shock solutions for surfaces at angles of attack, sweep, and dihedral (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Normal- and oblique-shock flow parameters for air in thermochemical equilibrium are tabulated as a function of shock angle for altitudes ranging from 15.24 km to 91.44 km in increments of 7.62 km at selected hypersonic speeds. Post-shock parameters tabulated include flow-deflection angle, velocity, Mach number, compressibility factor, isentropic exponent, viscosity, Reynolds number, entropy difference, and static pressure, temperature, density, and enthalpy ratios across the shock. A procedure is presented for obtaining oblique-shock flow properties in equilibrium air on surfaces at various angles of attack, sweep, and dihedral by use of the two-dimensional tabulations. Plots of the flow parameters against flow-deflection angle are presented at altitudes of 30.48, 60.96, and 91.44 km for various stream velocities.
    Keywords: AERODYNAMICS
    Type: NASA-SP-3093
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  • 173
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    Effects of upper-surface blowing and thrust vectoring on low-speed aerodynamic characteristics of a large-scale supersonic transport model (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Tests were conducted in the Langley full-scale tunnel to determine the low-speed aerodynamic characteristics of a large-scale arrow-wing supersonic transport configured with engines mounted above the wing for upper surface blowing, and conventional lower surface engines with provisions for thrust vectoring. A limited number of tests were conducted for the upper surface engine configuration in the high lift condition for beta = 10 in order to evaluate lateral directional characteristics, and with the right engine inoperative to evaluate the engine out condition.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72792
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  • 174
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    Inverse solutions for laminar boundary-layer flows with separation and reattachment (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Numerical solutions of the laminar, incompressible boundary layer equations are presented for flows involving separation and reattachment. Regular solutions are obtained with an inverse approach in which either the displacement thickness or the skin friction is specified; the pressure is deduced from the solution. A vorticity-stream-function formulation of the boundary layer equations is used to eliminate the unknown pressure. Solutions of the resulting finite difference equations, in which the flow direction is taken into account, are obtained by several global iteration schemes which are stable and have unconditional diagonal dominance. Results are compared with Klineberg and Steger's separated boundary layer calculations, and with Briley's solution of Navier-Stokes equations for a separated region. In addition, an approximate technique is presented in which the streamwise convection of vorticity is set equal to zero in the reversed flow region; such a technique results in a quick forward marching procedure for separated flows.
    Keywords: AERODYNAMICS
    Type: NASA-TR-R-447 , L-10336
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  • 175
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    Second-order small-disturbance solutions for hypersonic flow over power-law bodies (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Similarity solutions were found which give the adiabatic flow of an ideal gas about two-dimensional and axisymmetric power-law bodies at infinite Mach number to second order in the body slenderness parameter. The flow variables were expressed as a sum of zero-order and perturbation similarity functions for which the axial variations in the flow equations separated out. The resulting similarity equations were integrated numerically. The solutions, which are universal functions, are presented in graphic and tabular form. To avoid a singularity in the calculations, the results are limited to body power-law exponents greater than about 0.85 for the two-dimensional case and 0.75 for the axisymmetric case. Because of the entropy layer induced by the nose bluntness (for power-law bodies other than cones and wedges), only the pressure function is valid at the body surface. The similarity results give excellent agreement with the exact solutions for inviscid flow over wedges and cones having half-angles up to about 20 deg. They give good agreement with experimental shock-wave shapes and surface-pressure distributions for 3/4-power axisymmetric bodies, considering that Mach number and boundary-layer displacement effects are not included in the theory.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-7973 , L-10154
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  • 176
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    Upwash angles near engine inlets of an externally blown flap STOL transport (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: An investigation was conducted in the Langley V/STOL tunnel to determine the upwash flow angles in the region of the nacelle inlets of a representative powered-lift transport configuration operating at high lift coefficients. The upwash angles were indicated by tufts and measured from photographs. A potential-flow program was used to estimate these flow angles. Large upflow angles exist near the inlets of the nacelles; the highest value (67.3 deg) occurred with flaps at 15, 35, 55 deg, an angle of attack of 25.7 deg, and a thrust coefficient of 4. The upflow angle was found to be strongly dependent on the circulation lift, regardless of the flap deflection, angle of attack, or thrust coefficient used to generate this circulation lift. The potential-flow calculations away from the nacelle inlets agreed fairly well with the experimental data.
    Keywords: AERODYNAMICS
    Type: NASA-TN-D-8091 , L-10406
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  • 177
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    Steady subsonic flow around finite-thickness wings (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The general method for analyzing steady subsonic potential aerodynamic flow around a lifting body having arbitrary shape is presented. By using the Green function method, an integral representation for the potential is obtained. Under small perturbation assumption, the potential at any point, P, in the field depends only upon the values of the potential and its normal derivative on the surface of the body. Hence if the point P approaches the surface of the body, the representation reduces to an integral equation relating the potential and its normal derivative (which is known from the boundary conditions) on the surface. The question of uniqueness is examined and it is shown that, for thin wings, the operator becomes singular as the thickness approaches zero. This fact may yield numerical problems for very thin wings. However, numerical results obtained for a rectangular wing in subsonic flow show that these problems do not appear even for thickness ratio tau = .001. Comparison with existing results shows that the proposed method is at least as fast and accurate as the lifting surface theories.
    Keywords: AERODYNAMICS
    Type: NASA-CR-2616 , TR-73-02
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  • 178
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    A subvortex technique for the close approach to a discretized vortex sheet (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The close-approach problem associated with flow calculation methods based on vortex-lattice theory was examined numerically using two-dimensional discretized vortex sheets. The analysis first yields a near-field radius of approximately the distance apart of the vortices in the lattice; only within this distance from the sheet are the errors arising from the discretization significant. Various modifications to the discrete vortices are then considered with the objective of reducing the errors. This leads to a near-field model in which a vortex splits into an increasing number of subvortices as it is approached. The subvortices, whose strengths vary linearly from the vortex position, are evenly distributed along an interpolated curve passing through the basic vortices. This subvortex technique can be extended to the three-dimensional case and is efficient because the number of vortices is effectively increased, but only where and when needed.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62487 , A-6277
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  • 179
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    A design method for entrance sections of transonic wind tunnels with rectangular cross sections (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A mathematical technique developed to design entrance sections for transonic or high-speed subsonic wind tunnels with rectangular cross sections is discribed. The transition from a circular cross-section setting chamber to a rectangular test section is accomplished smoothly so as not to introduce secondary flows (vortices or boundary-layer separation) into a uniform test stream. The results of static-pressure measurements in the transition region and of static and total-pressure surveys in the test section of a pilot model for a new facility at the Ames Research Center are presented.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62490 , A-6293
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  • 180
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    Introduction to the aerodynamics of flight (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: General concepts of the aerodynamics of flight are discussed. Topics considered include: the atmosphere; fluid flow; subsonic flow effects; transonic flow; supersonic flow; aircraft performance; and stability and control.
    Keywords: AERODYNAMICS
    Type: NASA-SP-367
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  • 181
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    Low speed aerodynamic characteristics of a transport model having 42.33 deg swept low wing with supercritical airfoil, double-slotted flaps, and T-tail or low tail (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A low-speed investigation was conducted in the Langley V/STOL tunnel over an angle-of-attack range of approximately 4 deg to 24 deg to determine the static longitudinal stability characteristics and high lift performance of a general research model which represented an advanced subsonic transport configuration. The model had a 42.33 deg swept, aspect ratio 7.05 wing with a supercritical airfoil and high lift system consisting of a leading edge device (slat or Kruger flap) and a double-slotted flap. The flaps were deflected for take off and landing configurations and were not deflected for tests of the clean configuration. The model was tested with the horizontal tail in either a T tail or low tail position. The effects of various arrangements of flowthrough nacelles which represent a three engine configuration (two large wing-mounted nacelles and a vertical tail mounted nacelle) and a four engine configuration (four smaller wing-mounted nacelles) were determined.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-3276 , L-9957
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  • 182
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    A technique for accelerating iterative convergence in numerical integration, with application in transonic aerodynamics (1975)
    In:  CASI
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    Publication Date: 2019-06-27
    Description: A technique is described for the efficient numerical solution of nonlinear partial differential equations by rapid iteration. In particular, a special approach is described for applying the Aitken acceleration formula (a simple Pade approximant) for accelerating the iterative convergence. The method finds the most appropriate successive approximations, which are in a most nearly geometric sequence, for use in the Aitken formula. Simple examples are given to illustrate the use of the method. The method is then applied to the mixed elliptic-hyperbolic problem of steady, inviscid, transonic flow over an airfoil in a subsonic free stream.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62496
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  • 183
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    An investigation of the increase in vortex induced rolling moment associated with landing gear wake (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Flight tests were conducted to verify the results found in ground base facilities of the effect of span lift load variation as well as the vortex attentuation of the high energy jet engine exhaust through proper thrust programming. During these flight tests a large increase in vortex strength was experienced as a result of extending the landing gear. Tests in the Langley Vortex Research Facility indicate that the wake produced by the landing gear may possibly form an aerodynamic endplate or reflection plane at the inboard edge of each inboard flap which increases the effective aspect ratio of the flap and thereby increases the strength of the flap outer edge vortex.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72786
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  • 184
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    Transonic pressure measurements and comparison of theory to experiment for an arrow-wing configuration. Volume 3: Data report, comparison of attached flow theories to experiment (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N76-11034.
    Keywords: AERODYNAMICS
    Type: NASA-CR-132729 , D6-42670-4-VOL-3
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  • 185
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    Transonic pressure measurements and comparison of theory to experiment for an arrow-wing configuration. Volume 2: Experimental data report, effects of control surface deflection (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: For abstract, see N76-11034.
    Keywords: AERODYNAMICS
    Type: NASA-CR-132728 , D6-42670-3-VOL-2
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  • 186
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    Development of a computer program to obtain ordinates for NACA 4-digit, 4-digit modified, 5-digit, and 16 series airfoils (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A computer program developed to calculate the ordinates and surface slopes of any thickness, symmetrical or cambered NACA airfoil of the 4-digit, 4-digit modified, 5-digit, and 16-series airfoil families is presented. The program produces plots of the airfoil nondimensional ordinates and a punch card output of ordinates in the input format of a readily available program for determining the pressure distributions of arbitrary airfoils in subsonic potential viscous flow.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-3284 , L-10375
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  • 187
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    Numerical methods for the calculation of three-dimensional nozzle exhaust flow fields (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: Numerical codes developed for the calculation of three-dimensional nozzle exhaust flow fields associated with hypersonic airbreathing aircraft are described. Both codes employ reference plane grid networks with respect to three coordinate systems. Program CHAR3D is a characteristic code utilizing a new wave preserving network within the reference planes, while program BIGMAC is a finite difference code utilizing conservation variables and a one-sided difference algorithm. Secondary waves are numerically captured by both codes, while the underexpansion shock and plume boundary are treated discretely. The exhaust gas properties consist of hydrogen-air combustion product mixtures in local chemical equilibrium. Nozzle contours are treated by a newly developed geometry package based on dual cubic splines. Results are presented for simple configurations demonstrating two- and three-dimensional multiple wave interactions.
    Keywords: AERODYNAMICS
    Type: NASA. Langley Res. Center Aerodynamic Analyses Requiring Advanced Computers, Pt. 1; p 659-701
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  • 188
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    Theoretical performance of cross-wind axis turbines with results for a catenary vertical axis configuration (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A general analysis capable of predicting performance characteristics of cross-wind axis turbines was developed, including the effects of airfoil geometry, support struts, blade aspect ratio, windmill solidity, blade interference and curved flow. The results were compared with available wind tunnel results for a catenary blade shape. A theoretical performance curve for an aerodynamically efficient straight blade configuration was also presented. In addition, a linearized analytical solution applicable for straight configurations was developed. A listing of the computer program developed for numerical solutions of the general performance equations is included in the appendix.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-72662
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  • 189
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    Flow visualization of vortex interactions in multiple vortex wakes behind aircraft (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A flow visualization technique was developed which allows the nature of lift-generated wakes behind aircraft models to be investigated. The technique was applied to models being towed underwater in a ship model basin. Seven different configurations of a small-scale model of a 747 transport aircraft were used to allow observation of typical vortex interactions and merging in multiple vortex wakes. It was established that the motion of the wake vortices is often sensitive to small changes in either wing span loading or model attitude. Landing gear deployement was found to cause a far-field reformation of vorticity behind a model configuration which dissipated concentrated vorticity in the near-field wake. Alleviation of wake vorticity is achievable by configuring the wing span loading to cause the wake vortices to move in paths that result in their interactions and merging. The vortices shed from the horizontal stabilizer always moved down rapidly into the wake and merged with the other vortices, primarily the inboard flap vortices.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-62459 , A-6193
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  • 190
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    An exploratory investigation of the cooling drag associated with general aviation propulsive systems (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A systematic investigation of the drag associated with cooling air flow in contemporary general aviation engine installations is proposed. Theoretical and experimental methods include a state-of-the-art survey, determination of cooling drag by flight tests, and establishment of relative magnitude and components of cooling drag.
    Keywords: AERODYNAMICS
    Type: Kansas Univ. Proc. of the NASA, Ind., Univ., Gen. Aviation Drag Reduction Workshop; p 263-272
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  • 191
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    Mean and fluctuating flow measurements of a fully-developed, non-adiabatic, hypersonic boundary layer (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: Extensive boundary-layer measurements have been made on a cone-ogive-cylinder model at a free-stream Mach number of 7.0 and momentum-thickness Reynolds number of 8500. Mean flow transformations and calculated turbulence correlations are presented which are in good agreement with previous incompressible results. New quantitative turbulence measurements including measurements of the first higher moment and probability density of fluctuations in mass flow and total temperature in hypersonic flow are also presented. The higher moment and probability density data show that the characters of the fluctuation modes of the mass flow and total temperature are significantly different in the wall region and in the outer part of the boundary layer. These differences together with data on the turbulence scale and lifetime obtained from autocorrelation and space-time correlation measurements are discussed.
    Keywords: AERODYNAMICS
    Type: Journal of Fluid Mechanics; 70; July 29
    Format: text
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  • 192
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    Investigation of two-dimensional wedge exhaust nozzles for advanced aircraft (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: Two-dimensional wedge nozzle performance characteristics were investigated in a series of wind-tunnel tests. An isolated single-engine/nozzle model was used to study the effects of internal expansion area ratio, aftbody cowl boattail angle, and wedge length. An integrated twin-engine/nozzle model, tested with and without empenage surfaces, included cruise, acceleration, thrust vectoring and thrust reversing nozzle operating modes. Results indicate that the thrust-minus-aftbody drag performance of the twin two-dimensional nozzle integration is significantly higher, for speeds greater than Mach 0.8, than the performance achieved with twin axisymmetric nozzle installations. Significant jet-induced lift was obtained on an aft-mounted lifting surface using a cambered wedge center body to vector thrust. The thrust reversing capabilities of reverser panels installed on the two-dimensional wedge center body were very effective for static or in-flight operation.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 75-1317
    Format: text
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  • 193
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    The effect of Reynolds number on the boattail drag of two wing-body configurations (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: An investigation has been conducted in the Langley 1/3-meter transonic cryogenic tunnel to determine the effects of varying Reynolds number on the boattail drag of wing-body configurations at subsonic speeds. Two boattailed cone-cylinder nacelle models were tested with a 60 deg delta wing at an angle of attack of 0 deg. Reynolds number, based on model length, was varied from about 2.5 million to 67 million. Even though the presence of the wing had large effects on the boattail pressure coefficients, the results of this investigation were similar to those previously found for a series of isolated boattails. Boattail pressure coefficients in the expansion region became more negative with increasing Reynolds number, while those in the recompression region became more positive. These two effects were compensating, and as a result, there was virtually no effect of Reynolds number on boattail pressure drag.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 75-1294
    Format: text
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  • 194
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    Inlet performance of the integrated Langley scramjet module (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: The inlet concept for the Langley Scramjet Module has been developed and proven in Langley wind tunnels over a Mach number range from 2.3 to 6.0 (flight simulation of Mach 2.6 to 7.6). This modular engine concept is designed to integrate with the airframe, which results in precompression of the engine airflow by the vehicle bow shock and additional expansion of the nozzle exhaust gas by the afterbody of the vehicle. With these integration advantages, the inlet can be designed with modest contraction ratios and fixed geometry. Also, the module nozzle exit area can be equal to the capture area, which permits the cowl to be alined with the local flow producing minimum external drag. The inlet leading edges and planar compression surfaces are swept at 48 deg, which provides spillage at low Mach numbers for starting and which reduces the pressure gradient on the top surface to permit ingestion of the vehicle forebody boundary layer into the inlet without separating. Three fuel injection struts provide for the use of a short combustor having low internal cooling requirements. Schedules for mass capture ratio, contraction ratio, and total pressure recovery are well within the acceptable range for a good scramjet propulsion device. The fixed geometry, minimum external drag design has proven to be a practical, high-performance inlet concept.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 75-1212
    Format: text
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  • 195
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    An experimental and analytical investigation of axisymmetric diffusers (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-27
    Description: A finite difference computer program for turbulent compressible flow was used to establish the performance of several diffuser shapes for experimental testing. The diffusers were designed to have a linear change in Mach number, a linear change in pressure or a curvature fitted by a quadratic equation. Testing was performed with M = 0.1 to 0.9 with and without boundary layer bleed. Above M = 0.6, data were obtained with a normal shock upstream of the diffuser entrance. Peak static pressure recovery occurred with a diffuser inlet M-0.75. The quadratic diffuser yielded the highest total pressure recovery.
    Keywords: AERODYNAMICS
    Type: AIAA PAPER 75-1211
    Format: text
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  • 196
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    Iterative techniques for the solution of large linear systems in computational aerodynamics (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Keywords: AERODYNAMICS
    Type: Journal of Aircraft; 12; Feb. 197
    Format: text
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  • 197
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    Model for the solution of supersonic viscous interaction effects on blunted cones (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 13; Mar. 197
    Format: text
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  • 198
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    Strong interaction associated with transonic flow past boattailed afterbodies (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-06-27
    Description: The problem of transonic flow past boattails was studied with the aid of numerical relaxative schemes. Preliminary calculations were restricted to a particular model configuration which had been tested in an experimental program. It was found that the full potential equation must be considered in the study. The final results agreed very well with the experimental data. The investigation illustrates the strong interaction character of the transonic flow past a boattailed afterbody.
    Keywords: AERODYNAMICS
    Type: AIAA Journal; 13; Jan. 197
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  • 199
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    Description of CASCOMP Comprehensive Airship Sizing and Performance Computer Program, Volume 2 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The computer program CASCOMP, which may be used in comparative design studies of lighter than air vehicles by rapidly providing airship size and mission performance data, was prepared and documented. The program can be used to define design requirements such as weight breakdown, required propulsive power, and physical dimensions of airships which are designed to meet specified mission requirements. The program is also useful in sensitivity studies involving both design trade-offs and performance trade-offs. The input to the program primarily consists of a series of single point values such as hull overall fineness ratio, number of engines, airship hull and empennage drag coefficients, description of the mission profile, and weights of fixed equipment, fixed useful load and payload. In order to minimize computation time, the program makes ample use of optional computation paths.
    Keywords: AERODYNAMICS
    Type: NASA-CR-137691-VOL-2 , D210-10953-2-VOL-2
    Format: application/pdf
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  • 200
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    Effect of wall suction on performance of a short annular diffuser at inlet Mach numbers up to 0.5 (1975)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: A short annular diffuser equipped with wall bleed (suction)capability was evaluated at inlet Mach numbers of 0.186 to 0.5. The diffuser had an area ratio of 4.0 and a length-to-inlet height ratio of 1.6. Test results show that the exit velocity profiles, typical of annular jet flow without suction, could be considerably flattened by application of wall suction. This improved performance was also reflected in diffuser effectiveness (static-pressure recovery) and total-pressure loss results. At the inlet Mach number of 0.5 diffuser static-pressure recovery is equal to or better than at lower inlet Mach numbers for comparable suction rates.
    Keywords: AERODYNAMICS
    Type: NASA-TM-X-3302 , E-8393
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