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At least 127 records · Page 7

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NASA NTRS2009

Establishing a Ballistic Test Methodology for Documenting the Containment Capability of Small Gas Turbine Engine Compressors

A test methodology currently employed for large engines was extended to quantify the ballistic containment capability of a small turboshaft engine compressor case. The approach involved impacting the inside of a compressor case with a compressor blade. A gas gun propelled the blade into the case at energy levels representative of failed compressor blades. The test target was a full compressor case. The aft flange was rigidly attached to a test stand and the forward flange was attached to a main frame to provide accurate boundary conditions. A window machined in the case allowed the projectile to pass through and impact the case wall from the inside with the orientation, direction and speed that would occur in a blade-out event. High-peed, digital-video cameras provided accurate velocity and orientation data. Calibrated cameras and digital image correlation software generated full field displacement and strain information at the back side of the impact point.

Heady, Joel↗

NASA NTRS2012

NASA Environmentally Responsible Aviation High Overall Pressure Ratio Compressor Research Pre-Test CFD

This paper describes a collaborative and cost-shared approach to reducing fuel burn under the NASA Environmentally Responsible Aviation project. NASA and General Electric (GE) Aviation are working together aa an integrated team to obtain compressor aerodynamic data that is mutually beneficial to both NASA and GE Aviation. The objective of the High OPR Compressor Task is to test a single stage then two stages of an advanced GE core compressor using state-of-the-art research instrumentation to investigate the loss mechanisms and interaction effects of embedded transonic highly-loaded compressor stages. This paper presents preliminary results from NASA's in-house multistage computational code, APNASA, in preparation for this advanced transonic compressor rig test.

Celestina, Mark L.↗

NASA NTRS2017

Toward Understanding Tip Leakage Flows in Small Compressor Cores Including Stator Leakage Flow

The focus of this work was to provide additional data to supplement the work reported in NASA/CR-2015-218868 (Berdanier and Key, 2015b). The aim of that project was to characterize the fundamental flow physics and the overall performance effects due to increased rotor tip clearance heights in axial compressors. Data have been collected in the three-stage axial research compressor at Purdue University with a specific focus on analyzing the multistage effects resulting from the tip leakage flow. Three separate rotor tip clearances were studied with nominal tip clearance gaps of 1.5 percent, 3.0 percent, and 4.0 percent based on a constant annulus height. Overall compressor performance was previously investigated at four corrected speedlines (100 percent, 90 percent, 80 percent, and 68 percent) for each of the three tip clearance configurations. This study extends the previously published results to include detailed steady and time-resolved pressure data at two loading conditions, nominal loading (NL) and high loading (HL), on the 100 percent corrected speedline for the intermediate clearance level (3.0 percent). Steady detailed radial traverses of total pressure at the exit of each stator row are supported by flow visualization techniques to identify regions of flow recirculation and separation. Furthermore, detailed radial traverses of time-resolved total pressures at the exit of each rotor row have been measured with a fast-response pressure probe. These data were combined with existing three-component velocity measurements to identify a novel technique for calculating blockage in a multistage compressor. Time-resolved static pressure measurements have been collected over the rotor tips for all rotors with each of the three tip clearance configurations for up to five loading conditions along the 100 percent corrected speedline using fast-response piezoresistive pressure sensors. These time-resolved static pressure measurements reveal new knowledge about the trajectory of the tip leakage flow through the rotor passage. Further, these data extend previous measurements identifying a modulation of the tip leakage flow due to upstream stator wake propagation. Finally, a novel instrumentation technique has been implemented to measure pressures in the shrouded stator cavities. These data provide boundary conditions relating to the flow across the shrouded stator knife seal teeth. Moreover, the utilization of fast-response pressure sensors provides a new look at the time-resolved pressure field, leading to instantaneous differential pressures across the seal teeth. Ultimately, the data collected for this project represent a unique data set which contributes to build a better understanding of the tip leakage flow field and its associated loss mechanisms. These data will facilitate future engine design goals leading to small blade heights in the rear stages of high pressure compressors and aid in the development of new blade designs which are desensitized to the performance penalties attributed to rotor tip leakage flows.

Berdanier, Reid A.↗

NASA NTRS1965

Aerodynamic Design of Axial Flow Compressors

An overview of 'Aerodynamic systems design of axial flow compressors' is presented. Numerous chapters cover topics such as compressor design, potential and viscous flow in two dimensional cascades, compressor stall and blade vibration, and compressor flow theory. Theoretical aspects of flow are also covered. This work is based substantially on a three-volume Confidential Research Memorandum f h t issued by the National Advisory Committee for Aeronautics in 1956 and declassified in .1958 for release to the general public. The material, which was of a fundamental character, has been updated for the present edition in the light of subsequent developments in the field and has been carefully reexamined for current validity by members of the staff of the NASA Lewis Research Center, Cleveland, Ohio. It is now published for the first time in a single volume to meet continuing demand for the information and to make it readily available in convenient form to the engineering community both within and outside the National Aeronautics and Space Administration.

Axial flow compressor↗

NASA NTRS1956

Effect of design over-all compressor pressure ratio division on acceleration characteristics of three hypothetical two-spool turbojet engines

The engines, each with a compressor overall total-pressure ratio of 12 and a design inner-turbine-inlet temperature of 2500 degrees R, were investigated at static sea-level conditions to determine the effect on transient performance of varying the desitn pressure ratio divisions 2-6, 3-4, and 4-3 between the outer and inner compressors. The transient considered was an acceleration from 40 to 100 percent design thrust. When the outer compressor of each engine reached design speed, the inner compressors were overspeeding, the maximum being only 1.7 over design mechanical speed. Acceleration times for the three engines were equal.

TURBINES - MATCHINE↗

NASA NTRS2015

Effects of Unsteady Flow Interactions on the Performance of a Highly-Loaded Transonic Compressor Stage

Effects of unsteady flow interactions on the aerodynamic performance of a highly-loaded transonic axial compressor are investigated in the present study. The primary focus of the study is to investigate how unsteady flow interactions between blade rows affect the aerodynamic performance of a highly-loaded transonic axial compressor. Recent experimental and numerical studies of current highly-loaded axial compressor performance indicated that predicting calculating the loss generation is very challenging with various analysis tools. In the present study, the effects of generation and transport of shock induced vortices on the compressor performance is investigated in detail.

Stall↗

NASA NTRSDate not supplied

Stall Margin Improvement in a Transonic Compressor with a Casing Treatment: Flow Mechanism

Effects of casing treatments for improved stall margin in transonic compressors are investigated. Two types of compressor casing treatments, circumferential grooves and axial slots casing treatments, are studied in detail. The primary objective of the current paper is to identify the underlying fluid mechanics that change the stall margin of transonic compressors with these two casing treatments. Large eddy simulation (LES) is applied to calculate the compressor flow fields with these two types of casing treatments. Calculated flow fields are first compared with measured experimental data to validate the LES simulation. Details of the calculated unsteady flow structures are examined to unearth the underlying flow mechanism for the improved stall margin. Further possible design optimization of the casing treatment for the stall margin extension without any performance penalty near the design condition is also explored.

compressor↗

NASA NTRS1971

Conceptual design study of a nuclear Brayton turboalternator-compressor

A comprehensive analysis and conceptual design study of the turboalternator-compressor components using HeXe as the working fluid was performed. The study was conducted in three phases: general configuration analysis (Phase 1), design variations (Phase 2), and conceptual design study (Phase 3). During the Phase 1 analysis, individual turbine, alternator, compressor, and bearing and seal designs were evaluated. Six turboalternator-compressor (TAC) configurations were completed. Phase 2 consisted of evaluating one selected Phase 1 TAC configuration to calculate its performance when operating under new cycle conditions, namely, one higher and one lower turbine inlet temperature and one case with krypton as the working fluid. Based on the Phase 1 and 2 results, a TAC configuration that incorporated a radial compressor, a radial turbine, a Lundell alternator, and gas bearings was selected. During Phase 3 a new layout of the TAC was prepared that reflects the cycle state points necessary to accommodate a zirconium hydride moderated reactor and a 400 Hz alternator. The final TAC design rotates at 24,000 rpm and produces 160 kWe, 480 V, 3-phase, 400 hertz power.

Source record↗

NASA NTRS1972

Nuclear Brayton turboalternator-compressor (TAC) conceptual design study

A comprehensive analysis and conceptual design study of the turboalternator-compressor components was performed using HeXe as the working fluid. Individual turbine, alternator, compressor, and bearing and seal designs were evaluated. Six turboalternator-compressor TAC configurations were completed. One TAC configuration was evaluated to calculate its performance when operating under new cycle conditions,namely, one higher and one lower turbine inlet temperature and one case with krypton as the working fluid. Based on the results, a TAC configuration that incorporated a radial compressor, a radial turbine, a Lundell Alternator, and gas bearings was selected. A new layout of the TAC was prepared that reflects the cycle state points necessary to accommodate a zirconium hydride moderated reactor and a 400 Hz alternator. The final TAC design rotates at 24,000 rpm and produces 160 kWe, 480V, 3-phase, 400 hertz power.

Mock, E. A.↗

NASA NTRS1973

Thermodynamic properties of UF sub 6 measured with a ballistic piston compressor

From experiments performed with a ballistic piston compressor, certain thermodynamic properties of uranium hexafluoride were investigated. Difficulties presented by the nonideal processes encountered in ballistic compressors are discussed and a computer code BCCC (Ballistic Compressor Computer Code) is developed to analyze the experimental data. The BCCC unfolds the thermodynamic properties of uranium hexafluoride from the helium-uranium hexafluoride mixture used as the test gas in the ballistic compressor. The thermodynamic properties deduced include the specific heat at constant volume, the ratio of specific heats for UF6, and the viscous coupling constant of helium-uranium hexafluoride mixtures.

Sterritt, D. E.↗

NASA NTRS1973

Advanced two-stage compressor program design of inlet stage

The aerodynamic design of an inlet stage for a two-stage, 10/1 pressure ratio, 2 lb/sec flow rate compressor is discussed. Initially a performance comparison was conducted for an axial, mixed flow and centrifugal second stage. A modified mixed flow configuration with tandem rotors and tandem stators was selected for the inlet stage. The term conical flow compressor was coined to describe a particular type of mixed flow compressor configuration which utilizes axial flow type blading and an increase in radius to increase the work input potential. Design details of the conical flow compressor are described.

Bryce, C. A.↗

NASA NTRS1974

Some comparisons of the flow characteristics of a turbofan compressor system with and without inlet pressure distortion

The measured effects of a circumferential distortion in inlet total pressure on the fan, low, and high compressor of an afterburning turbofan engine are presented and discussed. Extensive inner-stage instrumentation, combined with a unique test technique offered an accurate means of measuring the shifts in flow, performance, and stall mechanisms within the compressor. These effects are compared at one speed to the corresponding effects measured with undistorted inlet flow. The results show the rate at which the distorted flow areas were attenuated and rotated as well as the change in flow velocities that occurred at various points in the compressor. High response pressure traces indicated the location of stalls including the sequence of dynamic events from the onset and propagation of various stall-recovery events, to compressor surge, to the resulting hammershock.

Evans, D. G.↗

NASA NTRS1974

Some comparisons of the flow characteristics of a turbofan compressor system with and without inlet pressure distortion

The measured effects of a circumferential distortion in inlet total pressure on the fan, low, and high compressor of an afterburning turbofan engine are presented and discussed. Extensive inner-stage instrumentation, combined with a unique test technique offered an accurate means of measuring the shifts in flow, performance, and stall mechanisms within the compressor. These effects are compared at one speed to the corresponding effects measured with undistorted inlet flow. The results show the rate at which the distorted flow areas were attenuated and rotated, as well as the change in flow velocities that occurred at various points in the compressor. High response pressure traces indicated the location of stalls including the sequence of dynamic events from the onset and propagation of various stall-recovery events, to compressor surge, to the resulting hammershock.

Evans, D. G.↗

NASA NTRS1975

Turbofan compressor dynamics during afterburner transients

The effects of afterburner light-off and shut-down transients on compressor stability were investigated. Experimental results are based on detailed high-response pressure and temperature measurements on the Tf30-p-3 turbofan engine. The tests were performed in an altitude test chamber simulating high-altitude engine operation. It is shown that during both types of transients, flow breaks down in the forward part of the fan-bypass duct. At a sufficiently low engine inlet pressure this resulted in a compressor stall. Complete flow breakdown within the compressor was preceded by a rotating stall. At some locations in the compressor, rotating stall cells initially extended only through part of the blade span. For the shutdown transient, the time between first and last detected occurrence of rotating stall is related to the flow Reynolds number. An attempt was made to deduce the number and speed of propagation of rotating stall cells.

Kurkov, A. P.↗

NASA NTRS1975

Effects of inlet treatment location and treatment cavity depth on compressor noise

The ability of acoustic liners to reduce compressor noise inside and in front of an inlet was studied. An axial flow research compressor and a specially designed inlet were used inside an anechoic chamber. Acoustic and performance data were obtained for a range of inlet treatment locations and cavity depths to determine their effects on inlet noise over a range of blade passing frequencies. The greatest noise reductions in front of the inlet were obtained with acoustic treatment located close to the compressor and backed with the deepest cavities tested. Inside the inlet the maximum noise level reductions were obtained in the area of the treatment regardless of treatment location. No appreciable losses in compressor performance were measured.

Clark, L. R.↗

NASA NTRS1975

Turbofan compressor dynamics during afterburner transients

The effects of afterburner light-off and shut-down transients on the compressor stability are investigated. The reported experimental results are based on detailed high-response pressure and temperature measurements on the TF30-P-3 turbofan engine. The tests were performed in an altitude test chamber simulating high-altitude engine operation. It is shown that during both types of transients, flow breaks down in the forward part of the fan-bypass duct. At a sufficiently low engine inlet pressure this resulted in a compressor stall. Complete flow breakdown within the compressor was preceded by a rotating stall. At some locations in the compressor, rotating stall cells initially extended only through part of the blade span. For the shut-down transient the time between first and last detected occurrence of rotating stall is related to the flow Reynolds number. An attempt was made to deduce the number and speed of propagation of rotating stall cells.

Kurkov, A. P.↗

NASA NTRS1976

Turbofan compressor dynamics during afterburner transients

The effects of afterburner light-off and shut-down transients on the compressor stability are investigated. The reported experimental results are based on detailed high response pressure and temperature measurements on the TF30-P-3 turbofan engine. The tests were performed in an altitude test chamber simulating high altitude engine operation. It is shown that during both types of transients, flow breaks down in the forward part of the fan bypass duct. At a sufficiently low engine inlet pressure this resulted in a compressor stall. Complete flow breakdown within the compressor was preceded by a rotating stall. At some locations in the compressor, rotating stall cells initially extended only through part of the blade span. For the shutdown transient the time between first and last detected occurrence of rotating stall is related to the flow Reynolds number. An attempt was made to deduce the number and speed of propagation of rotating stall cells.

Kurkov, A. P.↗

NASA NTRS1975

Effects of casing treatment on a small, transonic axial-flow compressor

Improved axial compressor surge margin through effective rotor casing treatment has been identified from test results on large axial compressors. A modified scale of a large compressor was built and tested to determine if similar improvements in surge margin could be duplicated in small-size turbomachinery. In addition, the effects of rotor radial running clearance, both with and without casing treatment, were investigated and are discussed. Test results of the scale configuration are presented and compared to the parent compressor.

Holman, F. F.↗