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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1920

The problem of the turbo-compressor

In terminating the study of the adaptation of the engine to the airplane, we will examine the problem of the turbo-compressor,the first realization of which dates from the war; this will form an addition to the indications already given on supercharging at various altitudes. This subject is of great importance for the application of the turbo-compressor worked by the exhaust gases. As a matter of fact, a compressor increasing the pressure in the admission manifold may be controlled by the engine shaft by means of multiplication gear or by a turbine operated by the exhaust gas. Assuming that the increase of pressure in the admission manifold is the same in both cases, the pressure in the exhaust manifold would be greater in the case in which the compressor is worked by the exhaust gas and there would result a certain reduction of engine power which we must be able to calculate. On the other hand , if the compressor is controlled by the engine shaft, a certain fraction of the excess power supplied is utilized for the rotation of the compressor. In order to compare the two systems, it is there-fore necessary to determine the value of the reduction of power due to back pressure when the turbine is employed.

Devillers, Rene↗
NASA NTRS1993

Cooled spool piston compressor

A hydraulically powered gas compressor receives low pressure gas and outputs a high pressure gas. The housing of the compressor defines a cylinder with a center chamber having a cross-sectional area less than the cross-sectional area of a left end chamber and a right end chamber, and a spool-type piston assembly is movable within the cylinder and includes a left end closure, a right end closure, and a center body that are in sealing engagement with the respective cylinder walls as the piston reciprocates. First and second annual compression chambers are provided between the piston enclosures and center housing portion of the compressor, thereby minimizing the spacing between the core gas and a cooled surface of the compressor. Restricted flow passageways are provided in the piston closure members and a path is provided in the central body of the piston assembly, such that hydraulic fluid flows through the piston assembly to cool the piston assembly during its operation. The compressor of the present invention may be easily adapted for a particular application, and is capable of generating high gas pressures while maintaining both the compressed gas and the compressor components within acceptable temperature limits.

Morris, Brian G.↗
NASA NTRS1995

NASA Low-speed Axial Compressor for Fundamental Research

A low-speed multistage axial compressor built by the NASA Lewis Research Center is described. The purpose of this compressor is to increase the understanding of the complex flow phenomena within multistage axial compressors and to obtain detailed data from a multistage compressor environment for use in developing and verifying models for computational fluid dynamic code assessment. The compressor has extensive pressure instrumentation in both stationary and rotating frames of reference, and has provisions for flow visualization and laser velocimetry. The compressor will accommodate rotational speeds to 1050 rpm and is rated at a pressure ratio of 1.042.

Wasserbauer, Charles A.↗
NASA NTRS1996

Comparisons Between Unsteady Aerodynamic Events in a Gas Turbine Generator and an Identical Compressor Rig

Extensive testing done on a T55-L-712 turboshaft engine compressor in a compressor test rig is being followed by engine tests in progress as part of the Army Non-Recoverable Stall Program. Goals include a greater understanding of the gas turbine engine start cycle and compressor/engine operation in the regions 'beyond' the normal compressor stall line (rotating stall/surge). Rig steady state instrumentation consisted of 497 steady state pressure sensors and 153 temperature sensors. Engine instrumentation was placed in similar radial/axial locations and consists of 122 steady state pressure sensors and 65 temperature sensors. High response rig instrumentation consisted of 34 wall static pressure transducers. Rig and engine high response pressure transducers were located in the same axial/radial/circumferential locations in front of the first three stages. Additional engine high response instrumentation was placed in mach probes in front of the engine and on the compressor hub. This instrumentation allows for the generation of detailed stage characteristics, overall compressor mapping, and detailed analysis of dynamic compressor events.

Owen, A. Karl↗
NASA NTRS2001

An Investigation of Surge in a High-Speed Centrifugal Compressor Using Digital PIV

Compressor stall is a catastrophic breakdown of the flow in a compressor, which con lead to a loss of engine power, large pressure transients in the inlet/nacelle, and engine flameout. The implementation of active or passive strategies for controlling rotating stall and surge can significantly extend the stable operating range of a compressor without substantially sacrificing performance. It is crucial to identify the dynamic changes occurring in the flow field prior to rotating stall and surge in order to control these events successfully. Generally, pressure transducer measurements are made to capture the transient response of a compressor prior to rotating stall. In this investigation, Digital Particle Imaging Velocimetry (DPIV) is used in conjunction with dynamic pressure transducers to capture transient velocity and pressure measurements simultaneously in the nonstationary flow field during compressor surge. DPIV is an instantaneous, planar measurement technique that is ideally suited for studying transient flow phenomena in highspeed turbomachinery and has been used previously to map the stable operating point flow field in the diffuser of a high-speed centrifugal compressor. Through the acquisition of both DPIV images and transient pressure data, the time evolution of the unsteady flow during surge is revealed.

Wernet, Mark P.↗
NASA NTRS2002

An Investigation of Surge in a High-Speed Centrifugal Compressor Using Digital PIV

Compressor stall is a catastrophic breakdown of the flow in a compressor, which can lead to a loss of engine power, large pressure transients in the inlet/nacelle and engine flameout. The implementation of active or passive strategies for controlling rotating stall and surge can significantly extend the stable operating range of a compressor without substantially sacrificing performance. It is crucial to identify the dynamic changes occurring in the flow field prior to rotating stall and surge in order to successfully control these events. Generally, pressure transducer measurements are made to capture the transient response of a compressor prior to rotating stall. In this investigation, Digital Particle Imaging Velocimetry (DPIV) is used in conjunction with dynamic pressure transducers to simultaneously capture transient velocity and pressure measurements in the non-stationary flow field during compressor surge. DPIV is an instantaneous, planar measurement technique which is ideally suited for studying transient flow phenomena in high speed turbomachinery and has been used previously to successfully map the stable operating point flow field in the diffuser of a high speed centrifugal compressor. Through the acquisition of both DPIV images and transient pressure data, the time evolution of the unsteady flow during surge is revealed.

Wernet, Mark P.↗
NASA NTRS2003

Task II: Three-dimensional Rotating Stall Inception and Effects of Rotating Tip Clearance Asymmetry in Axial Compressors

The effects of two types of flow non-uniformity on stall inception behavior were assessed with linearized stability analyses of two compressor flow models. Response to rotating tip clearance asymmetries induced by a whirling rotor shaft or rotor height variations were investigated with a two-dimensional flow model. A 3-D compressor model was also developed to study the stability of both full-span and part-span rotating stall modes in annular geometries with radial flow variations. The studies focussed on (1) understanding what compressor designs were sensitive to these types of circumferential and spanwise flow non-uniformities, and (2) situations where 2-D stability theories were inadequate because of 3-D flow effects. Rotating tip clearance non-uniformity caused the greatest performance loss for shafts whirling at the rotating stall frequency. A whirling shaft displacement of 1 percent chord caused the stalling mass flow to rise by as much as 10 percent and the peak pressure rise to decrease by 6 percent. These changes were an order of magnitude larger than for equivalent-sized stationary or rotor-locked clearance asymmetries. Spanwise flow non-uniformity always destabilized the compressor, so that 2-D models over-predicted that stall margin compared to 3-D theory. The difference increased for compressors with larger spanwise variations of characteristic slope and reduced characteristic curvature near the peak. Differences between 2-D and 3-D stall point predictions were generally unacceptable (2 - 4 percent of flow coefficient) for single-stage configurations, but were less than 1 percent for multistage compressors. 2-D analyses predicted the wrong stall mode for specific cases of radial inlet flow distortion, mismatching and annulus area contraction, where higher-order radial modes led to stall. The stability behavior of flows with circumferential or radial non-uniformity was unified through a single stability criterion. The stall point for both cases was set by the integral around the annulus of the pressure rise characteristic slope, weighted by the amplitude of the mode shape. For the case of steady circumferential variations, this criterion reduced to the integrated mean slope (IMS) condition associated with steady inlet distortions. The rotating tip clearance asymmetry model was also used to demonstrate the feasibility of actively controlling the shaft position to suppress rotating stall. In axisymmetric mean flow, this method only stabilized the first harmonic mode, increasing the operating range until surge or higher harmonic modes became unstable.

Suder, Kenneth↗
NASA NTRS1997

Stator Indexing in Multistage Compressors

The relative circumferential location of stator rows (stator indexing) is an aspect of multistage compressor design that has not yet been explored for its potential impact on compressor aerodynamic performance. Although the inlet stages of multistage compressors usually have differing stator blade counts, the aft stages of core compressors can often have stage blocks with equal stator blade counts in successive stages. The potential impact of stator indexing is likely greatest in these stages. To assess the performance impact of stator indexing, researchers at the NASA Lewis Research Center used the 4 ft diameter, four-stage NASA Low Speed Axial Compressor for detailed experiments. This compressor has geometrically identical stages that can circumferentially index stator rows relative to each other in a controlled manner; thus it is an ideal test rig for such investigations.

Barankiewicz, Wendy S.↗
NASA NTRS1998

Jet Injection Used to Control Rotating Stall in a High-Speed Compressor

In a joint effort between the Massachusetts Institute of Technology (MIT) and the NASA Lewis Research Center, a new technology was demonstrated to identify and control rotating stall and surge in a single-stage, high-speed compressor. Through the use of highvelocity, high-frequency jet injectors, the instabilities of surge and stall were controlled in a high-speed compressor rig. Through the use of active stall control, modal instabilities that normally occur in the pressure measurements prior to stall were normalized and the range of the compressor was extended. Normally the events of rotating stall and surge instabilities limit the operation of the aeroengine compressor to a region below the surge line. To enhance the performance of the compressor, the Lewis/MIT team used active stall control methods to extend the normal operation of the compressor beyond the original stall point.

Bright, Michelle M.↗
NASA NTRS2000

New Compressor Added to Glenn's 450- psig Combustion Air System

In September 1999, the Central Process Systems Engineering Branch and the Maintenance and the Central Process Systems Operations Branch, released for service a new high pressure compressor to supplement the 450-psig Combustion Air System at the NASA Glenn Research Center at Lewis Field. The new compressor, designated C-18, is located in Glenn s Central Air Equipment Building and is remotely operated from the Central Control Building. C-18 can provide 40 pounds per second (pps) of airflow at pressure to our research customers. This capability augments our existing system capacity (compressors C 4 at 38 pps and C-5 at 32 pps), which is generated from Glenn's Engine Research Building. The C-18 compressor was originally part of Glenn's 21-Inch Hypersonic Tunnel, which was transferred from the Jet Propulsion Laboratory to Glenn in the mid-1980's. With the investment of construction of facilities funding, the compressor was modified, new mechanical and electrical support equipment were purchased, and the unit was installed in the basement of the Central Air Equipment Building. After several weeks of checkout and troubleshooting, the new compressor was ready for long-term, reliable operations. With a total of 110 pps in airflow now available, Glenn is well positioned to support the high-pressure air test requirements of our research customers.

Swan, Jeffrey A.↗
NASA NTRS1948

Investigation of X24C-2 10-Stage Axial-Flow Compressor. III - Surge Characteristics

Compressor operation at low air flows for a given speed is limited by unstable flow conditions, commonly called surge. An investigation of surge in centrifugal compressors (reference 1) showed that the pulsation of pressures and velocities occurred when the slope of the compressor characteristic curve was positive and that the magnitude and frequency, as well as the incidence of surge, depended on the capacity and resistance of the total system. Although the theory presented in reference 1 is applicable to axial-floe compressors, little experimental information is available on the surge characteristics of the individual stages of axial-flow compressors, or on the variation of the surge characteristics with operating conditions. During the investigation to determine the performance of the X24C-2 compressor (references 2 and 3), instrumentation was added to study the surge characteristics and to determine the effect of speed and inlet pressure on the frequency, amplitude, and phase relation of the pressure pulsations behind each stage.

Buckner, Howard A., Jr.↗
NASA NTRS2016

Effects of Double-Leakage Tip Clearance Flow on the Performance of a Compressor Stage with a Large Rotor Tip Gap

Effects of a large rotor tip gap on the performance of a one and half stage axial compressor are investigated in detail with a numerical simulation based on LES and available PIV data. The current paper studies the main flow physics, including why and how the loss generation is increased with the large rotor tip gap. The present study reveals that when the tip gap becomes large, tip clearance fluid goes over the tip clearance core vortex and enters into the next blade's tip gap, which is called double-leakage tip clearance flow. As the tip clearance flow enters into the adjacent blade's tip gap, a vortex rope with a lower pressure core is generated. This vortex rope breaks up the tip clearance core vortex of the adjacent blade, resulting in a large additional mixing. This double-leakage tip clearance flow occurs at all operating conditions, from design flow to near stall condition, with the large tip gap for the current compressor stage. The double-leakage tip clearance flow, its interaction with the tip clearance core vortex of the adjacent blade, and the resulting large mixing loss are the main flow mechanism of the large rotor tip gap in the compressor. When the tip clearance is smaller, flow near the end wall follows more closely with the main passage flow and this double-leakage tip clearance flow does not happen near the design flow condition for the current compressor stage. When the compressor with a large tip gap operates at near stall operation, a strong vortex rope is generated near the leading edge due to the double-leakage flow. Part of this vortex separates from the path of the tip clearance core vortex and travels from the suction side of the blade toward the pressure side of the blade. This vortex is generated periodically at near stall operation with a large tip gap. As the vortex travels from the suction side to the pressure side of the blade, a large fluctuation of local pressure forces blade vibration. Nonsynchronous blade vibration occurs due to this vortex as the frequency of this vortex generation is not the same as the rotor. The present investigation confirms that this vortex is a part of separated tip clearance vortex, which is caused by the double-leakage tip clearance flow.

flow↗
NASA NTRSDate not supplied

Flow Physics at a Large Rotor Tip Gap in a Multi-Stage Axial Compressor

The flow physics in a large rotor tip gap in a 1.5-stage axial compressor is investigated in the current study. The flow structure in the rotor tip region is complex with several dominant vortical structures of opposite rotation, resulting in inhomogeneous and highly anisotropic turbulence. Earlier measurements show that eddy viscosity is negative over large parts of the tip region and eddy viscosity varies among stress/strain components. The present study aims to understand how the complex nature of rotor tip leakage flow affects compressor performance when the tip gap size is greater than 4–5% of the rotor span, which is typical of advanced small core engines. Unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) techniques are applied to study flow physics in a large rotor tip gap (5.5% of rotor span) in a 1.5-stage axial compressor. Calculated flow fields from the two different approaches are compared with available measurements and examined in detail. LES calculates the pressure rise in the present compressor fairly well, while URANS with a standard two-equation turbulence closure underpredicts the pressure rise by 15-20% of the measured values. The current study shows that URANS with the current turbulence closure produces much higher all-positive eddy viscosity in the tip-gap region compared to measurements and LES. The distribution of eddy viscosity in the URANS simulation is also wrong. Consequently, the flow in the tip region is highly damped with significantly larger blockage generation, which results in the tip leakage vortex (TLV) staying closer to the blade suction side compared to the measurement. When the TLV stays closer to the blade, both flow turning and the pressure rise across the compressor are reduced compared to the measurements. It appears that this effect is amplified by a large rotor tip gap.

tip gap↗
DOE OSTI2021

Methane Exhaust Measurements at Gathering Compressor Stations in the United States

Unburned methane entrained in exhaust from natural gas-fired compressor engines (“combustion slip”) can account for a substantial portion of station-level methane emissions. A novel in-stack, tracer gas method was coupled with Fourier transform infrared (FTIR) species measurements to quantify combustion slip from natural gas compressor engines at 67 gathering and boosting stations owned or managed by nine “study partner” operators in 11 U.S. states. The mean methane emission rate from 63 four-stroke, lean-burn (4SLB) compressor engines was 5.62 kg/h (95% CI = 5.15–6.17 kg/h) and ranged from 0.3 to 12.6 kg/h. The mean methane emission rate from 39 fourstroke, rich-burn (4SRB) compressor engines was 0.40 kg/h (95% CI = 0.37– 0.42 kg/h) and ranged from 0.01 to 4.5 kg/h. Study results for 4SLB engines were lower than both the U.S. EPA compilation of air pollutant emission factors (AP-42) and Inventory of U.S. Greenhouse Gas Emissions and Sinks (GHGI) by 8 and 9%, respectively. Study results for 4SRB engines were 43% of the AP-42 emission factor and 8% of the GHGI emission factor, the latter of which does not distinguish between engine types. Total annual combustion slip from the U.S. natural gas gathering and boosting sector was modeled using measured emission rates and compressor unit counts from the U.S. EPA Greenhouse Gas Reporting Program. Modeled results [328 Gg/y (95% CI = 235–436 Gg/y) of unburned methane] would account for 24% (95% CI = 17–31%) of the 1391 Gg of methane emissions for “Gathering and Boosting Stations”, or 6% of the net emissions for “Natural Gas Systems” (5598 Gg) as reported in the 2020 U.S. EPA GHGI. In conclusion, gathering and boosting combustion slip emissions reported in the 2020 GHGI (374 Gg) fall within the uncertainty of this model.

37 INORGANIC, ORGANIC, PHYSICAL, AND ANALYTICAL CH↗
DOE OSTI2022

SZ3: A Modular Framework for Composing Prediction-Based Error-Bounded Lossy Compressors

Today's scientific simulations require a significant reduction of data volume because of extremely large amounts of data they produce and the limited I/O bandwidth and storage space. Here, error-bounded lossy compression has been considered one of the most effective solutions to the above problem. In practice, however, the best-fit compression method often needs to be customized or optimized in particular because of diverse characteristics in different datasets and various user requirements on the compression quality and performance. In this paper, we address this issue with a novel modular, composable compression framework named SZ3. Our contributions are four-folds. (1) We develop SZ3 which features an innovative modular abstraction for the prediction-based compression framework, such that compression modules can be plugged in easily to create new compressors based on characteristics of data and user requirements. (2) We create a new compression pipeline by SZ3 for GAMESS data, which significantly improves the compression ratios over state-of-the-art compressors. (3) We develop an adaptive compression pipeline by SZ3 for APS data with minimal efforts, which leads to the best rate-distortion among all existing error-bounded lossy compressors for any bit-rate. (4) We compare the sustainability of SZ3 with leading error-bounded prediction-based compressors, and then demonstrate the necessity of diverse pipelines by integrating and evaluating several compression pipelines on diverse scientific datasets from multiple disciplines. Experiments show that SZ3 incurs very limited overhead in compressor integration and our customized compression pipelines lead to up to 20% improvement in compression ratios under the same data distortion, when compared with the best existing approach.

97 MATHEMATICS AND COMPUTING↗
DOE OSTI2023

Building Technologies Office 03.02.02.38 Milestone Report—Compressor Technology Development for Refrigerants with <150 Global Warming Potential

The major purpose of this project is to enable the development of compressors specifically designed for compatibility with low–global warming potential (GWP) refrigerants (with GWP values below 150). These compressors are intended for applications in residential and light commercial air-conditioning and heat pump systems. This report contains an exhaustive review of the existing literature with the intention of discerning the most efficacious methodologies currently in use for compressor modeling. This fundamental stage was designed to foster a comprehensive understanding of the intricate dynamics and mechanics of compressors. As a result, a model (PDSim) was selected to simulate scroll, rotary and linear compressor technologies.

42 ENGINEERING↗
NASA NTRS1972

Evaluation of circumferential airflow uniformity entering combustors from compressors. Volume 1: Discussion of data and results

The compressor discharge airflow uniformity of two compressors from advanced engines, the J58 and F100/F401, was studied. Compressor discharge pressures and temperatures at up to 33 circumferential rake locations allowed the airflow distribution to be ascertained and computer plotted. Several flight conditions and compressor variables, i.e., inlet distortion, modified seals, etc., were analyzed. An unexpectedly high nonuniform airflow was found for both compressors. Circumferential airflow deviation differences of up to 52% from maximum to minimum were found for the J58, and up to 40% for the F100/F401. The effects of aerodynamic and thermal distortion were found to be additive. The data were analyzed for influence of exit guide vane wakes and found free of any effect. Data system errors were small in relation to the measured pressure and temperature variations.

Shadowen, J. H.↗
NASA NTRS1974

Formulation of a distortion index based on peak compressor pressure ratios

In order to effectively use a compressor face total-pressure distortion index as a measure of inlet-engine compatibility, a correlation of distortion amplitude with stall margin must be developed with minimal scatter. A recent analysis of data recorded in extensive distortion screen tests with the J85-GE-13 turbojet engine has resulted in a correlation based on compressor discharge pressure ratioed to the minimum pressure at the compressor face. Simply by determining compressor stall lines with a single hub radial distortion pattern, a single tip radial pattern, and with undistorted inflow, the overall compressor pressure ratio at stall for even the most complex distortion pattern was found to be predictable.

Calogeras, J. E.↗