Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “compressors”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 973 records · Page 54
All sourcesNASA NTRSDOE OSTI
NASA NTRS1947

Investigation of Axial-flow Fan and Compressor Rotors Designed for Three-dimensional Flow

An investigation has been conducted to determine whether three-dimensional flows may be utilized in axial-flow fan and compressor rotors so that the spanwise load distribution may be varied to obtain high pressure rise. Two rotors, one with approximately uniform and one with solid-body downstream tangential-velocity distributions, were designed and tested at the design blade angle. Radial surveys of total pressure, static pressure, and flow angle were made upstream and downstream of the test rotors through a quantity-coefficient range. Tests of the solid-body rotor were also conducted at a large value of tip clearance. The results indicated that the three-dimensional flows may be utilized with high efficiency and that the three-dimensional theory used in conjunction with two-dimensional cascade data is sufficiently accurate for design purposes. The tests also showed that the tip-clearance losses of rotors highly loaded at the tips are not excessive. The existing three-dimensional theory in simplified for and an illustrative rotor design are presented in appendixes.

Kahane, A↗
NASA NTRS1994

System Measures Pressures Aboard A Compressor Rotor

Rotating pressure-measuring instrumentation includes on-board calibration standard. Computer-controlled, multichannel instrumentation system acquires pressure measurements from sensors mounted in 1.52-m-diameter rotor of compressor. Includes 5 miniature, electronically scanned pressure (ESP) modules, each containing 48 piezoresistive pressure sensors, pneumatic calibration valve, and electronic circuits for addressing and amplifying output of each sensor. Modules mounted on centerline of rotor, on instrumentation tower located inside nose cone of rotor. Subsystem designed to convert analog signal to distinct frequency without significantly affecting accuracy.

Freedman, Robert J.↗
NASA NTRS1994

Piston-Skirt Lubrication System For Compressor

Piston-skirt lubrication system provides steady supply of oil to piston rings of gas compressor. No need for oil-filled crankcase or external oil pump. Instead, part of each piston acts as its own oil pump circulating oil from reservoir. Annular space at bottom of piston and cylinder constitutes working volume of small oil pump. Depending on application, reservoir open to atmosphere, or sealed and pressurized in bellows to prevent contact between oil and atmosphere. Filter removes particles worn away from piston rings and cylinder wall during normal operation.

Schroeder, Edgar C.↗
NASA NTRS1995

Enhanced capabilities and modified users manual for axial-flow compressor conceptual design code CSPAN

Modifications made to the axial-flow compressor conceptual design code CSPAN are documented in this report. Endwall blockage and stall margin predictions were added. The loss-coefficient model was upgraded. Default correlations for rotor and stator solidity and aspect-ratio inputs and for stator-exit tangential velocity inputs were included in the code along with defaults for aerodynamic design limits. A complete description of input and output along with sample cases are included.

Glassman, Arthur J.↗
NASA NTRS1995

The effect of adding roughness and thickness to a transonic axial compressor rotor

The performance deterioration of a high speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54-3.18 RMS microns (100-125 RMS microinches) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10% at the hub and 20% at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9% reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effect of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254-0.508 RMS microns (10-20 RMS microinches), compared to the bare metal blade surface finish of 0.508 RMS microns (20 RMS microinches). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60%, 80%, and 100% of design speed. The results indicate that thickness/roughness over the first 10% of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70% of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-3D Navier-Stokes flow solver which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that coating the blade causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

Suder, Kenneth L.↗
NASA NTRS1995

Experimental and Computational Investigation of the Tip Clearance Flow in a Transonic Axial Compressor Rotor

Experimental and computational techniques are used to investigate tip clearance flows in a transonic axial compressor rotor at design and part speed conditions. Laser anemometer data acquired in the endwall region are presented for operating conditions near peak efficiency and near stall at 100% design speed and at near peak efficiency at 60% design speed. The role of the passage shock/leakage vortex interaction in generating endwall blockage is discussed. As a result of the shock/vortex interaction at design speed, the radial influence of the tip clearance flow extends to 20 times the physical tip clearance height. At part speed, in the absence of the shock, the radial extent is only 5 times the tip clearance height. Both measurements and analysis indicate that under part-speed operating conditions a second vortex, which does not originate from the tip leakage flow, forms in the endwall region within the blade passage and exits the passage near midpitch. Mixing of the leakage vortex with primary flow downstream of the rotor at both design and part speed conditions is also discussed.

Suder, Kenneth L.↗
NASA NTRS1995

Better Gas-Gap Thermal Switches For Sorption Compressors

Gas-gap thermal switches associated with sorption compressors of some heat pumps and cryogenic systems designed for higher performance, according to proposal, by introducing controlled turbulent flows into gas gaps. Utilizes convection in turbulent flow to transfer heat at greater rate. Design takes advantage of flow of working fluid. Working fluid also serve as heat transfer medium in gas gap.

Bhandari, Pradeep↗
NASA NTRS1995

A packet data compressor

We are in the preliminary stages of creating an operational system for losslessly compressing packet data streams. The end goal is to reduce costs. Real world constraints include transmission in the presence of error, tradeoffs between the costs of compression and the costs of transmission and storage, and imperfect knowledge of the data streams to be transmitted. The overall method is to bring together packets of similar type, split the data into bit fields, and test a large number of compression algorithms. Preliminary results are very encouraging, typically offering compression factors substantially higher than those obtained with simpler generic byte stream compressors, such as Unix Compress and HA 0.98.

Grunes, Mitchell R.↗
NASA NTRS1995

Design geometry and design/off-design performance computer codes for compressors and turbines

This report summarizes some NASA Lewis (i.e., government owned) computer codes capable of being used for airbreathing propulsion system studies to determine the design geometry and to predict the design/off-design performance of compressors and turbines. These are not CFD codes; velocity-diagram energy and continuity computations are performed fore and aft of the blade rows using meanline, spanline, or streamline analyses. Losses are provided by empirical methods. Both axial-flow and radial-flow configurations are included.

Glassman, Arthur J.↗
NASA NTRS1995

Development of a fiber optic compressor blade sensor

A complete working prototype of the fiber optic blade tip sensor was first tested in the laboratory, followed by a thorough evaluation at NASA W8 Single Compressor Stage Facility in Lewis Research Center. Subsequently, a complete system with three parallel channels was fabricated and delivered to Dr. Kurkov. The final system was tested in the Subsonic Wind Tunnel Facility, in parallel with The General Electric Company's light probe system. The results at all operating speeds were comparable. This report provides a brief description of the system and presents a summary of the experimental results.

Dhadwal, Harbans Singh↗
NASA NTRS1996

Calculation of tip clearance effects in a transonic compressor rotor

The flow through the tip clearance region of a transonic compressor rotor (NASA rotor 37) was computed and compared to aerodynamic probe and laser anemometer data. Tip clearance effects were modeled both by gridding the clearance gap and by using a simple periodicity model across the ungridded gap. The simple model was run with both the full gap height, and with half the gap height to simulate a vena-contracta effect. Comparisons between computed and measured performance maps and downstream profiles were used to validate the models and to assess the effects of gap height on the simple clearance model. Recommendations were made concerning the use of the simple clearance model. Detailed comparisons were made between the gridded clearance gap solution and the laser anemometer data near the tip at two operating points. The computer results agreed fairly well with the data but overpredicted the extent of the casing separation and underpredicted the wake decay rate. The computations were then used to describe the interaction of the tip vortex, the passage shock, and the casing boundary layer.

Chima, R. V.↗
NASA NTRS1996

Variable-Speed Induction Motor Drives for Aircraft Environmental Control Compressors

New, more-efficient designs for aircraft jet engines are not capable of supplying the large quantities of bleed air necessary to provide pressurization and air conditioning for the environmental control systems (ECS) of the next generation of large passenger aircraft. System analysis and engineering have determined that electrically-driven ECS can help to maintain the improved fuel efficiencies; and electronic controllers and induction motors are now being developed in a NASA/NPD SBIR Program to drive both types of ECS compressors. Previous variable-speed induction motor/controller system developments and publications have primarily focused on field-oriented control, with large transient reserve power, for maximum acceleration and optimum response in actuator and robotics systems. The application area addressed herein is characterized by slowly-changing inputs and outputs, small reserve power capability for acceleration, and optimization for maximum efficiency. This paper therefore focuses on the differences between this case and the optimum response case, and shows the development of this new motor/controller approach. It starts with the creation of a new set of controller requirements. In response to those requirements, new control algorithms are being developed and implemented in an embedded computer, which is integrated into the motor controller closed loop. Buffered logic outputs are used to drive the power switches in a resonant-technology, power processor/motor-controller, at switching/resonant frequencies high enough to support efficient high-frequency induction motor operation at speeds up to 50,000-RPA

Mildice, J. W.↗
NASA NTRS1996

User's Guide for MSAP2D: A Program for Unsteady Aerodynamic and Aeroelastic (Flutter and Forced Response) Analysis of Multistage Compressors and Turbines

This guide describes the input data required for using MSAP2D (Multi Stage Aeroelastic analysis Program - Two Dimensional) computer code. MSAP2D can be used for steady, unsteady aerodynamic, and aeroelastic (flutter and forced response) analysis of bladed disks arranged in multiple blade rows such as those found in compressors, turbines, counter rotating propellers or propfans. The code can also be run for single blade row. MSAP2D code is an extension of the original NPHASE code for multiblade row aerodynamic and aeroelastic analysis. Euler equations are used to obtain aerodynamic forces. The structural dynamic equations are written for a rigid typical section undergoing pitching (torsion) and plunging (bending) motion. The aeroelastic equations are solved in time domain. For single blade row analysis, frequency domain analysis is also provided to obtain unsteady aerodynamic coefficients required in an eigen analysis for flutter. In this manual, sample input and output are provided for a single blade row example, two blade row example with equal and unequal number of blades in the blade rows.

Reddy, T. S. R.↗
NASA NTRS1997

Demonstration of PIV in a Transonic Compressor

Particle Imaging Velocimetry (PIV) is a powerful measurement technique which can be used as an alternative or complementary approach to Laser Doppler Velocimetry (LDV) in a wide range of research applications. PIV data are measured simultaneously at multiple points in space, which enables the investigation of the non-stationary spatial structures typically encountered in turbomachinery. Many of the same issues encountered in the application of LDV techniques to rotating machinery apply in the application of PIV. Preliminary results from the successful application of the standard 2-D PIV technique to a transonic axial compressor are presented. The lessons learned from the application of the 2-D PIV technique will serve as the basis for applying 3-component PIV techniques to turbomachinery.

Wernet, Mark P.↗
NASA NTRS1998

Demonstration of PIV in a Transonic Compressor

Particle Imaging Velocimetry (PIV) is a powerful measurement technique which can be used as an alternative or complementary approach to Laser Doppler Velocimetry (LDV) in a wide range of research applications. PIV data are measured simultaneously at multiple points in space, which enables the investigation of the non-stationary spatial structures typically encountered in turbomachinery. Many of the same issues encountered in the application of LDV techniques to rotating machinery apply in the application of PIV. Preliminary results from the successful application of the standard 2-D PIV technique to a transonic axial compressor are presented. The lessons learned from the application of the 2-D PIV technique will serve as the basis for applying 3-component PIV techniques to turbomachinery.

Wernet, Mark P.↗
NASA NTRS1998

Wave Augmented Diffusers for Centrifugal Compressors

A conceptual device is introduced which would utilize unsteady wave motion to slow and turn flows in the diffuser section of a centrifugal compressor. The envisioned device would substantially reduce the size of conventional centrifugal diffusers by eliminating the relatively large ninety degree bend needed to turn the flow from the radial/tangential to the axial direction. The bend would be replaced by a wall and the flow would instead exit through a series of rotating ports located on a disk, adjacent to the diffuser hub, and fixed to the impeller shaft. The ports would generate both expansion and compression waves which would rapidly transition from the hub/shroud (axial) direction to the radial/tangential direction. The waves would in turn induce radial/tangential and axial flow. This paper presents a detailed description of the device. Simplified cycle analysis and performance results are presented which were obtained using a time accurate, quasi-one-dimensional CFD code with models for turning, port flow conditions, and losses due to wall shear stress. The results indicate that a periodic wave system can be established which yields diffuser performance comparable to a conventional diffuser. Discussion concerning feasibility, accuracy, and integration follow.

Paxson, Daniel E.↗
NASA NTRS1959

Performance of Typical Rear-Stage Axial-Flow Compressor Rotor Blade Row at Three Different Blade Setting Angles

A comparison of the performance of a single-stage rotor run at three different blade setting angles is presented. The rotor was of a design typical for a last stage of a multistage compressor. At each setting angle, the rotor blade row was operated from 53 to 100 percent of equivalent maximum speed (850 ft/sec tip speed) at constant inlet pressure. Hot-wire anemometry was used to observe rotating-stall and surge patterns in time unsteady flow. Results indicated that an increase in peak pressure ratio and an increase in maximum equivalent weight flow were obtained at each speed investigated when the blade setting angle was decreased. An increase in peak efficiency was achieved with decrease in blade setting angle for part of the range of speeds investigated. However, the peak efficiencies for the three blade setting angles were approximately the same at the maximum speed investigated. The flow ranges for all three configurations were about the same at minimum speed and decreased at almost the same rate when the rotative speed was increased through part of the range of speeds investigated. At maximum speed, the flow range for the smallest setting angle was considerably less than the flow range for the other two configurations. A decrease in efficiency and flow range for the smallest blade setting angle at maximum speed can be attributed primarily to a Mach number effect. In addition, because of the difference in projected axial chord lengths at the casing wall, some effect on performance could be expected from the change in three-dimensional flow occurring at the tip. Rotating-stall characteristics for the two smaller blade setting angles were essentially the same. Only surge could be detected for the largest blade setting angle in the unstable-flow region of operation.

Kussoy, Marvin I.↗