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At least 163 records · Page 9

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NASA NTRS1975

Performance of a J85-13 compressor with clean and distorted inlet flow

The results presented are of a series of experimental tests in which a J85-13 turbojet engine was subjected to both distorted and undistorted inlet total pressure conditions. A distinctive feature of the data base obtained is that it includes compressor interstage information not previously recorded for a J85-13 engine. Each of the eight compressor stages was instrumented to obtain the characteristics of the individual stages for undistorted inlet conditions, and these data are documented in the report along with the undistorted compressor overall performance. Also included in the report is the overall performance of the compressor exposed to 14 different distorted-inlet conditions - 10 circumferential patterns and 4 radial patterns. The distortion patterns were introduced using screens that spoiled from 8 to 50 percent of the compressor face area; the distortion screen density, or the area blocked by the screen wire per unit area of screen, varied from 26 to 69 percent.

Milner, E. J.↗

NASA NTRS1976

Effect of slotted casing treatment on performance of a multistage compressor

A J85-13 engine was equipped with a compressor case that allowed changes to the case wall over the rotor tips of six of its eight stages. The engine was tested with four inlet configurations: uniform inlet flow, 180 degree circumferential distortion, hub radial distortion, and tip radial distortion. Slotted inserts were installed in the first three stages, and the compressor was mapped under similar conditions. Overall compressor performance obtained with tip treatment was inferior to the performance for the compressor's normal operating range. Pumping capacity with the slotted inserts was reduced. Overall compressor efficiency was reduced 1 to 2 percent with the slotted rings installed for 90 and 100 percent corrected engine speeds.

Moss, J. E., Jr.↗

NASA NTRS1976

Single-stage experimental evaluation of low aspect ratio, highly loaded blading for compressors. Part 9: Stage F and stage G, volume 1

Two single-stage, 0.77 hub/tip ratio axial-flow compressors were tested to evaluate the effectiveness of low aspect ratio blading as a means of obtaining higher stage loadings. One compressor, designated Stage F, was comprised of circular arc blading with an aspect ratio of 0.9 for both the rotor and stator. This compressor was tested with uniform inlet flow, hub radial, tip radial, and 180 deg arc circumferential inlet distortion. The second compressor, designated Stage G, was comprised of multiple circular arc blading with an aspect ratio of 1.0 for both the rotor and stator. This compressor was tested with uniform inlet flow only. Design rotor tip speeds for Rotor F and Rotor G were 285 m/sec (934 ft/sec) and 327 m/sec (1,074 ft/sec) respectively. Both stages operated at high loading levels with adequate efficiency and operating range. The peak efficiencies and corresponding average stage diffusion factors for Stages F and G at design rotor speed were 86.4% and 84.1% and 0.59 and 0.55 respectively. The surge margin at peak efficiency for Stage F was 12.6% and the corresponding value for Stage G was 16.5%. Both stages experienced a loss in efficiency with increasing rotor speed; however, the multiple circular arc rotor delayed the characteristic loss in efficiency within increasing Mach number to higher Mach number.

Cheatham, J. G.↗

NASA NTRS1977

Internal flow characteristics of a multistage compressor with inlet pressure distortion

The measured distribution of compressor interstage pressures and temperatures resulting from a 180 deg inlet-total-pressure distortion for a J85-13 turbojet engine is reported. Extensive inner stage instrumentation combined with stepwise rotation of the inlet distortion gave data of high circumferential resolution. The steady-state pressures and temperatures along with the amplitude, extent, and location of the distorted areas are given. Data for 80, 90, and 100 percent of rotor design speed are compared with clean (undistorted) inlet flow conditions to show pressure and temperature behavior within the compressor. Both overall and stagewise compressor performances vary only slightly when clean and distorted inlet conditions are compared. Total and static pressure distortions increase in amplitude in the first few stages of the compressor and then attenuate fairly uniformly to zero at the discharge. Total-temperature distortion induced by the pressure distortion reached a maximum amplitude by the first two stages and decayed only a little through the rest of the compressor. Distortion amplitude tended to peak in line with the screen edges, and, except for total and static pressure in the tip zone, there was little swirl in the axial direction.

Debogdan, C. E.↗

NASA NTRS1979

Causes of high pressure compressor deterioration in service

The high mechanical reliability and low deterioration rate of the JT9D high-pressure compressor results in long utilization without exceeding engine operational limits. The increasing cost and decreasing supply of fuel have focused attention on the fuel burned implications of such high time use without refurbishment. The paper presents the results of JT9D high pressure compressor studies. The mechanical deterioration of the JT9D high-pressure compressor gaspath parts versus increasing service usage, documented from inspection of service parts, is presented and discussed including changes in airfoil roughness, blade length, airfoil contour and outer air seal trench characteristics. An estimate of the performance loss versus usage is related to each type of damage. The combined estimated high-pressure compressor performance loss for all mechanisms determined from part inspection is compared to historical engine test data to establish the validity of the predicted loss levels. The effect of cold section refurbishment on engine fuel consumption recovery and the results of an optimization study to determine the appropriate interval for high pressure compressor refurbishment are also reported.

Richardson, J. H.↗

NASA NTRS1980

Steady-state performance of J85-21 compressor at 100 percent of design speed with and without interstage rake blockage

Internal compressor instrumentation blockage effects on steady state J85-21 compressor performance at 100 percent of design speed are determined. The blockage was generated by instrumented vanes for the first three compressor stages and by removal rakes for stages 4 to 9. Individual flow passage blockages ranged up to 4.5 percent with the instrumented vanes and up to 22 percent with the removable interstage rakes. At a Reynolds number index of 1.0, pressure ratio and airflow remained unchanged with insertion of the interstage rakes, but efficiency dropped 0.3 percentage point. Compressor exit profiles, compressor stage static pressure rise coefficients, turbine exit temperature, and fuel flow are also presented.

Werner, R. A.↗

NASA NTRS1980

Practical experience with unstable compressors

Using analytical mathematical modeling techniques for the system components, an attempt is made to gauge the destabilizing effects in a number of compressor designs. In particular the overhung (or cantilevered) compressor designs and the straddle-mounted (or simply supported) compressor designs are examined. Recommendations are made, based on experiences with stable and unstable compressors, which can be used as guides in future designs. High and low pressure compressors which operate well above their fundamental rotor-bearing lateral natural frequencies can suffer from destructive subsynchronous vibration. Usually the elements in the system design which contribute to this vibration, other than the shafting and the bearings, are the seals (both gas labyrinth and oil breakdown bushings) and the aerodynamic components.

Malanoski, S. B.↗

NASA NTRS1982

CF6 Jet Engine Diagnostics Program: High pressure compressor clearance investigation

The effects of high pressure compressor clearance changes on engine performance were experimentally determined on a CF6 core engine. The results indicate that a one percent reduction in normalized average clearance, expressed as a fraction of airfoil length, improves compressor efficiency by one percent. Compressor clearances are reduced by the application of rotor bore cooling, insulation of the stator casing, and use of a low coefficient of expansion material in the aft stages. This improvement amounts to a reduction of normalized average clearance of 0.78 percent, relative to CF6-60 compressor, which is equivalent to an improvement in compressor efficiency of 0.78 percent.

Radomski, M. A.↗

NASA NTRS1983

Determining compressor-inlet airflow in the compact multimission aircraft propulsion simulators in wind-tunnel applications

Test techniques are being developed to use the Compact Multimission Aircraft Propulsion Simulator (CMAPS) in wind-tunnel tests. The CMAPS allows simultaneous simulation of inlet and exhaust flow fields. Static tests have been conducted to acquire CMAPS performance data and to evaluate four methods of determining compressor inlet airflow during wind-tunnel tests. The first method measures airflow at the compressor face; the second measures airflow in the compressor discharge duct; the third deduces the compressor airflow from the calibrated nozzle, turbine supply, and turbine bleed airflows; and the fourth correlates compressor airflow with the mixer pressure. Test results and the advantages, disadvantages, and expected accuracy of each method are discussed.

Smith, S. C.↗

NASA NTRS1985

The design of a small linear-resonant, split Stirling cryogenic refrigerator compressor

The development of a small linear-resonant compressor for use in a 1/4-watt, 78K, split Stirling cryogenic refrigerator is discussed. The compressor contains the following special features: (1) a permanent-magnet linear motor; (2) resonant dynamics; (3) dynamic balancing; and (4) a close-clearance seal between the compressor piston and cylinder. This paper describes the design of the compressor, and presents component test data and system test data for the compressor driving a 1/4-watt expander.

Ackermann, R. A.↗

NASA NTRS1987

Performance of two 10-lb/sec centrifugal compressors with different blade and shroud thicknesses operating over a range of Reynolds numbers

Centrifugal compressors often cannot be directly scaled to very small flow sizes because of structural and manufacturing limitations. The inability to directly scale all design parameters leads to a performance loss other than that which can be associated with the lower Reynolds number. A 10-lb/sec centrifugal compressor was scaled down to 2-lb/sec where adjustments to blade and shroud thickness and fillet radii were required. The modified 2-lb/sec compressor was then directly scaled back up to 10 lb/sec so that the effect of the modifications could be determined. The performance of the two 10-lb/sec compressors is compared over a range of speed and mass flow. The effect of variations in Reynolds number, impeller tip clearance, and shroud thickness on compressor performance is also presented.

Skoch, Gary J.↗

NASA NTRS1987

Performance of two 10-lb/sec centrifugal compressors with different blade and shroud thicknesses operating over a range of Reynolds numbers

Skoch, Gary J.↗

NASA NTRS1951

Analysis of thrust augmentation of turbojet engines by water injection at compressor inlet including charts for calculating compression processes with water injection

A psychrometric chart having total pressure (sum of partial pressures of air and water vapor) as a variable, a Mollier diagram for air saturated with water vapor, and charts showing the thermodynamic properties of various air-water vapor and exhaust gas-water vapor mixtures are presented as aids in calculating the thrust augmentation of a turbojet engine resulting from the injection of water at the compressor inlet. Curves are presented that show the theoretical performance of the augmentation method for various amounts of water injected and the effects of varying flight Mach number, altitude, ambient-air temperature, ambient relative humidity, compressor pressure ratio, and inlet-diffuser efficiency. Numerical examples, illustrating the use of the psychrometric chart and the Mollier diagram in calculating both compressor-inlet and compressor-outlet conditions when water is injected at the compressor inlet, are presented.

Wilcox, E Clinton↗

NASA NTRS1948

Altitude-Wind-Tunnel investigation of Westinghouse 19B-2, 19B-8, and 19XB-1 Jet-Propulsion Engines IV : analysis of compressor performance

Investigations were conducted in the NACA Cleveland altitude wind tunnel to determine the performance and operational characteristics of the 19B-2, 19B-6, and 19XB-1 Turbojet Engines. One objective of the investigations was to determine the effect of altitude, flight Mach number, and tail-pipe-nozzle area on the performance characteristics of the six-stage and ten-stage axial-flow compressors of the 19B-8 and 19XB-1 engines, respectively. The data were obtained over a range of simulated altitudes and flight Mach numbers. At each simulated flight condition the engine was run over its full operable range of speeds. Performance characteristics of the 19B-8 and 19XB-1 compressors for the range of operation obtainable in the turbojet-engine installation are presented. Compressor characteristics are presented as functions of air flow corrected to sea-level conditions, compressor Mach number, and compressor load coefficient.

Dietz, Robert O↗

NASA NTRS2000

The MEMS Knudsen Compressor as a Vacuum Pump for Space Exploration Applications

Several lander, probe and rover missions currently under study at the Jet Propulsion Laboratory (JPL) and especially in the Microdevices Laboratory (MDL) Center for Space Microelectronics Technology, focus on utilizing microelectromechanical systems (MEMS) based instruments for science data gathering. These small instruments and NASA's commitment to "faster, better, cheaper" type missions has brought about the need for novel approaches to satisfying mission requirements. Existing in-situ instrument systems clearly lack novel and integrated methods for satisfying their vacuum needs. One attractive candidate for a MEMS vacuum pump is the Knudsen Compressor, which operates based on thermal transpiration. Thermal transpiration describes gas flows induced by temperature differences maintained across orifices, porous membranes or capillary tubes under rarefied conditions. This device has two overwhelmingly attractive features as a MEMS vacuum pump - no moving parts and no fluids. An initial estimate of a Knudsen Compressor's pumping power requirements for a surface atmospheric sampling task on Mars is less than 80 mW, significantly below than alternative pumps. Due to the relatively low energy use for this task and the applicability of the Knudsen Compressor to other applications, the development of a Knudsen Compressor utilizing MEMS fabrication techniques has been initiated. This paper discusses the initial fabrication of a single-stage MEMS Knudsen Compressor vacuum pump, provides performance criteria such as pumping speed, size, energy use and ultimate pressure and details vacuum pump applications in several MDL related in-situ instruments.

Vargo, S. E.↗

NASA NTRS2000

Unsteady Analysis of Inlet-Compressor Acoustic Interactions Using Coupled 3-D and 1-D CFD Codes

It is well known that the dynamic response of a mixed compression supersonic inlet is very sensitive to the boundary condition imposed at the subsonic exit (engine face) of the inlet. In previous work, a 3-D computational fluid dynamics (CFD) inlet code (NPARC) was coupled at the engine face to a 3-D turbomachinery code (ADPAC) simulating an isolated rotor and the coupled simulation used to study the unsteady response of the inlet. The main problem with this approach is that the high fidelity turbomachinery simulation becomes prohibitively expensive as more stages are included in the simulation. In this paper, an alternative approach is explored, wherein the inlet code is coupled to a lesser fidelity 1-D transient compressor code (DYNTECC) which simulates the whole compressor. The specific application chosen for this evaluation is the collapsing bump experiment performed at the University of Cincinnati, wherein reflections of a large-amplitude acoustic pulse from a compressor were measured. The metrics for comparison are the pulse strength (time integral of the pulse amplitude) and wave form (shape). When the compressor is modeled by stage characteristics the computed strength is about ten percent greater than that for the experiment, but the wave shapes are in poor agreement. An alternate approach that uses a fixed rise in duct total pressure and temperature (so-called 'lossy' duct) to simulate a compressor gives good pulse shapes but the strength is about 30 percent low.

Suresh, A.↗

NASA NTRS1999

Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier-Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub comer stall. Both the measurements and the numerical solutions for the stator indicate that a strong twister-like vortex is formed near the rear part of the blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub comer stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

Hah, C.↗

NASA NTRS1947

Investigation of an Experimental Supersonic Axial-Flow Compressor

An investigation is in progress at the Langley Laboratory of the NACA to explore the possibilities of axial-flow compressors operating with supersonic velocities relative to the blade rows. The first phase of this investigation, a study of supersonic diffusers, has been reported. The second phase, an analysis of supersonic compressors, has also been reported. Preliminary calculations have shown that very high pressure ratios across a stage, together with somewhat increased mass flows, are possible with compressors which decelerate air through the speed of sound in their rotor blading. These performance characteristics are desirable in compressors for aircraft jet propulsion units, gas turbines, or superchargers. The third phase, presented here, is a preliminary experimental investigation of a supersonic compressor designed to produce a high pressure ratio in a single stage.

Erwin, John R.↗