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At least 1,099 records · Page 61

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NASA NTRS1992

Extra high speed modified Lundell alternator parameters and open/short-circuit characteristics from global 3D-FE magnetic field solutions

The combined magnetic vector potential - magnetic scalar potential method of computation of 3D magnetic fields by finite elements, introduced in a companion paper, is used for global 3D field analysis and machine performance computations under open-circuit and short-circuit conditions for an example 14.3 kVA modified Lundell alternator, whose magnetic field is of intrinsic 3D nature. The computed voltages and currents under these machine test conditions were verified and found to be in very good agreement with corresponding test data. Results of use of this modelling and computation method in the study of a design alteration example, in which the stator stack length of the example alternator is stretched in order to increase voltage and volt-ampere rating, are given here. These results demonstrate the inadequacy of conventional 2D-based design concepts and the imperative of use of this type of 3D magnetic field modelling in the design and investigation of such machines.

Wang, R.↗

NASA NTRS1991

Actively controlled superconducting bearings

Actively controlled conventional radial beating using copper winding and soft magnetic material can provide only up to 200 N/sq cm of pressure. Large cryogenic pumps for space applications operating at 30,000 rpm and high rpm machines may require larger magnetic pressure. We show that using superconducting winding in the rotor and the stator of a magnetic bearing system increases the pressure by an order of magnitude. The paper addresses winding configuration, stability, ac losses, and power requirement for the superconducting winding.

Eyssa, Yehia M.↗

NASA NTRS1993

Unsteady aerodynamic flow phenomena in a transonic compressor stage

A three-dimensional unsteady, viscous aerodynamic analysis has been developed for the flow inside a transonic, high-through-flow, single stage compressor. The compressor stage is comprised of a low-aspect-ratio rotor and a closely coupled stator. The analysis is based on a numerical method for solving the three-dimensional Navier-Stokes equation for unsteady viscous flow through multiple turbomachinery blade rows. The method solves the fully three-dimensional Navier-Stokes equation with an implicit scheme. A two-equation turbulence model with a low-Reynolds-number modification is applied for the turbulence closure. A third-order accurate upwinding scheme is used to approximate convection terms while a second-order accurate central difference scheme is used for the discretization of the viscous terms. A second-order accurate scheme is employed for the temporal discretization. The numerical method is applied to study the unsteady flow field inside a transonic, high-through-flow, axial compressor stage. The numerical results are compared with available experimental data.

Hah, C.↗

NASA NTRS1993

Design and test of a small two stage counter-rotating turbine for rocket engine application

The aerodynamic design and rig test evaluation of a small counter-rotating turbine system is described. The technology represented by this turbine is being developed for application in an advanced upper stage rocket engine turbopump. This engine will employ an oxygen/hydrogen expander cycle and achieve high performance through efficient combustion, high combustion pressure, and high area ratio exhaust nozzle expansion. Engine performance goals require that the turbopump drive turbines achieve high efficiency at low gas flow rates. The low flow rates result in very small airfoil diameter, height and chord. The high efficiency and small size requirements present a challenging turbine design problem. The unconventional approach employed to meet this challenge is described, along with the detailed design process and resulting airfoil configurations. The method and results of full scale aerodynamic performance evaluation testing of both one and two stage configurations, as well as operation without the secondary stage stator are presented. The overall results of this effort illustrate that advanced aerodynamic design tools and hardware fabrication techniques have provided improved capability to produce small high performance turbines for advanced rocket engines.

Huber, F. W.↗

NASA NTRS1993

Three-dimensional flow analysis inside turbomachinery stages with steady and unsteady Navier-Stokes method

This study presents a numerical method for solving the three-dimensional, Navier-Stokes equations for unsteady, viscous flow through multiple turbomachinery blade rows. The method solves the fully three-dimensional Navier-Stokes equations with an implicit scheme which is based on a control volume approach. A two-equation turbulence model with a low Reynolds number modification is employed in the present study. A third-order accurate upwinding scheme is used to approximate convection terms while a second order accurate central difference scheme is used for the discretization of viscous terms. A second-order accurate scheme is employed for the temporal discretization. The numerical method is applied to study the unsteady flow field of a subsonic turbine stage and the unsteady flow field inside a transonic, high-through-flow, axial compressor stage. The stage calculation is performed by coupling the stator and the rotor flow fields at each time step through an over-laid grid.

Copenhaver, W. W.↗

NASA NTRS1993

Aerodynamic performance of a supersonic throughflow fan stage

The NASA LeRC has begun a program to design, analyze, build, and test a fan stage that can operate with supersonic axial velocities throughout the stage. This report presents the results from the initial experimental evaluation of the supersonic throughflow fan stage. The stage was tested with flows at subsonic and supersonic inlet axial Mach numbers and with rotor speeds at 40 to 100 percent of design speed. Overall performance as well as detailed performance obtained from radially and circumferentially traversing probes downstream of the stator vanes is presented. At the design rotor tip speed of 1500 ft/s and the design Mach number of 2.0, the measured overall pressure ratio was 2.22, and the corresponding adiabatic efficiency was 0.694. Transition from subsonic to supersonic flow was successfully demonstrated at 75, 90, and 100 percent of design speed. In that speed range, the stage 'unstart' (transition from supersonic to subsonic inlet axial conditions) occurred at an inlet axial Mach number of about 1.36, which corresponds to the Mach number at which the stage started.

Royce D Moore↗

NASA NTRS1951

Effect of Blade-surface Finish on Performance of a Single-stage Axial-flow Compressor

A set of modified NACA 5509-34 rotor and stator blades was investigated with rough-machine, hand-filed, and highly polished surface finishes over a range of weight flows at six equivalent tip speeds from 672 to 1092 feet per second to determine the effect of blade-surface finish on the performance of a single-stage axial-flow compressor. Surface-finish effects decreased with increasing compressor speed and with decreasing flow at a given speed. In general, finishing blade surfaces below the roughness that may be considered aerodynamically smooth on the basis of an admissible-roughness formula will have no effect on compressor performance.

COMPRESSORS - AXIAL FLOW↗

NASA NTRS1954

Analytical Study of Losses at Off-design Conditions for a Fixed Geometry Turbine

An analytical investigation was made to determine the off-design loss characteristics of a fixed-geometry turbine of which the experimental performance was known. The method of analysis utilized an effective loss parameter and assumed that the velocity normal to the blade entrance angle was lost as a total-pressure loss. The method also assumed constant tangential component of velocity between the station just upstream and just downstream of the stator and rotor trailing edge. Good correlation between the analytically and experimentally obtained performance was found over the entire map until limiting loading was approached. The large decrease in efficiency at low-speed high pressure ratios and at high-speed low pressure ratios was found in the analysis to be almost entirely due to the rotor incidence and turbine exit whirl losses. From the results of the investigation it was concluded that for turbines designed to operate efficiently at more than one point, the design must compromise rotor incidence angle and exit whirl losses.

Stewart, Warner L↗

NASA NTRS1957

Results of turbojet engine operation tests using a 50-50 mixture of JP-4 and tributyl borate as the fuel

An experimental investigation was conducted on a centrifugal-type turbojet engine using a 50-50 mixture of tributyl borate and JP-4 as the fuel to determine the magnitude and the location of the boric oxide deposits in the engine as well as the effect of these deposits on the engine performance. Large deposits of boric acid formed in the combustor walls and on the turbine rotor and stator blades. The deposits had no effect on the engine thrust.

Schafer, Louis J , Jr↗

NASA NTRS1953

Investigation of Water-spray Cooling of Turbine Blades in a Turbojet Engine

An analytical and experimental investigation was made with a J33-A-9 engine to determine the effectiveness of spray cooling as a means of increasing thrust by permitting engine operation at inlet-gas temperatures and speeds above rated. With the assumption of adequate spray cooling at a coolant-to-gas flow ratio of 3 percent, calculations for the sea-level static condition indicated a thrust may be achieved by engine operation at an inlet-gas temperature of 2000 degrees F and an overspeed of 10 percent. Of the water-injection configurations investigated experimentally, those located in the inner ring of the stator diaphragm provided the best cooling at rated engine speed.

HEAT TRANSFER, AERODYNAMIC↗

NASA NTRS1953

Altitude Starting Characteristics of an Afterburner with Autoignition and Hot-streak Ignition

An investigation was conducted in an altitude test chamber at the NACA Lewis Laboratory to determine the altitude starting characteristics of an afterburner with autoignition and with hot-streak ignition. Transient afterburner ignition data were obtained over a range of altitudes from 30,000 to 50,000 feet at a flight Mach number of 0.60. Afterburner ignition with a torch igniter located axially at approximately the midpoint of the combustion chamber was possible over the entire range, but ignition ignition with a torch igniter located in the transition section 1 5/8 inches upstream of the turbine stators proved unsatisfactory at an altitude of 50,000 feet due to the inability to obtain flame through the turbine. Increasing the afterburner-inlet total pressure at a constant afterburner fuel-air ratio decreased the afterburner ignition time. Hot-streak ignition was possible within 2 seconds after the time required to obtain the preset, normal afterburner fuel pressure, whereas autoignition required 4 to 7 seconds for the range of altitudes investigated. Following the ignition there was a period of oscillatory operation existing in the engine-afterburner before steady-state operation was attained. The time required for steady-state stable operation decreased as afterburner inlet total pressure increased. The duration of oscillations also decreased with hot-streak ignition because the fuel-air mixture was ignited before a large volume of combustible mixture was accumulated in the afterburner.

Engines, Control - Turbojet↗

NASA NTRS1956

Investigation of Turbines Suitable for Use in a Turbojet Engine with High Compressor Pressure Ratio and Low Compressor-tip Speed VI: Experimental Performance of Two-stage Turbine

The brake internal efficiency of the highly loaded two-stage turbine was 0.79 equivalent design work and speed. The maximum brake internal efficiency was 0.84. A radial survey revealed these major defects: (1) the first-rotor throat area was too large, and a large area of underturned flow existed near the tip;(2) considerable underturning existed at the second-stator outlet; and (3) tangential components of velocity at the turbine outlet amounted to 2.5 points in turbine efficiency.

Davison, Elmer H↗

NASA NTRS1957

Investigation of a 0.6 hub-tip radius-ratio transonic turbine designed for secondary-flow study I : design and experimental performance of standard turbine

Detailed design information including overall performance parameters, velocity diagrams, and blade surface velocities is presented. Experimental performance includes maps based on rating as well as total-pressure ratios showing the effect of exit whirl. Also included are results of surveys at the stator exit and downstream of the rotor at design speed and specific work. This information will be used as a standard for comparison with subsequent secondary-flow work.

Rohlik, Harold E↗

NASA NTRS1956

Experimental Investigation of a High Subsonic Mach Number Turbine Having Low Rotor Suction-surface Diffusion

The quasi-three-dimensional method for the aerodynamic design of both stator and rotor blades is presented for a turbine having high weight flow and high specific work output. The turbine was designed for a relative critical velocity ratio of 0.82 at the rotor hub inlet. At the condition of equivalent design work output and blade speed, the efficiency based on the overall actual total-pressure ratio was 0.875, which is almost at high as those obtained for conservatively designed turbines.

Hauser, Cavour H↗

NASA NTRS1956

Investigation of Turbines Suitable for Use in a Turbojet Engine with High Compressor Pressure Ratio and Low Compressor-tip Speed VII : Experimental Performance of Modified Two-stage Turbine

Modifying the original turbine by closing down the first-rotor throat area and shrouding the first and second rotors resulted in an over-all increase in in efficiency of 3.5 percentage points. At equivalent design work and speed the rating and aerodynamic efficiences of the modified turbine were 0.825 and 0.846, respectively. The maximum rating and erodynamic efficiencies were 0.875 and 0.906, respectively. A radial survey indicated improved firstively. A radial survey indicated improved first and second-stage efficiencies but showed that the effective throat areas of the second stator and rotor were too large.

Davison, Elmer H↗

NASA NTRS1947

Analysis of Performance of Jet Engine from Characteristics of Components I : Aerodynamic and Matching Characteristics of Turbine Component Determined with Cold Air

The performance of the turbine component of an NACA research jet engine was investigated with cold air. The interaction and the matching of the turbine with the NACA eight-stage compressor were computed with the combination considered as a jet engine. The over-all performance of the engine was then determined. The internal aerodynamics were studied to the extent of investigating the performance of the first stator ring and its influence on the turbine performance. For this ring, the stream-filament method for computing velocity distribution permitted efficient sections to be designed, but the design condition of free-vortex flow with uniform axial velocities was not obtained.

Goldstein, Arthur W↗

NASA NTRS1948

Analysis of effect of basic design variables on subsonic axial-flow-compressor performance

A blade-element theory for axial-flow compressors has been developed and applied to the analysis of the effects of basic design variables such as Mach number, blade loading, and velocity distribution on compressor performance. A graphical method that is useful for approximate design calculations is presented. The relations among several efficiencies useful in compressor design are derived and discussed. The possible gains in useful operating range obtainable by the use of adjustable stator blades are discussed and a rapid approximate method of calculating blade-angle resettings is shown by an example. The relative Mach number is shown to be a dominant factor in determining the pressure ratio.

Sinnette, John T , Jr↗

NASA NTRS1949

Two-dimensional compressible flow in turbomachines with conic flow surfaces

A general method of analysis is developed for two-dimensional, steady, compressible flow in stators or rotors of radial and mixed flow turbomachines with conic flow surfaces (surfaces of right circular cones generated by center line of flow passage in the axial-radial plane). The variables taken into account are: (1) tip speed of the rotor, (2) flow rate, (3) blade shape, (4) variation in passage height with radius, (5) number of blades, and (6) cone angle of the flow surface. Relaxation methods are used to solve the nonlinear differential equation for the stream function. Two numerical examples are presented; one for compressible and the other for incompressible flow in a centrifugal compressor with thin, straight blades. The results of these examples are given by plots of the streamlines, constant velocity-ratio lines, and constant pressure-ratio line.

Stanitz, John D↗