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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 271 records · Page 15

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NASA NTRS1988

Measuring Fan-Blade-Tip Displacements

Magnets and Hall-effect transducers used to monitor tip-to-shroud clearances. Mounting permanent magnet in end of fan blade and monitoring tip-to-shroud distance by use of output of Hall-effect transducers affixed to surface of shroud. Transducers provide real-time information. By placing magnets in each fan-blade tip, complete set monitored sequentially as each blade rotates past sensor. Application in many rotating machines, providing surrounding materials nonmagnetic and temperatures not severe.

Berry, Robert F., Jr.↗

NASA NTRS1987

Localization of natural modes of vibration in bladed disks

A study is presented of the mode localization phenomenon in imperfect blade-disk and blade-shroud-disk assemblies. The results indicate that unshrouded blades mounted on stiff disks are especially susceptible, and even small blade imperfections within manufacturing tolerances are likely to trigger mode localization. Increasing the interblade coupling by adding shrouds or reducing the disk stiffness greatly reduces the localization susceptiblity, although certain modes may still become localized if the shrouds are free to slip.

Bendiksen, O. O.↗

NASA NTRS1988

Rotor-stator interaction in a diffuser pump

Steady and unsteady pressure measurements were obtained on the diffuser vanes and the shroud wall of vaned and vaneless diffuser pumps in order to investigate the rotor-stator interaction. It is shown that the magnitude of the fluctuating lift is greater than the steady lift, and that the pressure fluctuations are larger on the suction side than on the pressure side. Pressure fluctuations near the shroud were much smaller than those near the hub across the span of the vane. Pressure fluctuations on the shroud wall itself were found to be larger for the vaned diffuser than for the vaneless diffuser.

Acosta, A. J.↗

NASA NTRS1986

Heat transfer in the tip region of a rotor blade simulator

In gas turbines, the blades of axial turbine stages rotate in close proximity to a stationary peripheral wall. Differential expansion of the turbine wheel, blades, and the shroud causes variations in the size of the clearance gap between blade tip and stationary shroud. The necessity to tolerate this differential thermal expansion dictates that the clearance gap cannot be eliminated altogether, despite accurate engine machining. Pressure differences between the pressure and suction sides of a blade drives a flow through the clearance gap. This flow, the tip leakage flow, is detrimental to engine performance. The primary detrimental effect of tip leakage flow is the reduction of turbine stage efficiency, and a second is the convective heat transfer associated with the flow. The surface area at the blade tip in contact with the hot working gas represents an additional thermal loading on the blade which, together with heat transfer to the suction and pressure side surface area, must be removed by the blade internal cooling flows. Experimental results concerned with the local heat transfer characteristics on all surfaces of shrouded, rectangular cavities are reported. A brief discussion of the mass transfer system used is given.

Chyu, M. K.↗

NASA NTRS1992

A new approach for performing contamination control bakeouts in JPL thermal vacuum test chambers

Contamination control requirements for the Wide Field/Planetary Camera II (WF/PC II) are necessarily stringent to protect against post-launch contamination of the sensitive optical surfaces, particularly the cold charge coupled device (CCD) imaging surfaces. Typically, thermal vacuum test chambers have employed a liquid nitrogen (LN2) cold trap to collect outgassed contaminants. This approach has the disadvantage of risking recontamination of the test article from shroud offgassing during post-test warmup of the chamber or from any shroud warming of even a few degrees during the bakeout process. By using an enclave, essentially a chamber within a chamber, configured concentrically and internally within an LN2 shroud, a method was developed, based on a design concept by Taylor, for preventing recontamination of test articles during bakeouts and subsequent post-test warmup of the vacuum chamber. Enclaves for testing WF/PC II components were designed and fabricated, then installed in three of JPL's Environmental Test Lab chambers. The design concepts, operating procedures, and test results of this development are discussed.

Johnson, Kenneth R.↗

NASA NTRS1993

Integrity testing of brush seal in a T-700 engine

A split-ring brush seal was fabricated, installed between two labyrinth-honeycomb shroud seals, and tested in the fourth-stage turbine of a T-700 engine. The annealed Haynes 25 bristles rubbed directly against the nonconditioned, irregular Rene 80 turbine blade shroud surface. A total of 21 hr of cyclic and steady-state data were taken with surface speeds of 335 m/s (1100 ft/s) and shroud temperatures to 620 C (1150 F). Wear appeared to be rapid initially, with an orange flash of hot brush fragments during the first engine startup, to minimal after 10 hr of operation. The brush survived the testing but experienced some bristle pullouts and severe bristle wear; some turbine interface wear and possible material transfer was noted. Future design concerns center on tribological behavior at the interface with or without lubricants.

Hendricks, Robert C.↗

NASA NTRS1994

Experimental investigation of turbine blade-tip excitation forces

Results of a program to investigate the magnitude and parametric variations of rotordynamic forces which arise in high power turbines due to blade-tip leakage effects are presented. Five different unshrouded turbine configurations and one configuration shrouded with a labyrinth seal were tested with static offsets of the turbine shaft. The forces along and perpendicular to the offset were measured directly with a rotating dynometer. Exploration of casing pressure and flow velocity distributions was used to investigate the force-generating mechanisms. For unshrouded turbines, the cross-forces originate mainly from the classical Alford mechanisms while the direct forces arise mainly from a slightly skewed pressure pattern. The Alford coefficient for cross-force was found to vary between 2.4 and 4.0, while the similar direct force coefficient varied from 1.5 to 3.5. The cross-forces are found to increase substantially when the gap is reduced from 3.0 to 1.9% of blade height, probably due to viscous blade-tip effects. The forces also increase when the hub gap between stator and rotor decreases. The force coefficient decreased with operating flow coefficient. In the case of the shrouded turbine, most of the forces arise from nonuniform seal pressures. This includes about 80% for the transverse forces. The rest appears to come from uneven work extraction. Their level is about 50% higher in the shrouded case.

Martinez-Sanchez, Manuel↗

NASA NTRS1960

Transonic Performance Characteristics of Several Jet Noise Suppressors

An investigation of the transonic performance characteristics of several noise-suppressor configurations has been conducted in the Langley 16-foot transonic tunnel. The models were tested statically and over a Mach number range from 0.70 to 1.05 at an angle of attack of 0 deg. The primary jet total-pressure ratio was varied from 1.0 (jet off) to about 4.5. The effect of secondary air flow on the performance of two of the configurations was investigated. A hydrogen peroxide turbojet-engine simulator was used to supply the hot-jet exhaust. An 8-lobe afterbody with centerbody, short shroud, and secondary air had the highest thrust-minus-drag coefficients of the six noise-suppressor configurations tested. The 12-tube and 12-lobe afterbodies had the lowest internal losses. The presence of an ejector shroud partially shields the external pressure distribution of the 8-lobe after-body from the influence of the primary jet. A ring-airfoil shroud increased the static thrust of the annular nozzle but generally decreased the thrust minus drag at transonic Mach numbers.

Schmeer, James W.↗

NASA NTRS1960

Ranger Spacecraft

The thermal-control philosophy of the spacecraft currently under development by the Jet Propulsion Laboratory is design by passive means to maintain all components within the tolerances specified by cognizant engineers. Due to the complexity of the configurations, calculations are) of necessity, fairly generalized and final design is based upon tests in an environmental chamber. The Ranger series spacecraft is designed with a basic structure which is common to all models, with additional hardware to suit the individual mission. This basic structure of Rangers A-1 and A-2 is seen as the hexagonal instrument section, the erectable solar panels, the movable antenna, and the omniantenna. The Ranger A-1 and A-2 configuration is for engineering tests and space-exploration, with the scientific instrumentation isolation requirement dictating the spread-out design. The spacecraft stands 12 feet high, weighs 700 to 800 pounds, and has an internal power of 150 watts. Rangers A-3, A-4, and A-5 are designed to rough land a capsule on the moon. For these, a capsule and retrorocket replace the scientific instruments, occupying the space inside the tower structure. The spacecraft must survive many environments. Chronologically they are: 1) Folded configuration inside an aerodynamic shroud on the pad. 2) Thermal flux from shroud aerodynamically heated during boost phase. 3) Coasting up to 30 minutes attached to Agena stage after booster and shroud are separated. 4) Agena stage burning. 5) Coasting and tumbling after separation from Agena until it passes from earth's shadow. 6) Upon reaching sunlight, panels open and begin sun acquisition. 7) Antenna seeks earth after spacecraft locks onto sun. 8) Space phase- "steady state" with vehicle's vertical axis locked on sun, communicating with earth. The philosophy is to design for the sun-acquired mode, making allowances for the transient conditions.

Thostesen, T. O.↗

NASA NTRS1947

Effect of Various Methods of Boundary-Layer Control on Performance of V-1710-93 Engine-Stage Supercharger

Four methods of boundary-layer control were tried during an investigation to improve the flow in the impeller passages of a V-1710-93 engine-stage supercharger. The boundary layer along the impeller front shroud was removed by suction. In one method the removal was accomplished by recirculation of the air to the impeller inlet; in another method, by external removal. In the other methods, slots were cut through the impeller-blade faces first at 30 percent and then at 30 and 70 percent of the mean-flow-path length measured from leading edges of the rotating inlet guide vanes to introduce air from the high-pressure side of the blades into the region where stagnation and separation were suspected. A slight improvement in performance was obtained when the boundary layer was removed through the impeller front shroud. In general, this improvement become more pronounced as the amount of air removed was increased even though the excessive impeller frontal clearance maintained for these tests, together with an exaggerated negative pressure gradient, apparently induced flow separation on the diffuser front and rear walls as well as on the impeller front shroud. The use of slots in the impellers at the locations selected had a detrimental effect on the supercharger performance characteristics.

Kohl, Robert C.↗

NASA NTRS1998

Effect of Installation of Mixer/Ejector Nozzles on the Core Flow Exhaust of High-Bypass-Ratio Turbofan Engines

The aerospace industry is currently investigating the effect of installing mixer/ejector nozzles on the core flow exhaust of high-bypass-ratio turbofan engines. This effort includes both full-scale engine tests at sea level conditions and subscale tests in static test facilities. Subscale model tests are to be conducted prior to full-scale testing. With this approach, model results can be analyzed and compared with analytical predications. Problem areas can then be identified and design changes made and verified in subscale prior to committing to any final design configurations for engine ground tests. One of the subscale model test programs for the integrated mixer/ejector development was a joint test conducted by the NASA Lewis Research Center and Pratt & Whitney Aircraft. This test was conducted to study various mixer/ejector nozzle configurations installed on the core flow exhaust of advanced, high-bypass-ratio turbofan engines for subsonic, commercial applications. The mixer/ejector concept involves the introduction of largescale, low-loss, streamwise vortices that entrain large amounts of secondary air and rapidly mix it with the primary stream. This results in increased ejector pumping relative to conventional ejectors and in more complete mixing within the ejector shroud. The latter improves thrust performance through the efficient energy exchange between the primary and secondary streams. This experimental program was completed in April 1997 in Lewis' CE-22 static test facility. Variables tested included the nozzle area ratio (A9/A8), which ranged from 1.6 to 3.0. This ratio was varied by increasing or decreasing the nozzle throat area, A8. Primary nozzles tested included both lobed mixers and conical primaries. These configurations were tested with and without an outer shroud, and the shroud position was varied by inserting spacers in it. In addition, data were acquired with and without secondary flow.

Harrington, Douglas E.↗

NASA NTRS2004

Pulsed Ejector Thrust Amplification Tested and Modeled

There is currently much interest in pulsed detonation engines for aeronautical propulsion. This, in turn, has sparked renewed interest in pulsed ejectors to increase the thrust of such engines, since previous, though limited, research had indicated that pulsed ejectors could double the thrust in a short device. An experiment has been run at the NASA Glenn Research Center, using a shrouded Hartmann-Sprenger tube as a source of pulsed flow, to measure the thrust augmentation of a statistically designed set of ejectors. A Hartmann- Sprenger tube directs the flow from a supersonic nozzle (Mach 2 in the present experiment) into a closed tube. Under appropriate conditions, an oscillation is set up in which the jet flow alternately fills the tube and then spills around flow emerging from the tube. The tube length determines the frequency of oscillation. By shrouding the tube, the flow was directed out of the shroud as an axial stream. The set of ejectors comprised three different ejector lengths, three ejector diameters, and three nose radii. The thrust of the jet alone, and then of the jet plus ejector, was measured using a thrust plate. The arrangement is shown in this photograph. Thrust augmentation is defined as the thrust of the jet with an ejector divided by the thrust of the jet alone. The experiments exhibited an optimum ejector diameter and length for maximizing the thrust augmentation, but little dependence on nose radius. Different frequencies were produced by changing the length of the Hartmann-Sprenger tube, and the experiment was run at a total of four frequencies. Additional measurements showed that the major feature of the pulsed jet was a starting vortex ring. The size of the vortex ring depended on the frequency, as did the optimum ejector diameter.

Wilson, Jack↗

NASA NTRS2007

Shock Characteristics Measured Upstream of Both a Forward-Swept and an Aft-Swept Fan

Three different types of diagnostic data-blade surface flow visualization, shroud unsteady pressure, and laser Doppler velocimeter (LDV)--were obtained on two fans, one forward-swept and one aft-swept, in order to learn more about the shocks which propagate upstream of these rotors when they are operated at transonic tip speeds. Flow visualization data are presented for the forward-swept fan operating at 13831 rpm(sub c), and for the aft-swept fan operating at 12500 and 13831 rpm(sub c) (corresponding to tip rotational Mach numbers of 1.07 and 1.19, respectively). The flow visualization data identify where the shocks occur on the suction side of the rotor blades. These data show that at the takeoff speed, 13831 rpm(sub c), the shocks occurring in the tip region of the forward-swept fan are further downstream in the blade passage than with the aft-swept fan. Shroud unsteady pressure measurements were acquired using a linear array of 15 equally-spaced pressure transducers extending from two tip axial chords upstream to 0.8 tip axial chords downstream of the static position of the tip leading edge of each rotor. Such data are presented for each fan operating at one subsonic and five transonic tip speeds. The unsteady pressure data show relatively strong detached shocks propagating upstream of the aft-swept rotor at the three lowest transonic tip speeds, and weak, oblique pressure disturbances attached to the tip of the aft-swept fan at the two highest transonic tip speeds. The unsteady pressure measurements made with the forward-swept fan do not show strong shocks propagating upstream of that rotor at any of the tested speeds. A comparison of the forward-swept and aft-swept shroud unsteady pressure measurements indicates that at any given transonic speed the pressure disturbance just upstream of the tip of the forward-swept fan is much weaker than that of the aft-swept fan. The LDV data suggest that at 12500 and 13831 rpm(sub c), the forward-swept fan swallowed the passage shocks occurring in the tip region of the blades, whereas the aft-swept fan did not. Due to this difference, the flows just upstream of the two fans were found to be quite different at both of these transonic speeds. Nevertheless, despite distinct differences just upstream of the two rotors, the two fan flows were much more alike about one axial blade chord further upstream. As a result, the LDV data suggest that it is unwise to attempt to determine the effect that the shocks have on far field noise by focusing only on measurements (or CFD predictions) made very near the rotor. Instead, these data suggest that it is important to track the shocks throughout the inlet.

Podboy, Gary G.↗

NASA NTRS1994

High temperature turbine engine structure

A high temperature ceramic/metallic turbine engine includes a metallic housing which journals a rotor member of the turbine engine. A ceramic disk-like shroud portion of the engine is supported on the metallic housing portion and maintains a close running clearance with the rotor member. A ceramic spacer assembly maintains the close running clearance of the shroud portion and rotor member despite differential thermal movements between the shroud portion and metallic housing portion.

Carruthers, William D.↗

NASA NTRS2000

Scan mirror remote temperature sensing system and method

A remote temperature sensing system (10) for a scanning mirror (7). The system (10) includes a sensor which detects heat radiated by the mirror and provides a signal in response thereto. In the illustrative implementation, the system (10) includes a thermistor mounted within a housing. The housing is contoured to maximize the receipt of thermal energy thereby. A mounting assembly maintains the thermistor a predetermined nonzero distance from the scanning mirror (7). The invention includes a shroud (12) mounted on the mirror (7) for shielding the thermistor and a support tube connected to the thermistor housing on a first end and to a base on the second end thereof. The support tube is adapted to remain stationary within the shroud as the scanning mirror and the shroud rotate due to the scanning of the mirror. Wires are connected to the thermistor on a first end thereof and are wrapped around the support tube. The wires include a length of electrically conductive material having a resistivity which has a low sensitivity to temperature variations. The sensor output is processed in a conventional manner to provide an output indicative of the temperature of the mirror.

VanDyk, Steven G.↗

NASA NTRS1993

High temperature turbine engine structure

Carruthers, William D.↗

NASA NTRS1992

High temperature turbine engine structure

Carruthers, William D.↗

NASA NTRS1995

Active noise control using noise source having adaptive resonant frequency tuning through variable ring loading

A noise source for an aircraft engine active noise cancellation system in which the resonant frequency of noise radiating structure is tuned to permit noise cancellation over a wide range of frequencies. The resonant frequency of the noise radiating structure is tuned by a plurality of drivers arranged to contact the noise radiating structure. Excitation of the drivers causes expansion or contraction of the drivers, thereby varying the edge loading applied to the noise radiating structure. The drivers are actuated by a controller which receives input of a feedback signal proportional to displacement of the noise radiating element and a signal corresponding to the blade passage frequency of the engine's fan. In response, the controller determines a control signal which is sent to the drivers, causing them to expand or contract. The noise radiating structure may be either the outer shroud of the engine or a ring mounted flush with an inner wall of the shroud or disposed in the interior of the shroud.

Pla, Frederic G.↗