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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 361 records · Page 20
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Aerodynamically forced vibration analysis of turbomachines

An account is given of theoretical considerations for the determination of turbomachine response under aerodynamic excitation, as in the cases of advanced turboprop rotors with highly swept blades and axial flow compressors. Dynamic response is characterized in terms of the normal modal coordinates of tuned rotating cyclic structures for both rigid and flexible hubs/disks. Implementation of the scheme is in NASTRAN; coriolis and centripetal accelerations are included, together with differential stiffness effects. Analytically predicted vibratory stresses for the blades of two different advanced turboprops are compared to those determined by wind tunnel tests conducted by NASA-Lewis.

Elchuri, V.↗
NASA NTRS1985

Modal forced vibration analysis of aerodynamically excited turbosystems

Theoretical aspects of a new capability to determine the vibratory response of turbosystems subjected to aerodynamic excitation are presented. Turbosystems such as advanced turbopropellers with highly swept blades, and axial-flow compressors and turbines can be analyzed using this capability. The capability has been developed and implemented in the April 1984 release of the general purpose finite element program NASTRAN. The dynamic response problem is addressed in terms of the normal modal coordinates of these tuned rotating cyclic structures. Both rigid and flexible hubs/disks are considered. Coriolis and centripetal accelerations, as well as differential stiffness effects are included. Generally non-uniform steady inflow fields and uniform flow fields arbitrarily inclined at small angles with respect to the axis of rotation of the turbosystem are considered sources of aerodynamic excitation. The spatial non-uniformities are considered to be small deviations from a principally uniform inflow. Subsonic and supersonic relative inflows are addressed, with provision for linearly interpolating transonic airloads.

Elchuri, V.↗
NASA NTRS1985

NASTRAN supplemental documentation for modal forced vibration analysis of aerodynamically excited turbosystems

This report is a supplemental NASTRAN document for a new capability to determine the vibratory response of turbosystems subjected to aerodynamic excitation. Supplements of NASTRAN Theoretical, User's, Programmer's, and Demonstration Manuals are included. Turbosystems such as advanced turbopropellers with highly swept blades, and axial-flow compressors and turbines can be analyzed using this capability, which has been developed and implemented in the April 1984 release of the general purpose finite element program NASTRAN. The dynamic response problem is addressed in terms of the normal modal coordinates of these tuned rotating cyclic structures. Both rigid and flexible hubs/disks are considered. Coriolis and centripetal accelerations, as well as differential stiffness effects are included. Generally nonuniform steady inflow fields and uniform flow fields arbitrarily inclined at small angles with respect to the axis of rotation of the turbosystem are considered as the sources of aerodynamic excitation. The spatial nonuniformities are considered to be small deviations from a principally uniform inflow. Subsonic relative inflows are addressed, with provision for linearly interpolating transonic airloads.

Elchuri, V.↗
NASA NTRS1988

Automated design of controlled-diffusion blades

A numerical automation procedure has been developed to be used in conjunction with an inverse hodograph method for the design of controlled diffusion blades. With this procedure a cascade of airfoils with a prescribed solidity, inlet Mach number, inlet air flow angle, and air flow turning can be produced automatically. The trailing edge thickness of the airfoil, an important quantity in inverse methods, is also prescribed. The automation procedure consists of a multidimensional Newton iteration in which the objective design conditions are achieved by acting on the hodograph input parameters of the underlying inverse code. The method, although more general in scope, is applied in this paper to the design of axial flow compressor blade sections, and a wide range of examples is presented.

Sanz, J. M.↗
NASA NTRS1950

Altitude Performance Characteristics of Tail-pipe Burner with Convergingconical Burner Section on J47 Turbojet Engine

An investigation of turbojet-engine thrust augmentation by means of tail-pipe burning was conducted in the NACA Lewis altitude wind tunnel. Performance data were obtained with a tail-pipe burner having a converging conical burner section installed on an axial-flow-compressor type turbojet engine over a range of simulated flight conditions and tail-pipe fuel-air ratios with a fixed-area exhaust nozzle. A maximum tail-pipe combustion efficiency of 0.86 was obtained at an altitude of 15,000 feet and a flight Mach number of 0.23. Tail-pipe burner operation was possible up to an altitude of 45,000 feet at a flight Mach number of 0.23.

Prince, William R↗
NASA NTRS1951

Investigation of Altitude Starting and Acceleration Characteristics of J47 Turbojet Engine

An investigation was conducted on an axial-flow-compressor type turbojet engine in the NACA Lewis altitude wind tunnel to determine the operational characteristics of several ignition systems, cross-fire tube configurations and fuel systems over a range of simulated flight conditions. The opposite-polarity-type spark plug provided the most satisfactory ignition. Increasing the cross-fire-tube diameter improved intercombustor flame propagation. At high windmilling speeds, accelerations to approximately 6200 rpm could be made at a preset constant throttle position. The use of a variable-area nozzle reduced acceleration time.

IGNITION SYSTEMS↗
NASA NTRS1950

Effect of fuel volatility on altitude starting limits of a turbojet engine

The effect of fuel volatility on altitude starting limits of an axial-flow-compressor-type turbojet engine was investigated using fuels with Reid vapor pressures of 1.1 and 5.4 pounds per square inch. At flight Mach numbers from 0.40 to 0.85, the AN-F-58 fuel allowed consistent windmilling at altitudes 2000 to 8000 feet higher than was obtained with the 1.1-pound Reid vapor pressure fuel. At a flight Mach number of 0.25, ignition could not be established at any altitude with the lower-volatility fuel.

COMBUSTION - IGNITION OF GASES↗
NASA NTRS1955

Performance of a High-solidity High-pressure-ratio Transonic Rotor

A high-solidity low-aspect-ratio transonic rotor with an inlet hub-tip ratio of 0.52 was investigated experimentally. The rotor developed a total-pressure tatio of 1.93 and an efficiency of 92 percent at an equivalent wheel tip speed of 1036 feet per second. The feasibility of using this rotor as a component of an inlet stage of an axial-flow compressor was investigated.

Neumann, Harvey E↗
NASA NTRS1995

Improved Speed Control System for the 87,000 HP Wind Tunnel Drive

This paper describes the design, installation, and integrated systems tests for a new drive motor speed control system which was part of a recent rehab project for the NASA Lewis 8x6 Supersonic Wind Tunnel. The tunnel drive consists of three mechanically-coupled 29,000 HP wound rotor induction motors driving an axial flow compressor. Liquid rheostats are used to vary the impedance of the rotor circuits, thus varying the speed of the drive system. The new design utilizes a distributed digital control system with a dual touch screen CRT operator console to provide alarm monitoring, logging, and trending. The liquid rheostats are driven by brushtype servomotor systems with magnetostrictive linear displacement transducers used for position feedback. The new system achieved all goals for speed variations with load, motor load balance, and control of total power.

Becks, Edward A.↗
NASA NTRS1996

Aircraft Noise Prediction Program (ANOPP) Fan Noise Prediction for Small Engines

The Fan Noise Module of ANOPP is used to predict the broadband noise and pure tones for axial flow compressors or fans. The module, based on the method developed by M. F. Heidmann, uses empirical functions to predict fan noise spectra as a function of frequency and polar directivity. Previous studies have determined the need to modify the module to better correlate measurements of fan noise from engines in the 3000- to 6000-pound thrust class. Additional measurements made by AlliedSignal have confirmed the need to revise the ANOPP fan noise method for smaller engines. This report describes the revisions to the fan noise method which have been verified with measured data from three separate AlliedSignal fan engines. Comparisons of the revised prediction show a significant improvement in overall and spectral noise predictions.

Hough, Joe W.↗
NASA NTRS1959

Experimental Investigation of a 0.35 Hub-Tip Radius Ratio Transonic Axial Flow Rotor Designed for 40 Pounds per Second per Square Foot with a Design Tip Diffusion Factor of 0.20

In order to determine the effect of a low design diffusion factor on the performance of a transonic axial-flow compressor rotor, a high-specific-flow rotor with a 0.35 hub-tip radius ratio was designed, fabricated and tested. This rotor used a design tip diffusion factor of 0.20 with a design corrected specific weight flow of 40 pounds per second per square foot of frontal area, a total-pressure ratio of 1.27, and an adiabatic efficiency of 0.96. The design, rotor performance, and blade element performance are presented with a discussion on rotor shock losses and a comparison with a similarly designed rotor with a tip diffusion factor of 0.35. At the design corrected tip speed of 1100 feet per second, a peak rotor adiabatic efficiency of 0.88 was attained at a corrected specific weight flow of 39 pounds per second per square foot of frontal area with a mass-averaged total-pressure ratio of 1.27. The blade element tip diffusion factor was 0.281, which is 0.08 higher than the design value of 0.20. Peak efficiencies of 0.95, 0.91, 0.89, and 0.85 were obtained at 70, 80, 90, and 110 percent of design speed, respectively. Comparison of the performance of the rotor reported herein and a similarly designed rotor with increased blade loading indicates that higher blade loading results in a more desirable rotor because of a higher pressure ratio and equivalent efficiency. Computed values of shock losses at the rotor tip section indicate that the losses at peak efficiency are primarily a function of shock losses since the profile losses are only a small percentage of the total loss.

Yasaki, Paul T.↗
NASA NTRS1961

Experimental Investigation of an Air-Cooled Turbine Operating in a Turbojet Engine at Turbine Inlet Temperatures up to 2500 F

An experimental investigation was made of an air-cooled turbine at average turbine inlet temperatures up to 2500 F. A modified production-model 12-stage axial-flow-compressor turbojet engine operating in a static sea-level stand was used as the test vehicle. The modifications to the engine consisted of the substitution of special combustor and turbine assemblies and double-walled exhaust ducting for the standard parts of the engine. All of these special parts were air-cooled to withstand the high operating temperatures of the investigation. The air-cooled turbine stator and rotor blades were of the corrugated-insert type. Leading-edge tip caps were installed on the rotor blades to improve leading-edge cooling by diverting the discharge of coolant to regions of lower gas pressure toward the trailing edge of the blade tip. Caps varying in length from 0.15- to 0.55-chord length were used in an attempt to determine the optimum cap length for this blade. The engine was operated over a range of average turbine inlet temperatures from about 1600 to about 2500 F, and a range of average coolant-flow ratios of 0.012 to 0.065. Temperatures of the air-cooled turbine rotor blades were measured at all test conditions by the use of thermocouples and temperature-indicating paints. The results of the investigation indicated that this type of blade is feasible for operation in turbojet engines at the average turbine inlet temperatures and stress levels tested(maximums of 2500 F and 24,000 psi, respectively). An average one-third-span blade temperature of 1300 F could be maintained on 0.35-chord tip cap blades with an average coolant-flow ratio of about 0.022 when the average turbine inlet temperature was 2500 F and cooling-air temperature was about 260 F. All of the leading-edge tip cap lengths improved the cooling of the leading-edge region of the blades, particularly at low average coolant-flow ratios. At high gas temperatures, such parts as the turbine stator and the combustor liners are likely to be as critical as the turbine rotor blades.

Cochran, Reeves P.↗
NASA NTRS2000

Blading System and Method For Controlling Structural Vibrations

A new blading system for controlling the structural vibrations in axial-flow compressors, turbines, or fans, as in aircraft engines and like turbomachines including a stator disc and a rotor disc is presented. The rotor disc defines several radial hubs that retain the rotor blading systems. Each blading system includes a blade formed of an airfoil, and a root attachment which is dimensioned to fit within, and to engage a corresponding hub. Viscoelastic dampers are selectively applied to the outer surfaces of the root attachment on which compressive or shear forces are likely to develop, intermediate the root attachment and the hub, for compression therebetween upon rotation of the rotor disc, in order to dampen structural vibrations. One advantage presented by the viscoelastic dampers lies in its simplicity, efficiency, cost effectiveness, and its ability to be retrofitted into existing turbomachines with minor surface treatment of the root attachments. Furthermore, since the dampers are not exposed to the inflowing airstream, they do not affect the aerodynamic performance of the turbomachine. Another feature of the damping system is that it provides a significant source of damping to minimize destructive structural vibrations, thereby increasing the durability of the turbomachine, and reducing acoustic noise accompanying high amplitude vibrations.

Nguyen, Nhan↗
NASA NTRS1945

An Experimental Investigation of NACA Submerged-Duct Entrances

The results of a preliminary investigation of submerged duct entrances are presented. It is shown that an entrance of this type possess desirable critical speed and pressure recovery characteristics when used on a fuselage or nacelle in a region of low incremental velocity and thin boundary layer. The data obtained indicate that submerged entrances are most suitable for use with internal-flow systems which diffuse the air only a small amount: for example, those used with jet motors which have axial-flow compressors. Where complete diffusion of the air is required, fuselage-nose or wing leading edge inlets may prove to be superior. The results of the investigation have been prepared in such a form as to permit their use by a designer and the application of these data to a specific design is discussed.

Frick, Charles W.↗
NASA NTRS1995

Coupled Aerodynamic-Thermal-Structural (CATS) Analysis

Coupled Aerodynamic-Thermal-Structural (CATS) Analysis is a focused effort within the Numerical Propulsion System Simulation (NPSS) program to streamline multidisciplinary analysis of aeropropulsion components and assemblies. Multidisciplinary analysis of axial-flow compressor performance has been selected for the initial focus of this project. CATS will permit more accurate compressor system analysis by enabling users to include thermal and mechanical effects as an integral part of the aerodynamic analysis of the compressor primary flowpath. Thus, critical details, such as the variation of blade tip clearances and the deformation of the flowpath geometry, can be more accurately modeled and included in the aerodynamic analyses. The benefits of this coupled analysis capability are (1) performance and stall line predictions are improved by the inclusion of tip clearances and hot geometries, (2) design alternatives can be readily analyzed, and (3) higher fidelity analysis by researchers in various disciplines is possible. The goals for this project are a 10-percent improvement in stall margin predictions and a 2:1 speed-up in multidisciplinary analysis times. Working cooperatively with Pratt & Whitney, the Lewis CATS team defined the engineering processes and identified the software products necessary for streamlining these processes. The basic approach is to integrate the aerodynamic, thermal, and structural computational analyses by using data management and Non-Uniform Rational B-Splines (NURBS) based data mapping. Five software products have been defined for this task: (1) a primary flowpath data mapper, (2) a two-dimensional data mapper, (3) a database interface, (4) a blade structural pre- and post-processor, and (5) a computational fluid dynamics code for aerothermal analysis of the drum rotor. Thus far (1) a cooperative agreement has been established with Pratt & Whitney, (2) a Primary Flowpath Data Mapper has been prototyped and delivered to General Electric Aircraft Engines and Pratt & Whitney for evaluation, (3) a collaborative effort has been initiated with the National Institute of Standards and Testing to develop a Standard Data Access Interface, and (4) a blade tip clearance capability has been implemented into the Structural Airfoil Blade Engineering Routine (SABER) program. We plan to continue to develop the data mappers and data management tools. As progress is made, additional efforts will be made to apply these tools to propulsion system applications.

Source record↗
NASA NTRS1997

Turbomachinery Flows Modeled

Last year, researchers at the NASA Lewis Research Center used the average passage code APNASA to complete the largest three-dimensional simulation of a multistage axial flow compressor to date. Consisting of 29 blade rows, the configuration is typical of those found in aeroengines today. The simulation, which was executed on the High Performance Computing and Communications (HPCC) Program IBM SP2 parallel computer located at the NASA Ames Research Center, took nearly 90 hr to complete. Since the completion of this activity, a fine-grain, parallel version of APNASA has been written by a team of researchers from General Electric, NASA Lewis, and NYMA. Timing studies performed on the SP2 have shown that, with eight processors assigned to each blade row, the simulation time is reduced by a factor of six. For this configuration, the simulation time would be 15 hr. The reduction in computing time indicates that an overnight turnaround of a multistage configuration simulation is feasible. In addition, average passage forms of two-equation turbulence models were formulated. These models are currently being incorporated into APNASA.

Adamczyk, John J.↗
NASA NTRS1994

Rotating stall suppression

Rotating stall in an axial-flow compressor is suppressed by the positioning of a fixed inlet flow divider in the annular inlet flow passage upstream of the compressor. The inlet flow divider is aligned with the flow of fluid through the duct and acts to block or interfere with any rotating wave in the inlet and thereby suppresses rotating stall in the compressor.

Moore, Franklin K.↗
NASA NTRS1952

Effect of Various Blade Modifications on Performance of a 16-Stage High-pressure-ratio Axial-flow Compressor.Angles 3 deg: Effect on Over-all Performance Characteristics of Decreasing Twelfth Through Fifteenth State Stator-blade Angles 3 deg - 1

The stator-blade angles in the twelfth to fifteenth stages of a 16-stage high-pressure-ratio axial-flow compressor were decreased 3 deg The over-all performance of this compressor is compared with the performance of the same compressor with standard blade angles. The matching characteristics of the modified compressor and a two-stage turbine were also obtained and compared with those of the compressor with the original blade angles and the same turbine.

Medeiros, Arthur A.↗