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At least 757 records · Page 42

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NASA NTRS1981

Effects of blade-vane ratio and rotor-stator spacing of fan noise with forward velocity

A research fan stage was acoustically tested in an anechoic wind tunnel with a 41 m/sec tunnel flow. Two stator vane numbers giving cut-on and cut-off conditions were tested at three rotor-stator spacings ranging from about 0.5 to 2.0 rotor chords. These two stators were designed for similar aerodynamic performance. Hot film anemometer turbulence measurements were made at the leading edge of the stator for each spacing. The cut-off criterion strongly controlled the fundamental tone level at all spacings. The trends with spacing of the wake defect upwash component at the stator tip showed good agreement with the corresponding cut-on acoustic tone levels.

Woodward, R. P.↗

NASA NTRS1981

A full-scale wind tunnel investigation of a helicopter bearingless main rotor

A helicopter bearingless main rotor was tested. Areas of investigation included aeroelastic stability, aerodynamic performance, and rotor loads as a function of collective pitch setting, RPM, airspeed and shaft angle. The rotor/support system was tested with the wind tunnel balance dampers installed and, subsequently, removed. Modifications to the rotor hub were tested. These included a reduction in the rotor control system stiffness and increased flexbeam structural damping. The primary objective of the test was to determine aeroelastic stability of the fundamental flexbeam/blade chordwise bending mode. The rotor was stable for all conditions. Damping of the rotor chordwise bending mode increases with increased collective pitch angle at constant operating conditions. No significant decrease in rotor damping occured due to frequency coalescence between the blade chordwise fundamental bending mode and the support system.

Warmbrodt, W.↗

NASA NTRS1981

An analytical approach to airfoil icing

An analytical procedure has been developed to predict rime ice growth on unprotected airfoil sections and to evaluate the aerodynamic performance. A time stepping method is used in which: (1) water droplet trajectories are calculated, (2) a rime ice shape determined, (3) the flowfield around the iced airfoil is recalculated, and (4) the build-up process iterated upon until the desired icing time is reached. The performance of the iced airfoil shapes are then determined from existing analytic methods. Rime ice shapes determined in the NASA Lewis Icing Research Tunnel on a modified NACA 64 series airfoil agree well with the shapes predicted by the analytical method. Measured and predicted increases in drag due to the rime ice also agree favorably. A simplified scaling analysis is also presented and verified which provides the duplication of full scale results of rime ice accretions in small scale model tests.

Bragg, M. B.↗

NASA NTRS1981

Design and development of the combustor inlet diffuser for the NASA/GE energy efficient engine

Results of an experimental investigation of the aerodynamic performance of a split duct annular combustor inlet diffuser system are presented. Several diffuser configurations were investigated in 3X-scale water table tests and the preferred design was evaluated in full-scale annular airflow model tests. Pressure recovery and flow losses were determined as a function of prediffuser inlet velocity profile, flow extraction at the prediffuser exit, and distribution of flow in the combustor. Inlet velocity profile and turbulence levels were found to have a pronounced effect on system performance. Flow extraction at the prediffuser exit was found to have little influence on system performance. Generally, the annular split duct diffuser system was found to satisfy the performance objectives for the engine.

Sabla, P. E.↗

NASA NTRS1983

Supersonic-Wing Nonlinear Aerodynamics

Supersonic Wing Nonlinear Aerodynamics computer program, LTSTAR, estimates nonlinear aerodynamics characteristics of wing at supersonic speeds. Program provides generally improved predictions of wing overall force and moment coefficients. LTSTAR is useful in design studies aimed at aerodynamic performance optimization and providing more-realistic tradeoff information for selection of wing planform geometry and airfoil section parameters.

Carlson, H. W.↗

NASA NTRS1981

Low speed testing of the inlets designed for a tamden-fan V/STOL nacelle

An approximately 0.25 scale model of a tandem fan nacelle, designed for a subsonic V/STOL aircraft, was tested in a Lewis wind tunnel. Model variables included long and short aft inlet cowls and the addition of exterior strakes to the short inlet cowl. Inlet pressure recoveries and distortion were measured at pitch angles to 40 deg and at combinations of pitch and yaw to 30 deg. Airspeeds covered a range to 135 knots (69 m/sec). The short aft inlet with added strakes had the best aerodynamic performance and is considered suitable for the intended V/STOL application.

Williams, R. C.↗

NASA NTRS1981

Airframe-propulsion system aerodynamic interference predictions at high transonic Mach numbers including off-design engine airflow effects

The transonic speed regime for airplanes at conditions where inlet spillage takes place is discussed. A wind tunnel test program to evaluate aerodynamic performance penalties associated with propulsion system installation and operation at subsonic through low supersonic speeds was conducted. The accuracy of analytic methods for predicting transonic engine airframe interference effects was assessed. Study variables included Mach number, angle of attack, relative nacelle location, and nacelle mass flow ratio. Results include test theory comparisons of forces as well as induced pressure fields. Prediction capability of induced shock wave strength and locations is assessed. It was found that large interference forces due to engine location and flow spillage occur at transonic speeds, that theory explains these effects; and that theory can predict quantitatively these effects.

Kulfan, R. M.↗

NASA NTRS1982

Jet V/STOL wind-tunnel simulation and groundplane effects

The pretest preparation necessary to define the objectives of an appropriate investigation into the jet V/STOL wind tunnel simulation and ground plane effects were examined. Low speed wind tunnel testing of V/STOL aircraft concepts to determine the aerodynamic propulsion interaction effects during the transition between hover and wingborne flight is a necessary step in the development cycle of this type of aircraft. Powered models are normally used to determine the aerodynamic performance characteristics. Several factors which influence the selection of the model concept and the engine simulator are discussed. Some of the test techniques important for this class of aircraft model are examined. Wind tunnel wall effects important to this aircraft testing with special emphasis on groundplane effects are reviewed.

Margason, R. J.↗

NASA NTRS1982

A FORTRAN program for calculating three dimensional, inviscid and rotational flows with shock waves in axial compressor blade rows: User's manual

A FORTRAN-IV computer program was developed for the calculation of the inviscid transonic/supersonic flow field in a fully three dimensional blade passage of an axial compressor rotor or stator. Rotors may have dampers (part span shrouds). MacCormack's explicit time marching method is used to solve the unsteady Euler equations on a finite difference mesh. This technique captures shocks and smears them over several grid points. Input quantities are blade row geometry, operating conditions and thermodynamic quanities. Output quantities are three velocity components, density and internal energy at each mesh point. Other flow quanities are calculated from these variables. A short graphics package is included with the code, and may be used to display the finite difference grid, blade geometry and static pressure contour plots on blade to blade calculation surfaces or blade suction and pressure surfaces. The flow in a low aspect ratio transonic compressor was analyzed and compared with high response total pressure probe measurements and gas fluorescence static density measurements made in the MIT blowdown wind tunnel. These comparisons show that the computed flow fields accurately model the measured shock wave locations and overall aerodynamic performance.

Thompkins, W. T., Jr.↗

NASA NTRS1982

Viscous effect on airfoils for unsteady transonic flows

The viscous effect on aerodynamic performance of an arbitrary airfoil executing low frequency maneuvers during transonic flight was investigated. The small disturbance code, LTRAN2, was modified by using a conventional integral method, BLAYER, for the boundary layer and an empirical relation, viscous wedge, for simulating the suddenly thickened boundary layer behind the shock. Before the shock, only the boundary layer displacement thickness was evaluated. After the shock, the empirical wedge thickness was superimposed on the boundary layer thickness along the surface as well as in the wake region. The pressure coefficients were calculated for both steady and unsteady states. The viscous solution takes fewer iterations to obtain the converged steady state solution. Comparisons made with experimental data and the inviscid solution show that the viscous solution agrees better with the experimental data with about the same (or slightly less) amount of computational time.

Lee, S. C.↗

NASA NTRS1982

Small axial turbine stator technology program

An experimental investigation was conducted to determine the effects of surface finish, fillet radius, inlet boundary layer thickness, and free-stream inlet turbulence level on the aerodynamic performance of a small axial flow turbine stator. The principal objective of this program was to help understand why large turbine efficiency is not maintained when a large turbine is scaled to a smaller size. The stator used in this program as a one-sixth scale of a 762 mm (30 in.) diameter stator design with 50 vanes having a vane height of 17 mm (0.666 in.) and an aspect ratio of 1.77. A comprehensive overall test matrix was used to provide a complete engineering understanding of the effects of each variable over the full range of all the other variables. The range of each variable investigated was as follows: surface finish 0.1 micro (4 micro in.) to 2.4 micro (95 micro in.); boundary layer thickness 2 to 25 percent of channel height at each wall; fillet radius 0 mm (0 in.) to 1.0 mm (.040 in.) and turbulence 2 to 12 percent.

Brockett, W.↗

NASA NTRS1981

Jet V/STOL wind-tunnel simulation and groundplane effects

Low-speed wind-tunnel testing of V/STOL aircraft concepts to determine the aerodynamic-propulsion interaction effects during the transition between hover and wingborne flight is a necessary step in the development cycle of this type of aircraft. Powered models are normally used to determine the aerodynamic performance characteristics. This paper examines some of the pretest preparation necessary to define the objectives of an appropriate investigation. Several factors which influence the selection of the model concept and the engine simulator are discussed. In addition, some of the test techniques important for this class of aircraft model are examined. Finally, the paper reviews some of the wind-tunnel wall effects important to this type of aircraft testing with special emphasis on groundplane effects.

Margason, R. J.↗

NASA NTRS1981

A summary of V/STOL inlet analysis methods

Methods to analyze the aerodynamic performance of V/STOL inlets at the NASA Lewis Research Center are discussed with emphasis on recent extensions and applications. Methods include the specification of the Kutta condition for a slotted inlet, the calculation of suction and tangential blowing for boundary layer control, and the analysis of auxiliary inlet geometries at angles of attack. An optimum diffuser velocity distribution is also developed, which can result in the shortest no-boundary layer control inlet and the lowest loss for the required amount of diffusion.

Hwang, D. P.↗

NASA NTRS1982

Vortex lift augmentation by suction on a 60 deg swept Gothic wing

An experimental investigation was conducted in the Langley high-speed 7- by 10-foot wind tunnel to determine the aerodynamic performance of suction applied near the wing tips above the trailing edge of a 60 deg swept Gothic wing. Moveable suction inlets were symmetrically mounted in the proximity of the trailing edge, and the amount of suction was varied to maximize wing lift. Tests were conducted at Mach 0.15, 0.30, and 0.45, and the angle of attack was varied from -4 to 50 deg. The suction augmentation increases the lift coefficient over the entire range of angle of attack. The lift improvement exceeds the unaugmented wing lift by over 20%. Moreover, the augmented lift exceeds the lift predicted by vortex lattice theory to 30 deg angle of attack. Suction augmentation is postulated to strengthen the vortex system by increasing its velocity and making it more concentrated. This causes the vortex breakdown to be delayed to a higher angle of attack

Taylor, A. H.↗

NASA NTRS1981

Low speed testing of the inlets designed for a tandem-fan V/STOL nacelle

Williams, R. C.↗

NASA NTRS1982

Numerical design of the contoured wind-tunnel liner for the NASA swept-wing LFC test

A contoured, nonporous, wind-tunnel liner has been designed in order to simulate a free-flight infinite yawed-wing, transonic-flow condition about a large-chord, supercritical-section, laminar-flow-control (LFC), swept-wing test panel. The numerical procedure developed for this aerodynamic liner design is based upon the simple idea of streamlining and incorporates several existing transonic and boundary-layer analysis codes. A summary of the entire procedure is presented to indicate: what was done and why, the sequence of steps, and the overall data flow. The liner is being installed in the NASA Langley 8-Foot Transonic Pressure Tunnel (TPT). Test results indicating the aerodynamic performance of the liner are not yet available; thus, the liner design results given here are examples of the calculated requirements and the hardware implementation.

Newman, P. A.↗

NASA NTRS1982

A summary of V/STOL inlet analysis methods

The methods used to analyze the aerodynamic performance of V/STOL inlets at the NASA Lewis Research Center are briefly described. Recent extensions and applications of the method are emphasized. They include the specification of the Kutta condition for a slotted inlet, the calculation of suction and tangential blowing for boundary layer control, and the analysis of auxiliary inlet geometries. A comparison is made with experiment for the slotted inlet and also for tangential blowing. Finally, an optimum inlet diffuser velocity distribution is developed.

Hwang, D. P.↗

NASA NTRS1983

Aeroacoutic characteristics of a large, variable-pitch, variable-speed fan system

The acoustic and aerodynamic performance of the new drive fans for the NASA Ames 40- by 80-/80- by 120-foot wind tunnel was investigated. Results show that a fan system with variable-speed and variable-pitch rotor blades allows the operator to control noise and energy consumption, at a given mass flow rate, through the choice of blade speed and pitch. A low speed and high blade pitch will generally create the least noise at the least energy cost. An empirical method is described which predicts the sound power of this fan system reasonably well.

Soderman, P. T.↗