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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 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRS1950

Effect of Compressor-outlet Bleedoff on Turbojet-engine Performance

An investigation was conducted in the NACA Lewis altitude wind tunnel to determine the effect of compressor-outlet bleedoff on the performance of an axial-flow turbojet engine equipped with a variable-area exhaust nozzle. Results presented indicate the effect of compressor-outlet bleedoff on performance at altitudes of 25,000 and 40,000 feet and a flight Mach number of 0.53. Variation of performance with bleedoff flow is indicated for operation with fixed- and variable-area exhaust nozzles. Temperature and pressure losses through the bleedoff ducting system are also discussed.

TURBINES, GAS - BLEED-OFF↗
NASA NTRS1952

Investigation of turbines for driving supersonic compressors II : performance of first configuration with 2.2 percent reduction in nozzle flow area / Warner L. Stewart, Harold J. Schum, Robert Y. Wong

The experimental performance of a modified turbine for driving a supersonic compressor is presented and compared with the performance of the original configuration to illustrate the effect of small changes in the ratio of nozzle-throat area to rotor-throat area. Performance is based on the performance of turbines designed to operate with both blade rows close to choking. On the basis of the results of this investigation, the ratio of areas is concluded to become especially critical in the design of turbines such as those designed to drive high-speed, high-specific weight-flow compressors where the turbine nozzles and rotor are both very close to choking.

TURBINES--MIXED FLOW↗
NASA NTRS1954

Review of high-performance axial-flow-compressor blade-element theory

This report presents a review of current compressor blade-element theory with particular emphasis on application to the transonic high-performance compressor. A discussion of the significant parameters of total-pressure loss and deviation angle is presented, and an indication of the extent of available knowledge and the problems involved in the determination of blade-element characteristics is given. Some recent results and considerations in this pursuit and suggestions for further avenues of investigation are indicated.

Lieblein, Seymour↗
NASA NTRS1956

Aerodynamic Design of Axial-flow Compressors

Available experimental two-dimensional-cascade data for conventional compressor blade sections are correlated. The two-dimensional cascade and some of the principal aerodynamic factors involved in its operation are first briefly described. Then the data are analyzed by examining the variation of cascade performance at a reference incidence angle in the region of minimum loss. Variations of reference incidence angle, total-pressure loss, and deviation angle with cascade geometry, inlet Mach number, and Reynolds number are investigated. From the analysis and the correlations of the available data, rules and relations are evolved for the prediction of the magnitude of the reference total-pressure loss and the reference deviation and incidence angles for conventional blade profiles. These relations are developed in simplified forms readily applicable to compressor design procedures.

Source record↗
NASA NTRS1956

Analysis of Limitations Imposed on One-spool Turboprop-engine Designs by Compressors and Turbines at Flight Mach Numbers of 0, 0.6, and 0.8

Turbine centrifugal stress is a limiting factor for all flight conditions studied. This stress is more severe for sea-level operations than for subsonic flight at the tropopause. Turbines designed for a stress of 30,000 psi are capable of driving a light, compact, high-spedd compressor but only at high values of specific fuel consumption. An increase in turbine-inlet temperature is accompanied by an increase in turbine centrifugal stress. If stresses in excess of 50,000 psi can be tolerated, compressor aerodynamics may become a primary limitation.

Cavicchi, Richard H↗
NASA NTRS1957

Effect of Interstage Bleed on Rotating Stall and Blade Vibration in a 13-stage Axial-flow Compressor in a Turbojet Engine

The compressor case of a 13-stage axial-flow compressor was modified to incorporate air-bleed systems over the fifth and tenth rotor stages. Rotating stall-exited rotor-blade vibrations were measured in the first and second stages at 60 and 68 percent of rated speed, respectively. Either the fifth- or the tenth-stage bleeds satisfactorily eliminated the peak vibratory stresses.

Calvert, Howard F↗
NASA NTRS1952

Discussion of Boundary-Layer Characteristics Near the Wall of an Axial-Flow Compressor

The boundary-layer velocity profiles in the tip region of an axial-flow compressor downstream of the guide vanes and downstream of the rotor were measured by use of total-pressure and claw-type yaw probes. These velocities were resolved into two components: one along the streamline of the flow outside the boundary layer, and the other perpendicular to it. The affinity among all profiles was thus demonstrated with the boundary-layer thickness and the deflection of the boundary layer at the wall as the generalizing parameters. By use of these results and the momentum-integral equations, boundary-layer characteristics on the walls of an axial-flow compressor were qualitatively evaluated.

Mager, Artur↗
NASA NTRS1994

Cooled-Spool Piston Compressor

Proposed cooled-spool piston compressor driven by hydraulic power and features internal cooling of piston by flowing hydraulic fluid to limit temperature of compressed gas. Provides sufficient cooling for higher compression ratios or reactive gases. Unlike conventional piston compressors, all parts of compressed gas lie at all times within relatively short distance of cooled surface so that gas cooled more effectively.

Morris, Brian G.↗
NASA NTRS1994

Hermetically Sealed Compressor

Proposed hermetically sealed pump compresses fluid to pressure up to 4,000 atm (400 MPa). Pump employs linear electric motor instead of rotary motor to avoid need for leakage-prone rotary seals. In addition, linear-motor-powered pump would not require packings to seal its piston. Concept thus eliminates major cause of friction and wear. Pump is double-ended diaphragm-type compressor. All moving parts sealed within compressor housing.

Holtzapple, Mark T.↗
NASA NTRS1993

Development of the seeding system used for laser velocimeter surveys of the NASA Low-Speed Centrifugal Compressor flow field

An atomizer-based system for distributing high-volume rates of seed material was developed to support laser velocimeter investigations of the NASA Low-Speed Centrifugal Compressor flow field. The seeding system and the major concerns that were addressed during its development are described. Of primary importance were that the seed material be dispersed as single particles and that the liquid carrier used be completely evaporated before entering the compressor.

Wasserbauer, Charles A.↗
NASA NTRS1994

Magnetic power piston fluid compressor

A compressor with no moving parts in the traditional sense having a housing having an inlet end allowing a low pressure fluid to enter and an outlet end allowing a high pressure fluid to exit is described. Within the compressor housing is at least one compression stage to increase the pressure of the fluid within the housing. The compression stage has a quantity of magnetic powder within the housing, is supported by a screen that allows passage of the fluid, and a coil for selectively providing a magnetic field across the magnetic powder such that when the magnetic field is not present the individual particles of the powder are separated allowing the fluid to flow through the powder and when the magnetic field is present the individual particles of the powder pack together causing the powder mass to expand preventing the fluid from flowing through the powder and causing a pressure pulse to compress the fluid.

Gasser, Max G.↗
NASA NTRS1994

Design of a Low Aspect Ratio Transonic Compressor Stage Using CFD Techniques

A transonic compressor stage has been designed for the Naval Postgraduate School Turbopropulsion Laboratory. The design relied heavily on CFD techniques while minimizing conventional empirical design methods. The low aspect ratio (1.2) rotor has been designed for a specific head ratio of .25 and a tip relative inlet Mach number of 1.3. Overall stage pressure ratio is 1.56. The rotor was designed using an Euler code augmented by a distributed body force model to account for viscous effects. This provided a relatively quick-running design tool, and was used for both rotor and stator calculations. The initial stator sections were sized using a compressible, cascade panel code. In addition to being used as a case study for teaching purposes, the compressor stage will be used as a research stage. Detailed measurements, including non-intrusive LDV, will be compared with the design computations, and with the results of other CFD codes, as a means of assessing and improving the computational codes as design tools.

Sanger, Nelson L.↗
NASA NTRS1994

Rotor dynamic behaviour of a high-speed oil-free motor compressor with a rigid coupling supported on four radial magnetic bearings

The combination of a high-speed motor, dry gas seals, and magnetic bearings realized in this unit facilitates the elimination of oil. The motor is coupled with a quill shaft to the compressor. This yields higher natural frequencies of the rotor than with the use of a diaphragm coupling and helps to maintain a sufficient margin of the maximum speed to the frequency of the second compressor bending mode. However, the controller of each bearing then has to take the combined modes of both machines into account. The requirements for the controller to ensure stability and sufficient damping of all critical speeds are designed and compared with the implemented controller. The calculated closed loop behavior was confirmed experimentally, except the stability of some higher modes due to slight frequency deviations of the rotor model to the actual rotor. The influence of a mechanical damper as a device to provide additional damping to high models is demonstrated theoretically. After all, it was not necessary to install the damper, since all modes cold be stabilized by the controller.

Schmied, J.↗
NASA NTRS1994

Inlet Flow Test Calibration for a Small Axial Compressor Facility. Part 1: Design and Experimental Results

An axial compressor test rig has been designed for the operation of small turbomachines. The inlet region consisted of a long flowpath region with two series of support struts and a flapped inlet guide vane. A flow test was run to calibrate and determine the source and magnitudes of the loss mechanisms in the inlet for a highly loaded two-stage axial compressor test. Several flow conditions and IGV angle settings were established in which detailed surveys were completed. Boundary layer bleed was also provided along the casing of the inlet behind the support struts and ahead of the IGV. A detailed discussion of the flowpath design along with a summary of the experimental results are provided in Part 1.

Miller, D. P.↗
NASA NTRS1994

Relative performance comparison between baseline labyrinth and dual-brush compressor discharge seals in a T-700 engine test

In separate series of T-700 engine tests, direct comparisons were made between the forward-facing labyrinth and dual-brush compressor discharge seals. Compressor speeds to 43,000 rpm, surface speeds to 160 m/s (530 ft/s), pressures to 1 MPa (145 psi), and temperatures to 680 K (765 F) characterized these tests. The wear estimate for 40 hr of engine operations was less than 0.025 mm (0.001 in.) of the Haynes 25 alloy bristles running against a chromium-oxide-coated rub runner. The pressure drops were higher for the dual-brush than for the forward-facing labyrinth seal, implying better seal characteristics and engine performance for the brush seal. Modification of the secondary flow path requires that changes in cooling air and engine dynamics be accounted for.

Hendricks, Robert C.↗
NASA NTRS1994

Experimental/analytical approach to understanding mistuning in a transonic wind tunnel compressor

This paper will briefly set forth some of the basic tenets of mistuned rotating bladed-disk assemblies. The experience with an existing three stage compressor in a transonic wind tunnel will be documented. The manner in which the theoretical properties manifest themselves in this non-ideal compressor will be described. A description of mistuning behaviors that can and cannot be accurately substantiated will be discussed.

Kaiser, Teri↗
NASA NTRS1995

Relative performance comparison between baseline labyrinth and dual-brush compressor discharge seals in a T-700 engine test

In separate series of YT-700 engine tests, direct comparisons were made between the forward-facing labyrinth and dual brush compressor discharge seals. Compressor speeds to 43 000 rpm, surface speeds to 160 m/s (530 ft/s), pressures to 1 MPa (145 psi), and temperatures to 680 K (765 F) characterized these tests. The wear estimate for 46 hr of engine operations was less than 0.025 mm (0.001 in.) of the Haynes 25 alloy bristles running against a chromium-carbide-coated rub runner. The pressure drops were higher for the dual-brush seal than for the forward-facing labyrinth seal and leakage was lower-with the labyrinth seal leakage being 2-1/2 times greater-implying better seal characteristics, better secondary airflow distribution, and better engine performance (3 percent at high pressure to 5 percent at lower pressure) for the brush seal. (However, as brush seals wear down (after 500 to 1000 hr of engine operation), their leakage rates will increase.) Modification of the secondary flow path requires that changes in cooling air and engine dynamics be accounted for.

Hendricks, Robert C.↗
NASA NTRS1996

Experimental Investigation of the Flow Field in a Transonic, Axial Flow Compressor with Respect to the Development of Blockage and Loss

A detailed experimental investigation to understand and quantify the development of loss and blockage in the flow field of a transonic, axial flow compressor rotor has been undertaken. Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at design and off-design conditions. The rotor was operated at 100%, 85%, 80%, and 60% of design speed which provided inlet relative Mach numbers at the blade tip of 1.48, 1.26, 1.18, and 0.89 respectively. At design speed the blockage is evaluated ahead of the rotor passage shock, downstream of the rotor passage shock, and near the trailing edge of the blade row. The blockage is evaluated in the core flow area as well as in the casing endwall region. Similarly at pm speed conditions for the cases of (1) where the rotor passage shock is much weaker than that at design speed and (2) where there is no rotor passage shock, the blockage and loss are evaluated and compared to the results at design speed. Specifically, the impact of the rotor passage shock on the blockage and loss development, pertaining to both the shock/boundary layer interactions and the shock/tip clearance flow interactions, is discussed. In addition, the blockage evaluated from the experimental data is compared to (1) an existing correlation of blockage development which was based on computational results, and (2) computational results on a limited basis. The results indicate that for this rotor the blockage in the endwall region is 2-3 times that of the core flow region and the blockage in the core flow region more than doubles when the shock strength is sufficient to separate the suction surface boundary layer. The distribution of losses in the care flow region indicate that the total loss is primarily comprised of the shock loss when the shock strength is not sufficient to separate the suction surface boundary layer. However, when the shock strength is sufficient to separate the suction surface boundary layer, the profile loss is comparable to the shock loss and can exceed the shock loss.

Suder, Kenneth L.↗