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At least 73 records · Page 4

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NASA NTRS1978

Acoustical effects of blade tip shape changes on a full scale helicopter rotor in a wind tunnel

Four tip shapes were tested. They were rectangular, swept, tapered, and swept-tapered. The measured data covered a wide range of operating conditions. The range of advancing tip Mach numbers were between 0.72 to 0.96, and the advance ratios were from 0.2 to 0.375. At low and moderate advancing tip Mach number, the data in the dbA scale appear to indicate the swept tip is the quietest, swept tapered the second, tapered third and rectangular the most noisy. Above an advancing tip Mach number of about 0.89, a distinct acoustical pulse can be observed, which dominates the acoustical waveform. The pulse shape is symmetric at moderate tip Mach number, changing to a sawtooth shape at high advancing tip Mach numbers. Based on the amplitude of the impulsive noise, it appears the swept-tapered tip is the quietest, tapered tip the second, swept tip third and square tip the most noisy. The data presented in this report should be useful as data bases for modeling and evaluating helicopter impulsive noise.

Lee, A.↗

NASA NTRS1985

Analysis of the free-tip rotor wind-tunnel test results

The results from a wind tunnel test of a small scale free-tip rotor are analyzed. The free-tip rotor has blade tips that are free to weathervane into the tip's relative wind, thus producing a more uniform lift around the azimuth. The free-tip assembly, which includes the controller, functioned flawlessly throughout the test. In a test of the free-tip's response after passing through a vertical air jet, the tip pitched freely and in a controlled manner. Analysis of the tip's response characteristics showed the free-tip system's damped natural frequency to be 5.2 per rev. Tip pitch angle responses to the local airstream are presented for an advance-ratio range of 0.1 to 0.397 and for a solidity weighted rotor lift coefficient range of 0.038 to 0.092. Harmonic analysis of the responses showed a dominance by the first harmonic. As a result of the tip being free, forward flight power requirements were reduced by 8% or more. More power reduction was recorded for high thrust conditions. In addition to the power reduction, flatwise blade bending moments were reduced by as much as 30% at the inboard blade stations.

Stroub, Robert H.↗

NASA NTRS2000

Heat Transfer and Flow on the Squealer Tip of a Gas Turbine Blade

Experimental investigations are performed to measure the detailed heat transfer coefficient and static pressure distributions on the squealer tip of a gas turbine blade in a five-bladed stationary linear cascade. The blade is a 2-dimensional model of a modem first stage gas turbine rotor blade with a blade tip profile of a GE-E(sup 3) aircraft gas turbine engine rotor blade. A squealer (recessed) tip with a 3.77% recess is considered here. The data on the squealer tip are also compared with a flat tip case. All measurements are made at three different tip gap clearances of about 1%, 1.5%, and 2.5% of the blade span. Two different turbulence intensities of 6.1% and 9.7% at the cascade inlet are also considered for heat transfer measurements. Static pressure measurements are made in the mid-span and near-tip regions, as well as on the shroud surface opposite to the blade tip surface. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and an exit Reynolds number based on the axial chord of 1.1 x 10(exp 6). A transient liquid crystal technique is used to measure the heat transfer coefficients. Results show that the heat transfer coefficient on the cavity surface and rim increases with an increase in tip clearance. 'Me heat transfer coefficient on the rim is higher than the cavity surface. The cavity surface has a higher heat transfer coefficient near the leading edge region than the trailing edge region. The heat transfer coefficient on the pressure side rim and trailing edge region is higher at a higher turbulence intensity level of 9.7% over 6.1 % case. However, no significant difference in local heat transfer coefficient is observed inside the cavity and the suction side rim for the two turbulence intensities. The squealer tip blade provides a lower overall heat transfer coefficient when compared to the flat tip blade.

Azad, Gm S.↗

NASA NTRS2017

Toward Understanding Tip Leakage Flows in Small Compressor Cores Including Stator Leakage Flow

The focus of this work was to provide additional data to supplement the work reported in NASA/CR-2015-218868 (Berdanier and Key, 2015b). The aim of that project 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 clearances were studied with nominal tip clearance gaps of 1.5 percent, 3.0 percent, and 4.0 percent based on a constant annulus height. Overall compressor performance was previously investigated at four corrected speedlines (100 percent, 90 percent, 80 percent, and 68 percent) for each of the three tip clearance configurations. This study extends the previously published results to include detailed steady and time-resolved pressure data at two loading conditions, nominal loading (NL) and high loading (HL), on the 100 percent corrected speedline for the intermediate clearance level (3.0 percent). Steady detailed radial traverses of total pressure at the exit of each stator row are 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 were combined with existing three-component velocity measurements to identify a novel technique for calculating blockage in a multistage compressor. 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 percent corrected speedline using fast-response piezoresistive pressure sensors. These time-resolved static pressure measurements reveal new knowledge about the trajectory of the tip leakage flow through the rotor passage. Further, these data extend previous measurements identifying a modulation of the tip leakage flow due to upstream stator wake propagation. Finally, a novel instrumentation technique has been implemented to measure pressures in the shrouded stator cavities. These data provide boundary conditions relating to the flow across the shrouded stator knife seal teeth. Moreover, the utilization of fast-response pressure sensors provides a new look at the time-resolved pressure field, leading to instantaneous differential pressures across the seal teeth. 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 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↗

NASA NTRS1986

An experimental investigation of free-tip response to a jet

The aerodynamic response of passively oscillating tips appended to a model helicopter rotor was investigated during a whirl test. Tip responsiveness was found to meet free-tip rotor requirements. Experimental and analytical estimates of the free-tip aerodynamic spring, mechanical spring, and aerodynamic damping were calculated and compared. The free tips were analytically demonstrated to be operating outside the tip resonant response region at full-scale tip speeds. Further, tip resonance was shown to be independent of tip speed, given the assumption that the tip forcing frequency is linearly dependent upon the rotor rotational speed.

Young, L. A.↗

NASA NTRS2000

Heat Transfer and Pressure Distributions on a Gas Turbine Blade Tip

Heat transfer coefficient and static pressure distributions are experimentally investigated on a gas turbine blade tip in a five-bladed stationary linear cascade. The blade is a 2-dimensional model of a first stage gas turbine rotor blade with a blade tip profile of a GE-E(sup 3) aircraft gas turbine engine rotor blade. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1 x 10(exp 6). The middle 3-blade has a variable tip gap clearance. All measurements are made at three different tip gap clearances of about 1%, 1.5%, and 2.5% of the blade span. Heat transfer measurements are also made at two different turbulence intensity levels of 6.1 % and 9.7% at the cascade inlet. Static pressure measurements are made in the mid-span and the near-tip regions as well as on the shroud surface, opposite the blade tip surface. Detailed heat transfer coefficient distributions on the plane tip surface are measured using a transient liquid crystal technique. Results show various regions of high and low heat transfer coefficient on the tip surface. Tip clearance has a significant influence on local tip beat transfer coefficient distribution. Heat transfer coefficient also increases about 15-20% along the leakage flow path at higher turbulence intensity level of 9.7% over 6.1 %.

Azad, Gm S.↗

NASA NTRS2002

Tip-Clearance Vortex Characterized With Three-Dimensional Digital Particle Image Velocimetry

An optical measurement technique known as Three-Dimensional Digital Particle Image Velocimetry (3-D DPIV) was used to characterize the tip clearance flow in NASA Glenn Research Center's low-speed axial compressor. 3-D DPIV is a technique in which a stereoscopic imaging system consisting of two cross-correlation cameras is used to record particles entrained in a flow as a laser light sheet is pulsed at two instances in time. Although 3-D DPIV has been used elsewhere, this is the first time it has been used to measure compressor tip clearance flows. In-house modifications of the DPIV system include the use of effective seeding technology and a novel system to perform a priori calibrations at all five measurement planes, greatly reducing facility run time. Computational fluid dynamics predictions, which are used to guide design changes toward improving the efficiency and operating range of turbomachinery, can be verified and improved by comparison with 3-D DPIV measurements of the actual tip clearance flow. This measurement campaign dealt with the characterization of the tip clearance vortex in the first stage of a four-stage axial compressor. The tip clearance vortex is formed in compressors operating with a clearance gap between the moving rotor blade tips and the stationary casing when a leakage flow, forced from the pressure side of the blade over the blade tip, forms a vortical structure on the suction side of the blade. 3-D DPIV is ideally suited to measure the clearance vortex for two reasons: (1) this technique captures the entire blade passage flow at one instant in time, so that wandering of the vortex during the measurement does not smear out velocity gradients in the flow field, and (2) the spanwise component of velocity changes sign across the vortex core, providing a more accurate measurement of the vortex location than was available with previous two-dimensional measurement approaches. These two attributes of the data will enable computational fluid dynamics researchers to validate their predictions to a level of accuracy not previously attainable. In an effort to understand the effects of this flow phenomenon on the operation of the low speed axial compressor, data were acquired at two mass flow coefficients: 0.395 (design operating point) and 0.35 (operating point just above stall). In order to identify the spatial extent, location, and magnitude of the tip clearance vortex for each mass flow condition, data were acquired at five equally spaced spanwise locations, from 90- to 100-percent span. The data presented shows a cross section of the tip clearance vortex at 94-percent span and a mass flow coefficient of 0.395, where the color contours represent the radial velocity component. The blue (flow towards the hub) and red (flow towards the casing) regions illustrate the extent of the tip clearance vortex. The interface between the blue and red regions delineates the location of the vortex core.

John, W. Trevor↗

NASA NTRS2010

Analysis of Unsteady Tip and Endwall Heat Transfer in a Highly Loaded Transonic Turbine Stage

In a previous study, vane-rotor shock interactions and heat transfer on the rotor blade of a highly loaded transonic turbine stage were simulated. The geometry consists of a high pressure turbine vane and downstream rotor blade. This study focuses on the physics of flow and heat transfer in the rotor tip, casing and hub regions. The simulation was performed using the Unsteady Reynolds-Averaged Navier-Stokes (URANS) code MSU-TURBO. A low Reynolds number k-epsilon model was utilized to model turbulence. The rotor blade in question has a tip gap height of 2.1 percent of the blade height. The Reynolds number of the flow is approximately 3x10(exp 6) per meter. Unsteadiness was observed at the tip surface that results in intermittent "hot spots". It is demonstrated that unsteadiness in the tip gap is governed by inviscid effects due to high speed flow and is not strongly dependent on pressure ratio across the tip gap contrary to published observations that have primarily dealt with subsonic tip flows. The high relative Mach numbers in the tip gap lead to a choking of the leakage flow that translates to a relative attenuation of losses at higher loading. The efficacy of new tip geometry is discussed to minimize heat flux at the tip while maintaining choked conditions. In addition, an explanation is provided that shows the mechanism behind the rise in stagnation temperature on the casing to values above the absolute total temperature at the inlet. It is concluded that even in steady mode, work transfer to the near tip fluid occurs due to relative shearing by the casing. This is believed to be the first such explanation of the work transfer phenomenon in the open literature. The difference in pattern between steady and time-averaged heat flux at the hub is also explained.

Shyam, Vikram↗

NASA NTRS2012

Investigation of Turbulent Tip Leakage Vortex in an Axial Water Jet Pump with Large Eddy Simulation

Detailed steady and unsteady numerical studies were performed to investigate tip clearance flow in an axial water jet pump. The primary objective is to understand physics of unsteady tip clearance flow, unsteady tip leakage vortex, and cavitation inception in an axial water jet pump. Steady pressure field and resulting steady tip leakage vortex from a steady flow analysis do not seem to explain measured cavitation inception correctly. The measured flow field near the tip is unsteady and measured cavitation inception is highly transient. Flow visualization with cavitation bubbles shows that the leakage vortex is oscillating significantly and many intermittent vortex ropes are present between the suction side of the blade and the tip leakage core vortex. Although the flow field is highly transient, the overall flow structure is stable and a characteristic frequency seems to exist. To capture relevant flow physics as much as possible, a Reynolds-averaged Navier-Stokes (RANS) calculation and a Large Eddy Simulation (LES) were applied for the current investigation. The present study reveals that several vortices from the tip leakage vortex system cross the tip gap of the adjacent blade periodically. Sudden changes in local pressure field inside tip gap due to these vortices create vortex ropes. The instantaneous pressure filed inside the tip gap is drastically different from that of the steady flow simulation. Unsteady flow simulation which can calculate unsteady vortex motion is necessary to calculate cavitation inception accurately even at design flow condition in such a water jet pump.

Hah, Chunill↗

NASA NTRS2018

Active Turbine Tip Clearance Control Trade Space Analysis of an Advanced Geared Turbofan Engine

Tip clearance within the high pressure turbine of a gas turbine engine is a significant factor in engine performance and efficiency. In the pursuit of higher efficiency, aero-engine designs are migrating toward compact gas turbine (CGT) technology that seeks to increase the bypass ratio of the gas turbine engine without increasing the size of the fan, which is constrained by its underwing location. The reduced size of CGTs invoke concern over increased sensitivity of engine performance due to turbine tip clearance gap that makes an argument for advanced tip clearance mitigation and control techniques to be employed. This paper evaluates the tip clearance trade space for a conceptual geared turbofan engine with a CGT core. This is accomplished through a modeling and simulation approach that includes a sensitivity analysis of engine performance in response to high pressure turbine tip clearance as well as an evaluation of the sensitivity of tip clearance to various design parameters, including material properties and component cooling characteristics. Also included is a parametric study of actuators that provides preliminary requirements for implementation of active turbine tip clearance control actuation systems. The results produced from these studies are meant to be informative, with special emphasis on the demonstration of a systematic approach. The modeling approach appears to capture expected trends. The studies suggest that the tip clearance gap will have a greater impact on the new CGT engines and that a relatively slow, actively controlled actuation system may be sufficient as long as it has control authority to both open and close the tip clearance gap.

tip clearance control↗

NASA NTRS1983

Oil-flow study of a Space Shuttle orbiter tip-fin controller

Possible use of tip-fin controllers instead of a vertical tail on advanced winged entry vehicles was examined. Elimination of the vertical tail and using tip-fins offers the advantages of positive yaw control at high angles of attack and a potential weight savings. Oil-flow technique was used to obtain surface flow patterns on a tip-fin installed on a 0.01-scale Space Shuttle orbiter model for the purpose of assessing the extent of flow interference effects on the wing and tip-fin which might lead to serious heating problems. Tests were conducted in air at Mach 10 for a free-stream Reynolds numbers of .000113 at 20, 30, and 40 degree angle of attack and sideslip angles of 0 and 2 degree. Elevon deflections of -10, 0, and 10 degree and tip-fin control-surface deflections of 0, 20, and 40 degree were employed. Test results were also used to aid in the interpretation of heating data obtained on a Shuttle orbiter tip-fin on another model in a different facility. A limited comparison of oil-flow patterns and heat-transfer data is included. It was determined that elevon deflection angles from -10 to 10 degree and sideslip angles up to 2 degree have very little effect on tip-fin surface flow patterns. Also, there is a minimum of interference between the tip-fin and the wing. The most significant flow interactions occur on the tip-fin onboard surface as a result of its control-surface deflections.

Helms, V. T., III↗

NASA NTRS1985

A three-dimensional viscous flow analysis for the helicopter tip vortex generation problem

The tip vortex flow field occurring in the vicinity of the tip region of a a helicopter rotor blade is a very complicated three-dimensional, viscous flow phenomenon. The details of the flow in the tip region can have a major effect in determining the generated rotor noise and can significantly affect the performance and dynamic loading of the rotor blade. The three-dimensional viscous subsonic tip vortex generation processes is investigated by a numerical procedure which allows spatial forward-marching integration, utilizing flow approximations from the velocity-decomposition approach of Briley and McDonald. The approach has been applied to compute the laminar and turbulent tip vortex flows for a constant thickness slab airfoil with a square tip, a constant thickness slab airfoil with a half round tip and a NACA 0012 airfoil with a half round tip. The basic mechanism of the tip vortex generation process as well as the prediction of vortex appearance, strength and secondary flow shown by the calculations are in qualitative agreement with experimental results.

Lin, S. J.↗

NASA NTRS1986

Numerical simulation of tip vortices of wings in subsonic and transonic flows

A multi block zonal algorithm which solves the thin-layer Navier-Stokes and the Euler equations is used to numerically simulate the formation and roll-up of the tip vortex in both subsonic and transonic flows. Four test cases which used small and large aspect ratio wings have been considered to examine the influence of the tip-cap shape, the tip planform and the free-stream Mach number. It appears that both the tip-planform and the tip-cap shape have some influence on the formation of the tip vortex, but its subsequent roll-up seems to be more influenced by the tip-planform shape. In general, a good definition of the formation and the roll-up of the tip vortex has been observed for all the cases considered here. Comparions of the numerical results with the limited, available experimental data show good agreement with both the surface pressures and the tip-vortex strength.

Srinivasan, G. R.↗

NASA NTRS1999

Heat Transfer and Flow on the First Stage Blade Tip of a Power Generation Gas Turbine: Experimental Results - Part 1

A combined computational and experimental study has been performed to investigate the detailed distribution of convective heat transfer coefficients on the first stage blade tip surface for a geometry typical of large power generation turbines(>100MW). This paper is concerned with the design and execution of the experimental portion of the study. A stationary blade cascade experiment has been run consisting of three airfoils, the center airfoil having a variable tip gap clearance. The airfoil models the aerodynamic tip section of a high pressure turbine blade with inlet Mach number of 0.30, exit Mach number of 0.75, pressure ratio of 1.45, exit Reynolds number based on axial chord of 2.57 x 10(exp 6), and total turning of about 110 degrees. A hue detection based liquid crystal method is used to obtain the detailed heat transfer coefficient distribution on the blade tip surface for flat, smooth tip surfaces with both sharp and rounded edges. The cascade inlet turbulence intensity level took on values of either 5% or 9%. The cascade also models the casing recess in the shroud surface ahead of the blade. Experimental results are shown for the pressure distribution measurements on the airfoil near the tip gap, on the blade tip surface, and on the opposite shroud surface. Tip surface heat transfer coefficient distributions are shown for sharp-edge and rounded-edge tip geometries at each of the inlet turbulence intensity levels.

Bunker, Ronald S.↗

NASA NTRS1992

Aerodynamic Spring and Damping of Free-Pitching Tips on a Semispan Wing

A test was conducted in the NASA Ames 7- by 10-Foot Wind Tunnel to derive aerodynamic spring and damping estimates for free-pitching tips on a semispan wing. The test model was a rectangular planform semispan wing with wing tips that had a single rigid-body pitch degree of freedom with respect to the inboard wing. A number of different tip planform geometries were tested, incorporating a range of quarter-chord sweep angles and taper ratios. The wing-tip dynamic response characteristics were measured at several wing angles of attack and tunnel dynamic pressures. The tip oscillations were initiated by releasing the tips from prescribed angles of attack. A new method to isolate Coulomb damping from aerodynamic damping from these tip-motion time histories is developed and applied. Correlations were performed between the experimentally derived wing-tip aerodynamic spring and damping values and predictions from a semiempirical analysis based on steady-state tip aerodynamic loads.

Young, Larry A.↗

NASA NTRS1999

Dual-Laser Probe for Measuring Blade-Tip Clearance Tested

A dual-probe, integrated fiber-optic laser system for measuring blade tip clearance in rotating turbomachinery was developed cooperatively with Integrated Fiber Optic Systems, Inc., and tested at the NASA Lewis Research Center. Because the probes are nearly flush with the casing inner lining, there is minimal flow disturbance. The two probes are closely spaced in a circumferential plane and are slanted at an angle relative to each other so that the time it takes the blade tip to traverse the space between the two beams varies with the tip radius, making it possible to determine the tip clearance at rotor operating conditions. The tip clearance can be obtained for all the blades in a rotor with a single system, provided there are no synchronous vibrations present at a particular operating condition. The two probes for the laser system were installed in two holders. For one, the included angle between the probes was 20 degrees, and for the other, it was 40 degrees. The two configurations were calibrated in a vacuum spin-rig facility that can reproduce realistic blade-tip speeds. A specially designed, nondeflecting rotor was used to calibrate the probes. This rotor consisted of a tapered titanium bar with three teeth on each end, the thickness of which was representative of a typical blade-tip thickness. The bar and measuring teeth were in a horizontal plane in the spin rig, and the axis of rotation was vertical. To set the clearances, we used a remotely controlled optical stage that could be traversed vertically. The 40 probe system also was used to measure the tip clearance during an engine fan prototype test in a wind tunnel. Using the spin-rig calibrations, we estimated the accuracy of the tip-clearance measurement in the test to be 0.005 in. This program is ongoing.

Dhadwal, Harbans S.↗

NASA NTRS2016

Integrated Turbine Tip Clearance and Gas Turbine Engine Simulation

Gas turbine compressor and turbine blade tip clearance (i.e., the radial distance between the blade tip of an axial compressor or turbine and the containment structure) is a major contributing factor to gas path sealing, and can significantly affect engine efficiency and operational temperature. This paper details the creation of a generic but realistic high pressure turbine tip clearance model that may be used to facilitate active tip clearance control system research. This model uses a first principles approach to approximate thermal and mechanical deformations of the turbine system, taking into account the rotor, shroud, and blade tip components. Validation of the tip clearance model shows that the results are realistic and reflect values found in literature. In addition, this model has been integrated with a gas turbine engine simulation, creating a platform to explore engine performance as tip clearance is adjusted. Results from the integrated model explore the effects of tip clearance on engine operation and highlight advantages of tip clearance management.

aerothermodynamics↗