SEARCH · Search NASA

Results for “Compressor”

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.

SourceResearch connectionFrom yearThrough year

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

At least 109 records · Page 6

All sources NASA NTRS DOE OSTI

NASA NTRS1984

Determination of compressor in-stall characteristics from engine surge transients

A technique for extracting the in-stall pumping characteristics for an axial flow compressor operating in an engine system environment is developed. The technique utilizes a Hybrid computer simulation of the compressor momentum equation into which actual transient data are used to provide all terms but the desired compressor characteristic. The compressor force characteristic as a function of corrected flow and speed result from the computation. The critical problem of data filtering is addressed. Results for a compressor operating in a turbofan engine are presented and comparison is made with the conventional compressor map. The relationship of the compressor surge characteristic with its rotating stall characteristic is explored. Initial interpretation of the measured results is presented.

Lorenzo, C. F.↗

NASA NTRS1984

Determination of compressor in-stall characteristics from engine surge transients

A technique for extracting the in-stall pumping characteristics for an axial flow compressor operating in an engine system environment is developed. The technique utilzes a Hybrid computer simulation of the compressor momentum equation in which actual transient data are used to provide all terms but the desired compressor characteristic. The compressor force characteristic as a function of corrected flow and speed results from the computation. The critical problem of data filtering is addressed. Results for a compressor operating in a turbofan engine are presented and comparison is made with the conventional compressor map. The relationship of the compressor surge charcteristic with its rotating stall characteristic is explored. Initial interpretation of the measured results is presented. Previously announced in STAR as N84-22566

Lorenzo, C. F.↗

NASA NTRS1989

Small Displacement, Long Life On-Orbit Compressor Design and Fabrication

The focus is the generation of technology and fabrication of prototype hardware applicable to seven Space Station compressor system applications. The compressors are of the single acting reciprocating piston type and, in general, may be termed miniature in size compared with normal commercially available equipment. The initial technology development is focused on improved valve designs, and the control of pulsations and heating effects in order to increase compressor efficiency and reduce cycle temperatures, thus permitting significantly increased stage pressure ratios. The initial test compressor was successfully operated at pressure ratios of up to 50:1, and this significant extension of allowable pressure ratio will result in a reduction of the number of required stages and, hence, total hardware thereby reducing system weight and volume. These experiments have also identified the need to employ low shaft speeds, on the order of 250 to 500 rpm, to enhance heat transfer and increase life. The prototype compressor currently being designed, is to be driven by a low-speed brushless dc motor sealed in a case common to the compressor drive mechanism case. The compressor and motor case will communicate with stage suction pressure so that any minor gas leakage past the piston rings will be returned to the suction. Emphasis in this prototype design is being placed on simplicity, durability, commonality of components, and high efficiency.

Gerlach, C. R.↗

NASA NTRS1996

Dimension Determination of Precursive Stall Events in a Single Stage High Speed Compressor

This paper presents a study of the dynamics for a single-stage, axial-flow, high speed compressor core, specifically, the NASA Lewis rotor stage 37. Due to the overall blading design for this advanced core compressor, each stage has considerable tip loading and higher speed than most compressor designs, thus, the compressor operates closer to the stall margin. The onset of rotating stall is explained as bifurcations in the dynamics of axial compressors. Data taken from the compressor during a rotating stall event is analyzed. Through the use of a box-assisted correlation dimension methodology, the attractor dimension is determined during the bifurcations leading to rotating stall. The intent of this study is to examine the behavior of precursive stall events so as to predict the entrance into rotating stall. This information may provide a better means to identify, avoid or control the undesirable event of rotating stall formation in high speed compressor cores.

Bright, Michelle M.↗

NASA NTRS2002

Centrifugal Compressor Aeroelastic Analysis Code

Centrifugal compressors are very widely used in the turbomachine industry where low mass flow rates are required. Gas turbine engines for tanks, rotorcraft and small jets rely extensively on centrifugal compressors for rugged and compact design. These compressors experience problems related with unsteadiness of flowfields, such as stall flutter, separation at the trailing edge over diffuser guide vanes, tip vortex unsteadiness, etc., leading to rotating stall and surge. Considerable interest exists in small gas turbine engine manufacturers to understand and eventually eliminate the problems related to centrifugal compressors. The geometric complexity of centrifugal compressor blades and the twisting of the blade passages makes the linear methods inapplicable. Advanced computational fluid dynamics (CFD) methods are needed for accurate unsteady aerodynamic and aeroelastic analysis of centrifugal compressors. Most of the current day industrial turbomachines and small aircraft engines are designed with a centrifugal compressor. With such a large customer base and NASA Glenn Research Center being, the lead center for turbomachines, it is important that adequate emphasis be placed on this area as well. Currently, this activity is not supported under any project at NASA Glenn.

Keith, Theo G., Jr.↗

NASA NTRS2004

Experimental Investigation of Diffuser Hub Injection to Improve Centrifugal Compressor Stability

Results from a series of experiments to investigate whether centrifugal compressor stability could be improved by injecting air through the diffuser hub surface are reported. The research was conducted in a 4:1 pressure ratio centrifugal compressor configured with a vane-island diffuser. Injector nozzles were located just upstream of the leading edge of the diffuser vanes. Nozzle orientations were set to produce injected streams angled at 8, 0 and +8 degrees relative to the vane mean camber line. Several injection flow rates were tested using both an external air supply and recirculation from the diffuser exit. Compressor flow range did not improve at any injection flow rate that was tested. Compressor flow range did improve slightly at zero injection due to the flow resistance created by injector openings on the hub surface. Leading edge loading and semi-vaneless space diffusion showed trends similar to those reported earlier from shroud surface experiments that did improve compressor flow range. Opposite trends are seen for hub injection cases where compressor flow range decreased. The hub injection data further explain the range improvement provided by shroud-side injection and suggest that different hub-side techniques may produce range improvement in centrifugal compressors.

Skoch, Gary J.↗

NASA NTRS1948

Effect of Inlet Air Distortion on the Steady-State and Surge Characteristics of an Axial-Flow Turbojet Compressor

An investigation was conducted in an altitude test chamber to determine the effects of inlet airflow distortion on the compressor steady-state and surge characteristics of a high-pressure ratio, axial-flow turbojet engine. Circumferential-type inlet flow distortions were investigated, which covered a range of distortion sector angles from 20 deg to 168 deg and distortion levels up to 22 percent. The presence of inlet airflow distortions at the compressor face resulted in a substantial increase in the local pressure ratio in the distorted region, primarily for the inlet stages. The local pressure ratio in the distorted region for the inlet stages increased as either the distortion sector angle decreased or the percent distortion increased. The average compressor-surge pressure ratio was much more sensitive to inlet airflow distortions at lower engine speeds than at engine speeds near rated. Hence, compressor-surge margin reduction due to inlet airflow distortion was quite severe at the lower engine speeds. Although the average compressor-surge pressure ratio was generally reduced with inlet flow distortion, local pressure ratios across the distorted sector of the compressor were obtained during surge and were significantly greater than the normal compressor-surge pressure ratio. This was a result of increased loading of the inlet stages in the distorted region.

Ciepluch, Carl C.↗

NASA NTRS2002

Centrifugal Compressor Surge Margin Improved With Diffuser Hub Surface Air Injection

Aerodynamic stability is an important parameter in the design of compressors for aircraft gas turbine engines. Compression system instabilities can cause compressor surge, which may lead to the loss of an aircraft. As a result, engine designers include a margin of safety between the operating line of the engine and the stability limit line of the compressor. The margin of safety is typically referred to as "surge margin." Achieving the highest possible level of surge margin while meeting design point performance objectives is the goal of the compressor designer. However, performance goals often must be compromised in order to achieve adequate levels of surge margin. Techniques to improve surge margin will permit more aggressive compressor designs. Centrifugal compressor surge margin improvement was demonstrated at the NASA Glenn Research Center by injecting air into the vaned diffuser of a 4:1-pressure-ratio centrifugal compressor. Tests were performed using injector nozzles located on the diffuser hub surface of a vane-island diffuser in the vaneless region between the impeller trailing edge and the diffuser-vane leading edge. The nozzle flow path and discharge shape were designed to produce an air stream that remained tangent to the hub surface as it traveled into the diffuser passage. Injector nozzles were located near the leading edge of 23 of the 24 diffuser vanes. One passage did not contain an injector so that instrumentation located in that passage would be preserved. Several orientations of the injected stream relative to the diffuser vane leading edge were tested over a range of injected flow rates. Only steady flow (nonpulsed) air injection was tested. At 100 percent of the design speed, a 15-percent improvement in the baseline surge margin was achieved with a nozzle orientation that produced a jet that was bisected by the diffuser vane leading edge. Other orientations also improved the baseline surge margin. Tests were conducted at speeds below the design speed, and similar results were obtained. In most cases, the greatest improvement in surge margin occurred at fairly low levels of injected flow rate. Externally supplied injection air was used in these experiments. However, the injected flow rates that provided the greatest benefit could be produced using injection air that is recirculating between the diffuser discharge and nozzles located in the diffuser vaneless region. Future experiments will evaluate the effectiveness of recirculating air injection.

Skoch, Gary J.↗

NASA NTRS1955

Preliminary Results of the Determination of Inlet-Pressure Distortion Effects on Compressor Stall and Altitude Operating Limits of the J57-P-1 Turbojet Engine

During an investigation of the J57-P-1 turbojet engine in the Lewis altitude wind tunnel, effects of inlet-flow distortion on engine stall characteristics and operating limits were determined. In addition to a uniform inlet-flow profile, the inlet-pressure distortions imposed included two radial, two circumferential, and one combined radial-circumferential profile. Data were obtained over a range of compressor speeds at an altitude of 50,000 and a flight Mach number of 0.8; in addition, the high- and low-speed engine operating limits were investigated up to the maximum operable altitude. The effect of changing the compressor bleed position on the stall and operating limits was determined for one of the inlet distortions. The circumferential distortions lowered the compressor stall pressure ratios; this resulted in less fuel-flow margin between steady-state operation and compressor stall. Consequently, the altitude operating Limits with circumferential distortions were reduced compared with the uniform inlet profile. Radial inlet-pressure distortions increased the pressure ratio required for compressor stall over that obtained with uniform inlet flow; this resulted in higher altitude operating limits. Likewise, the stall-limit fuel flows required with the radial inlet-pressure distortions were considerably higher than those obtained with the uniform inlet-pressure profile. A combined radial-circumferential inlet distortion had effects on the engine similar to the circumferential distortion. Bleeding air between the two compressors eliminated the low-speed stall limit and thus permitted higher altitude operation than was possible without compressor bleed.

Wallner, L. E.↗

NASA NTRS2011

Modeling the Effects of Ice Accretion on the Low Pressure Compressor and the Overall Turbofan Engine System Performance

The focus of this study is on utilizing a mean line compressor flow analysis code coupled to an engine system thermodynamic code, to estimate the effects of ice accretion on the low pressure compressor, and quantifying its effects on the engine system throughout a notional flight trajectory. In this paper a temperature range in which engine icing would occur was assumed. This provided a mechanism to locate potential component icing sites and allow the computational tools to add blockages due to ice accretion in a parametric fashion. Ultimately the location and level of blockage due to icing would be provided by an ice accretion code. To proceed, an engine system modeling code and a mean line compressor flow analysis code were utilized to calculate the flow conditions in the fan-core and low pressure compressor and to identify potential locations within the compressor where ice may accrete. In this study, an "additional blockage" due to the accretion of ice on the metal surfaces, has been added to the baseline aerodynamic blockage due to boundary layer, as well as the blade metal blockage. Once the potential locations of ice accretion are identified, the levels of additional blockage due to accretion were parametrically varied to estimate the effects on the low pressure compressor blade row performance operating within the engine system environment. This study includes detailed analysis of compressor and engine performance during cruise and descent operating conditions at several altitudes within the notional flight trajectory. The purpose of this effort is to develop the computer codes to provide a predictive capability to forecast the onset of engine icing events, such that they could ultimately help in the avoidance of these events.

Veres, Joseph P.↗

NASA NTRS2015

An Experimental Investigation of the Flow Physics Associated With End Wall Losses and Large Rotor Tip Clearances as Found in the Rear Stages of a High Pressure Compressor

The focus of this work 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 clearance heights were studied with nominal tip clearance heights of 1.5%, 3.0%, and 4.0% based on a constant annulus height. Overall compressor performance was investigated at four corrected speedlines (100%, 90%, 80%, and 68%) for each of the three tip clearance configurations using total pressure and total temperature rakes distributed throughout the compressor. The results have confirmed results from previous authors showing a decrease of total pressure rise, isentropic efficiency, and stall margin which is approximately linear with increasing tip clearance height. The stall inception mechanisms have also been evaluated at the same corrected speeds for each of the tip clearance configurations. Detailed flow field measurements have been collected at two loading conditions, nominal loading (NL) and high loading (HL), on the 100% corrected speedline for the smallest and largest tip clearance heights (1.5% and 4.0%). Steady detailed radial traverses of total pressure at the exit of each stator row have been 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 have helped to quantify the size of the leakage flow at the exit of each rotor. Thermal anemometry has also been implemented to evaluate the time-resolved three-dimensional components of velocity throughout the compressor and calculate blockage due to the rotor tip leakage flow throughout the compressor. These measurements have also been used to calculate streamwise vorticity. 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% corrected speedline using fast-response piezoresistive pressure sensors. These time-resolved static pressure measurements, as well as the time-resolved total pressures and velocities have helped to reveal a profound influence of the upstream stator vane on the size and shape of the rotor tip leakage flow. Finally, a novel particle image velocimetry (PIV) technique has been developed as a proof-of- concept. In contrast to PIV methods that have been typically been utilized for turbomachinery applications in the past, the method used for this study introduced the laser light through the same access window that was also used to image the flow. This new method addresses potential concerns related to the intrusive laser-introducing techniques that have typically been utilized by other authors in the past. 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 NTRS1952

Preliminary Investigation of Flow Fluctuations During Surge and Blade Row Stall in Axial-flow Compressors / c Merle C. Huppert

A preliminary investigation of the flow fluctuations of surge and blade row stall was conducted with three single-stage axial-flow compressors with hub-tip ratios of 0.9,0.8, and 0.5 and with a multistage axial-flow compressor. Flow fluctuations of large amplitude associated with stall were detected in all compressors investigated. The fluctuations were caused by low flow regions affecting 25 to 40 percent of the annulus area and propagating in the direction of compressor rotation, but at a lower speed. Mild audible surge was obtained with the single-stage compressors with hub-tip ratios of 0.9 and 0.8. No audible surge was detected with the 0.5 hub-tip ratio stage. A severe audible surge was obtained with the multistage axial-flow compressor. The surge point was coincident with the sharp drop in pressure ratio due to stall.

ENGINES, TURBO-JET↗

NASA NTRS2015

Stage-by-Stage and Parallel Flow Path Compressor Modeling for a Variable Cycle Engine

Parallel Flow Path modeling↗

NASA NTRS2015

Stage-by-Stage and Parallel Flow Path Compressor Modeling for a Variable Cycle Engine, NASA Advanced Air Vehicles Program - Commercial Supersonic Technology Project - AeroServoElasticity

This paper covers the development of stage-by-stage and parallel flow path compressor modeling approaches for a Variable Cycle Engine. The stage-by-stage compressor modeling approach is an extension of a technique for lumped volume dynamics and performance characteristic modeling. It was developed to improve the accuracy of axial compressor dynamics over lumped volume dynamics modeling. The stage-by-stage compressor model presented here is formulated into a parallel flow path model that includes both axial and rotational dynamics. This is done to enable the study of compressor and propulsion system dynamic performance under flow distortion conditions. The approaches utilized here are generic and should be applicable for the modeling of any axial flow compressor design accurate time domain simulations. The objective of this work is as follows. Given the parameters describing the conditions of atmospheric disturbances, and utilizing the derived formulations, directly compute the transfer function poles and zeros describing these disturbances for acoustic velocity, temperature, pressure, and density. Time domain simulations of representative atmospheric turbulence can then be developed by utilizing these computed transfer functions together with the disturbance frequencies of interest.

Compressor Modeling Modeling↗

NASA NTRSDate not supplied

GT2024-139262: Development of a Computational Fluid Dynamics Model for a High-Speed Centrifugal Compressor

Computational fluid dynamics (CFD) has become widely used in the design and analysis of turbomachinery components such as centrifugal compressors. However, CFD is only a limited representation of experimental cases and struggles to model complex flows or lack model details to increase computation speed. The High Efficiency Centrifugal Compressor (HECC) was designed by United Technologies Research Center for NASA to investigate the difficulties behind improving centrifugal compressor technology and to provide an open case to the turbomachinery research community. The CFD technology seen in previous numerical models of the compressor have improved since then. In this study, a high-fidelity computational model for the vaneless configuration of HECC operating at its design point was developed and compared to an equivalent experimental case. This model incorporated GPU versions of mesh generation code WAND and solver code LEO from AeroDynamic Solutions, Inc. Blade sections extracted from the solid model formed the basis of the impeller mesh, while the remainder of the stage was modeled as ducts. Standard day aerodynamic conditions, design rotational speed, design mass flow rate, and inlet flow angles served as additional inputs. A mesh sensitivity study was performed prior to using the model for analysis. The accuracy of the model’s predictions of impeller and stage performance were assessed by comparing the total pressure ratio, total temperature rise ratio, and adiabatic efficiency to the corresponding parameters observed in the experimental case. Results showed a general overprediction of compressor performance by the model. Predictions were more accurate for the impeller performance compared to those for the full compressor stage. Spanwise analyses showed that the model tended to follow flow behavior patterns exhibited in the experiment. Beyond this study, this model was used to examine HECC aerodynamics and performance in research following thereafter.

centrifugal compressor↗

NASA NTRS2014

ERA Advanced Core Compressor Technology Program

As part of the NASA Integrated Systems Research Program, Environmentally Responsible Aviation Project, NASA GRC and GE Aviation have partnered to investigate the technology barriers associated with improved fuel economy of large gas turbine engines. Crucial to improving fuel economy is the increase in Pressure Ratio (PR) of the core compressor. General Electric is in the process of defining new advanced technology engine concepts and related engine architectures for both commercial and military subsonic and supersonic flight applications. These new concepts and architectures offer the potential of providing substantial improvements in mission fuel burn, emissions, and noise levels relative to SOTA engines and support the needs and goals of future Air force, Navy, and NASA aircraft systems. A key aspect of many of these new engine concepts is an ultra high pressure ratio core compressor. Current SOTA engines such as the GENx utilize a 23:1 class pressure ratio core compressor to achieve an overall pressure ratio in the 50:1 class. Future engines with overall pressure ratios in the 70– 80:1 class will need core compressors with pressure ratios in the 30:1 class. These single spool, ultra high pressure ratio compressors will have inherent challenges. The goal of this work is to provide a dedicated component test vehicle to study the very challenging compressor front stage performance and operability issues that arise in these aggressive machines. With a view to providing highly detailed aerodynamic data that will provide, subject to the rights restrictions for test data afforded under NASA Contract NNC10BA06B/NNC10TA64T, an industry standard data set for CFD validation.

Axial Compressors↗

NASA NTRS1974

Frequency response of an axial-flow compressor exposed to inlet pressure perturbations

Experimental results of a series of engine tests designed to obtain the stage dynamics of an eight-stage axial-flow compressor over the frequency range of 0.5 to 200 hertz are presented. The total pressure at the compressor face was varied by means of a secondary air jet system installed in the engine inlet and positioned to oppose the primary airflow. Total-pressure probes located at each compressor stage were used to obtain the frequency response of each compressor-stage total pressure to the average compressor-inlet total pressure. The engine operating conditions were chosen to illustrate the effects of changing the rotor speed, changing the exhaust nozzle area, and isolating the compressor discharge pressure perturbations from the fuel control and hence, the fuel flow.

Milner, E. J.↗

NASA NTRS1975

Effect of casing treatment on performance of a multistage compressor

AJ85-GE-13 engine was equipped with a compressor case which allowed changes to the case wall over the rotor tips of six of its stages. The engine was tested with four inlet configurations: undistorted and with 180 deg circumferential, hub radial, and tip radial distortions. Baseline data defining compressor performance and stall regions were taken for these inlet configurations with solid (untreated) compressor case inserts. Circumferentially grooved inserts were installed in the first three and last three stages, and the compressor was mapped under similar conditions. The compressor was mapped a third time with untreated inserts in the first three stages and inserts having slots conforming to blade angles in the last three stages. In most cases, the stall pressure ratio was the same as or lower than the baseline. Pumping capacity with the slotted inserts was reduced. Overall compressor efficiency with the grooved rings installed did not appreciably differ from the baseline, but it was 1 to 2 percentage points lower than the baseline with the slotted rings in place. Average stage characteristics for the undistorted inlet case showed little or no sensitivity to casing treatment.

Wenzel, L. M.↗