Kaplan turbine tip vortex cavitation - analysis and prevention

NASA Astrophysics Data System (ADS)

Motycak, L.; Skotak, A.; Kupcik, R.

2012-11-01

The work is focused on one type of Kaplan turbine runner cavitation - a tip vortex cavitation. For detailed description of the tip vortex, the CFD analysis is used. On the basis of this analysis it is possible to estimate the intensity of cavitating vortex core, danger of possible blade surface and runner chamber cavitation pitting. In the paper, the ways how to avoid the pitting effect of the tip vortex are described. In order to prevent the blade surface against pitting, the following possibilities as the change of geometry of the runner blade, dimension of tip clearance and finally the installation of the anti-cavitation lips are discussed. The knowledge of the shape and intensity of the tip vortex helps to design the anti-cavitation lips more sophistically. After all, the results of the model tests of the Kaplan runner with or without anti-cavitation lips and the results of the CFD analysis are compared.

76 FR 64293 - Airworthiness Directives; CFM International, S. A. Model CFM56-5B Series Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2011-10-18

... serial number (S/N) fan blades, part number (P/N) 338- 002-114-0. This proposed AD was prompted by a normal quality sampling at CFM that isolated a production batch of fan blades with nonconforming geometry of mid-span shroud tips of the fan blades. This defect would cause the upper panel of the fan blade...

Unsteady Turbine Blade and Tip Heat Transfer Due to Wake Passing

NASA Technical Reports Server (NTRS)

Ameri, Ali A.; Rigby, David L.; Steinthorsson, Erlendur; Heidmann, James; Fabian, John C.

2007-01-01

The geometry and the flow conditions of the first stage turbine blade of GE s E3 engine have been used to obtain the unsteady three-dimensional blade and tip heat transfer. The isothermal wall boundary condition was used. The effect of the upstream wake of the first stage vane was of interest and was simulated by provision of a gust type boundary condition upstream of the blades. A one blade periodic domain was used. The consequence of this choice was explored in a preliminary study which showed little difference in the time mean heat transfer between 1:1 and 2:3 vane/blade domains. The full three-dimensional computations are of the blade having a clearance gap of 2 percent the span. Comparison between the time averaged unsteady and steady heat transfer is provided. It is shown that there is a significant difference between the steady and time mean of unsteady blade heat transfer in localized regions. The differences on the suction side of the blade in the near hub and near tip regions were found to be rather significant. Steady analysis underestimated the blade heat transfer by as much as 20 percent as compared to the time average obtained from the unsteady analysis. As for the blade tip, the steady analysis and the unsteady analysis gave results to within 2 percent.

Exit blade geometry and part-load performance of small axial flow propeller turbines: An experimental investigation

DOE Office of Scientific and Technical Information (OSTI.GOV)

Singh, Punit; Nestmann, Franz

2010-09-15

A detailed experimental investigation of the effects of exit blade geometry on the part-load performance of low-head, axial flow propeller turbines is presented. Even as these turbines find important applications in small-scale energy generation using micro-hydro, the relationship between the layout of blade profile, geometry and turbine performance continues to be poorly characterized. The experimental results presented here help understand the relationship between exit tip angle, discharge through the turbine, shaft power, and efficiency. The modification was implemented on two different propeller runners and it was found that the power and efficiency gains from decreasing the exit tip angle couldmore » be explained by a theoretical model presented here based on classical theory of turbomachines. In particular, the focus is on the behaviour of internal parameters like the runner loss coefficient, relative flow angle at exit, mean axial flow velocity and net tangential flow velocity. The study concluded that the effects of exit tip modification were significant. The introspective discussion on the theoretical model's limitation and test facility suggests wider and continued experimentation pertaining to the internal parameters like inlet vortex profile and exit swirl profile. It also recommends thorough validation of the model and its improvement so that it can be made capable for accurate characterization of blade geometric effects. (author)« less

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

NASA Technical Reports Server (NTRS)

Shyam, Vikram; Ameri, Ali; Chen, Jen-Ping

2010-01-01

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.

Wind Turbine Blade Design System - Aerodynamic and Structural Analysis

NASA Astrophysics Data System (ADS)

Dey, Soumitr

2011-12-01

The ever increasing need for energy and the depletion of non-renewable energy resources has led to more advancement in the "Green Energy" field, including wind energy. An improvement in performance of a Wind Turbine will enhance its economic viability, which can be achieved by better aerodynamic designs. In the present study, a design system that has been under development for gas turbine turbomachinery has been modified for designing wind turbine blades. This is a very different approach for wind turbine blade design, but will allow it to benefit from the features inherent in the geometry flexibility and broad design space of the presented system. It starts with key overall design parameters and a low-fidelity model that is used to create the initial geometry parameters. The low-fidelity system includes the axisymmetric solver with loss models, T-Axi (Turbomachinery-AXIsymmetric), MISES blade-to-blade solver and 2D wing analysis code XFLR5. The geometry parameters are used to define sections along the span of the blade and connected to the CAD model of the wind turbine blade through CAPRI (Computational Analysis PRogramming Interface), a CAD neutral API that facilitates the use of parametric geometry definition with CAD. Either the sections or the CAD geometry is then available for CFD and Finite Element Analysis. The GE 1.5sle MW wind turbine and NERL NASA Phase VI wind turbine have been used as test cases. Details of the design system application are described, and the resulting wind turbine geometry and conditions are compared to the published results of the GE and NREL wind turbines. A 2D wing analysis code XFLR5, is used for to compare results from 2D analysis to blade-to-blade analysis and the 3D CFD analysis. This kind of comparison concludes that, from hub to 25% of the span blade to blade effects or the cascade effect has to be considered, from 25% to 75%, the blade acts as a 2d wing and from 75% to the tip 3D and tip effects have to be taken into account for design considerations. In addition, the benefits of this approach for wind turbine design and future efforts are discussed.

Transonic rotor tip design using numerical optimization

NASA Technical Reports Server (NTRS)

Tauber, Michael E.; Langhi, Ronald G.

1985-01-01

The aerodynamic design procedure for a new blade tip suitable for operation at transonic speeds is illustrated. For the first time, 3 dimensional numerical optimization was applied to rotor tip design, using the recent derivative of the ROT22 code, program R22OPT. Program R22OPT utilized an efficient quasi-Newton optimization algorithm. Multiple design objectives were specified. The delocalization of the shock wave was to be eliminated in forward flight for an advance ratio of 0.41 and a tip Mach number of 0.92 at psi = 90 deg. Simultaneously, it was sought to reduce torque requirements while maintaining effective restoring pitching moments. Only the outer 10 percent of the blade span was modified; the blade area was not to be reduced by more than 3 percent. The goal was to combine the advantages of both sweptback and sweptforward blade tips. A planform that featured inboard sweepback was combined with a sweptforward tip and a taper ratio of 0.5. Initially, the ROT22 code was used to find by trial and error a planform geometry which met the design goals. This configuration had an inboard section with a leading edge sweep of 20 deg and a tip section swept forward at 25 deg; in addition, the airfoils were modified.

Navier-Stokes analysis of an oxidizer turbine blade with tip clearance with and without a mini-shroud

NASA Technical Reports Server (NTRS)

Chan, Tony; Dejong, Frederik J.

1993-01-01

The Gas Generator Oxidizer Turbine (GGOT) Blade is being analyzed by various investigators under the NASA MSFC-sponsored Turbine Stage Technology Team design effort. The present work concentrates on the tip clearance region flow and associated losses; however, flow details for the passage region are also obtained in the simulations. The present calculations simulate the rotor blade row in a rotating reference frame with the appropriate coriolis and centrifugal acceleration term included in the momentum equations. The upstream computational boundary is located about one axial chord from the blade leading edge. The boundary conditions at this location have been determined by Pratt & Whitney using an Euler analysis without the vanes to obtain approximately the same flow profiles at the rotor as were obtained with the Euler stage analysis including the vanes. Inflow boundary layer profiles are then constructed assuming the skin friction coefficient at both the hub and the casing. The downstream computational boundary is located about one axial chord from the blade trailing edge, and the circumferentially averaged static pressure at this location was also obtained from the P&W Euler analysis. Results obtained for the 3-D baseline GGOT geometry at the full scale design Reynolds number show a region of high loss in the region near the casing. Particle traces in the near tip region show vortical flow behavior of the fluid which passes through the clearance region and exits at the downstream edge of the gap. In an effort to reduce clearance flow losses, the mini-shroud concept was proposed by the Pratt & Whitney design team. Calculations were performed on the GGO geometry with the mini-shroud. Results of these calculations indicate that the mini-shroud does not significantly affect the flow in the passage region, and although the tip clearance flow is different, the mini-shroud does not seem to prevent the above-mentioned vortical flow behavior. Since both flow distortion and total pressure losses are similar for both geometries, the addition of the mini-shroud does not seem to reduce the tip clearance flow effects.

"Fan-Tip-Drive" High-Power-Density, Permanent Magnet Electric Motor and Test Rig Designed for a Nonpolluting Aircraft Propulsion Program

NASA Technical Reports Server (NTRS)

Brown, Gerald V.; Kascak, Albert F.

2004-01-01

A scaled blade-tip-drive test rig was designed at the NASA Glenn Research Center. The rig is a scaled version of a direct-current brushless motor that would be located in the shroud of a thrust fan. This geometry is very attractive since the allowable speed of the armature is approximately the speed of the blade tips (Mach 1 or 1100 ft/s). The magnetic pressure generated in the motor acts over a large area and, thus, produces a large force or torque. This large force multiplied by the large velocity results in a high-power-density motor.

Structural Tailoring of Advanced Turboprops (STAT). Theoretical manual

NASA Technical Reports Server (NTRS)

Brown, K. W.

1992-01-01

This manual describes the theories in the Structural Tailoring of Advanced Turboprops (STAT) computer program, which was developed to perform numerical optimizations on highly swept propfan blades. The optimization procedure seeks to minimize an objective function, defined as either direct operating cost or aeroelastic differences between a blade and its scaled model, by tuning internal and external geometry variables that must satisfy realistic blade design constraints. The STAT analyses include an aerodynamic efficiency evaluation, a finite element stress and vibration analysis, an acoustic analysis, a flutter analysis, and a once-per-revolution (1-p) forced response life prediction capability. The STAT constraints include blade stresses, blade resonances, flutter, tip displacements, and a 1-P forced response life fraction. The STAT variables include all blade internal and external geometry parameters needed to define a composite material blade. The STAT objective function is dependent upon a blade baseline definition which the user supplies to describe a current blade design for cost optimization or for the tailoring of an aeroelastic scale model.

Structural Tailoring of Advanced Turboprops (STAT). Theoretical manual

NASA Astrophysics Data System (ADS)

Brown, K. W.

1992-10-01

This manual describes the theories in the Structural Tailoring of Advanced Turboprops (STAT) computer program, which was developed to perform numerical optimizations on highly swept propfan blades. The optimization procedure seeks to minimize an objective function, defined as either direct operating cost or aeroelastic differences between a blade and its scaled model, by tuning internal and external geometry variables that must satisfy realistic blade design constraints. The STAT analyses include an aerodynamic efficiency evaluation, a finite element stress and vibration analysis, an acoustic analysis, a flutter analysis, and a once-per-revolution (1-p) forced response life prediction capability. The STAT constraints include blade stresses, blade resonances, flutter, tip displacements, and a 1-P forced response life fraction. The STAT variables include all blade internal and external geometry parameters needed to define a composite material blade. The STAT objective function is dependent upon a blade baseline definition which the user supplies to describe a current blade design for cost optimization or for the tailoring of an aeroelastic scale model.

Studies in a transonic rotor aerodynamics and noise facility

NASA Technical Reports Server (NTRS)

Wright, S. E.; Lee, D. J.; Crosby, W.

1984-01-01

The design, construction and testing of a transonic rotor aerodynamics and noise facility was undertaken, using a rotating arm blade element support technique. This approach provides a research capability intermediate between that of a stationary element in a moving flow and that of a complete rotating blade system, and permits the acoustic properties of blade tip elements to be studied in isolation. This approach is an inexpensive means of obtaining data at high subsonic and transonic tip speeds on the effect of variations in tip geometry. The facility may be suitable for research on broad band noise and discrete noise in addition to high-speed noise. Initial tests were conducted over the Mach number range 0.3 to 0.93 and confirmed the adequacy of the acoustic treatment used in the facility to avoid reflection from the enclosure.

NPSS Multidisciplinary Integration and Analysis

NASA Technical Reports Server (NTRS)

Hall, Edward J.; Rasche, Joseph; Simons, Todd A.; Hoyniak, Daniel

2006-01-01

The objective of this task was to enhance the capability of the Numerical Propulsion System Simulation (NPSS) by expanding its reach into the high-fidelity multidisciplinary analysis area. This task investigated numerical techniques to convert between cold static to hot running geometry of compressor blades. Numerical calculations of blade deformations were iteratively done with high fidelity flow simulations together with high fidelity structural analysis of the compressor blade. The flow simulations were performed with the Advanced Ducted Propfan Analysis (ADPAC) code, while structural analyses were performed with the ANSYS code. High fidelity analyses were used to evaluate the effects on performance of: variations in tip clearance, uncertainty in manufacturing tolerance, variable inlet guide vane scheduling, and the effects of rotational speed on the hot running geometry of the compressor blades.

A Novel Method for Reducing Rotor Blade-Vortex Interaction

NASA Technical Reports Server (NTRS)

Glinka, A. T.

2000-01-01

One of the major hindrances to expansion of the rotorcraft market is the high-amplitude noise they produce, especially during low-speed descent, where blade-vortex interactions frequently occur. In an attempt to reduce the noise levels caused by blade-vortex interactions, the flip-tip rotor blade concept was devised. The flip-tip rotor increases the miss distance between the shed vortices and the rotor blades, reducing BVI noise. The distance is increased by rotating an outboard portion of the rotor tip either up or down depending on the flight condition. The proposed plan for the grant consisted of a computational simulation of the rotor aerodynamics and its wake geometry to determine the effectiveness of the concept, coupled with a series of wind tunnel experiments exploring the value of the device and validating the computer model. The computational model did in fact show that the miss distance could be increased, giving a measure of the effectiveness of the flip-tip rotor. However, the wind experiments were not able to be conducted. Increased outside demand for the 7'x lO' wind tunnel at NASA Ames and low priority at Ames for this project forced numerous postponements of the tests, eventually pushing the tests beyond the life of the grant. A design for the rotor blades to be tested in the wind tunnel was completed and an analysis of the strength of the model blades based on predicted loads, including dynamic forces, was done.

Effects of a Forward-swept Front Rotor on the Flowfield of a Counterrotation Propeller

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Podboy, Gary G.

1994-01-01

The effects of a forward-swept front rotor on the flowfield of a counterrotation model propeller at takeoff conditions at zero degree angle of attack are studied by solving the unsteady three-dimensional Euler equations. The configuration considered is an uneven blade count counterrotation model with twelve forward-swept blades on the fore rotor and ten aft-swept blades on the aft rotor. The flowfield is compared with that of a reference aft-swept counterrotation geometry and Laser Doppler Velocimeter (LDV) measurements. At the operating conditions considered, the forward-swept blade experiences a higher tip loading and produces a stronger tip vortex compared to the aft-swept blade, consistent with the LDV and acoustic measurements. Neither the solution nor the LDV data indicated the formation of a leading edge vortex. The predicted radial distribution of the circumferentially averaged axial velocity at the measurement station agreed very closely with LDV data, while crossflow velocities showed poor agreement. The discrepancy between prediction and LDV data of tangential and radial velocities is due in part to the insufficient mesh resolution in the region between the rotors and in the tip region to track the tip vortex. The vortex is diffused by the time it arrives at the measurement station. The uneven blade count configuration requires the solution to be carried out for six blade passages of the fore rotor and five passages of the aft rotor, thus making grid refinement prohibitive.

Unsteady Analysis of Blade and Tip Heat Transfer as Influenced by the Upstream Momentum and Thermal Wakes

NASA Technical Reports Server (NTRS)

Ameri, Ali A.; Rigby, David L.; Steinthorsson, Erlendur; Heidmann, James D.; Fabian, John C.

2008-01-01

The effect of the upstream wake on the blade heat transfer has been numerically examined. The geometry and the flow conditions of the first stage turbine blade of GE s E3 engine with a tip clearance equal to 2 percent of the span was utilized. Based on numerical calculations of the vane, a set of wake boundary conditions were approximated, which were subsequently imposed upon the downstream blade. This set consisted of the momentum and thermal wakes as well as the variation in modeled turbulence quantities of turbulence intensity and the length scale. Using a one-blade periodic domain, the distributions of unsteady heat transfer rate on the turbine blade and its tip, as affected by the wake, were determined. Such heat transfer coefficient distribution was computed using the wall heat flux and the adiabatic wall temperature to desensitize the heat transfer coefficient to the wall temperature. For the determination of the wall heat flux and the adiabatic wall temperatures, two sets of computations were required. The results were used in a phase-locked manner to compute the unsteady or steady heat transfer coefficients. It has been found that the unsteady wake has some effect on the distribution of the time averaged heat transfer coefficient on the blade and that this distribution is different from the distribution that is obtainable from a steady computation. This difference was found to be as large as 20 percent of the average heat transfer on the blade surface. On the tip surface, this difference is comparatively smaller and can be as large as four percent of the average.

A comprehensive method for preliminary design optimization of axial gas turbine stages. II - Code verification

NASA Technical Reports Server (NTRS)

Jenkins, R. M.

1983-01-01

The present effort represents an extension of previous work wherein a calculation model for performing rapid pitchline optimization of axial gas turbine geometry, including blade profiles, is developed. The model requires no specification of geometric constraints. Output includes aerodynamic performance (adiabatic efficiency), hub-tip flow-path geometry, blade chords, and estimates of blade shape. Presented herein is a verification of the aerodynamic performance portion of the model, whereby detailed turbine test-rig data, including rig geometry, is input to the model to determine whether tested performance can be predicted. An array of seven (7) NASA single-stage axial gas turbine configurations is investigated, ranging in size from 0.6 kg/s to 63.8 kg/s mass flow and in specific work output from 153 J/g to 558 J/g at design (hot) conditions; stage loading factor ranges from 1.15 to 4.66.

Quantifying uncertainties in the structural response of SSME blades

NASA Technical Reports Server (NTRS)

Nagpal, Vinod K.

1987-01-01

To quantify the uncertainties associated with the geometry and material properties of a Space Shuttle Main Engine (SSME) turbopump blade, a computer code known as STAEBL was used. A finite element model of the blade used 80 triangular shell elements with 55 nodes and five degrees of freedom per node. The whole study was simulated on the computer and no real experiments were conducted. The structural response has been evaluated in terms of three variables which are natural frequencies, root (maximum) stress, and blade tip displacements. The results of the study indicate that only the geometric uncertainties have significant effects on the response. Uncertainties in material properties have insignificant effects.

Investigation and Optimization of Blade Tip Winglets Using an Implicit Free Wake Vortex Method

NASA Astrophysics Data System (ADS)

Lawton, Stephen; Crawford, Curran

2014-06-01

Novel outer-blade geometries such as tip winglets can increase the aerodynamic power that can be extracted from the wind by tailoring the relative position and strengths of trailed vorticity. This design space is explored using both parameter studies and gradient-based optimization, with the aerodynamic analysis carried out using LibAero, a free wake vortex-based code introduced in previous work. The starting design is the NREL 5MW reference turbine, which allows comparison of the aerodynamic simulation for the unmodified blade with other codes. The code uses a Prandtl-Weissinger lifting line model to represent the blade, and vortex filaments as the flow elements. A fast multipole method is implemented to accelerate the influence calculations and reduce the computational cost. This results in higher fidelity aerodynamic simulations that can capture the effects of novel geometries while maintaining sufficiently fast run-times (on the order of an hour) to allow the use of optimization. Gradients of the objective function with respect to design variables are calculated using the complex step method which is accurate and efficient. Since the vortex structure behind the rotor is being resolved in detail, insight is also gained into the mechanisms by which these new blade designs affect performance. It is found that adding winglets can increase the power extracted from the wind by around 2%, with a similar increase in thrust. It is also possible to create a winglet that slightly lowers the thrust while maintaining very similar power compared to the standard straight blade.

Trailing Vortex Measurements in the Wake of a Hovering Rotor Blade with Various Tip Shapes

NASA Technical Reports Server (NTRS)

Martin, Preston B.; Leishman, J. Gordon

2003-01-01

This work examined the wake aerodynamics of a single helicopter rotor blade with several tip shapes operating on a hover test stand. Velocity field measurements were conducted using three-component laser Doppler velocimetry (LDV). The objective of these measurements was to document the vortex velocity profiles and then extract the core properties, such as the core radius, peak swirl velocity, and axial velocity. The measured test cases covered a wide range of wake-ages and several tip shapes, including rectangular, tapered, swept, and a subwing tip. One of the primary differences shown by the change in tip shape was the wake geometry. The effect of blade taper reduced the initial peak swirl velocity by a significant fraction. It appears that this is accomplished by decreasing the vortex strength for a given blade loading. The subwing measurements showed that the interaction and merging of the subwing and primary vortices created a less coherent vortical structure. A source of vortex core instability is shown to be the ratio of the peak swirl velocity to the axial velocity deficit. The results show that if there is a turbulence producing region of the vortex structure, it will be outside of the core boundary. The LDV measurements were supported by laser light-sheet flow visualization. The results provide several benchmark test cases for future validation of theoretical vortex models, numerical free-wake models, and computational fluid dynamics results.

Blade pitch optimization methods for vertical-axis wind turbines

NASA Astrophysics Data System (ADS)

Kozak, Peter

Vertical-axis wind turbines (VAWTs) offer an inherently simpler design than horizontal-axis machines, while their lower blade speed mitigates safety and noise concerns, potentially allowing for installation closer to populated and ecologically sensitive areas. While VAWTs do offer significant operational advantages, development has been hampered by the difficulty of modeling the aerodynamics involved, further complicated by their rotating geometry. This thesis presents results from a simulation of a baseline VAWT computed using Star-CCM+, a commercial finite-volume (FVM) code. VAWT aerodynamics are shown to be dominated at low tip-speed ratios by dynamic stall phenomena and at high tip-speed ratios by wake-blade interactions. Several optimization techniques have been developed for the adjustment of blade pitch based on finite-volume simulations and streamtube models. The effectiveness of the optimization procedure is evaluated and the basic architecture for a feedback control system is proposed. Implementation of variable blade pitch is shown to increase a baseline turbine's power output between 40%-100%, depending on the optimization technique, improving the turbine's competitiveness when compared with a commercially-available horizontal-axis turbine.

Computational Investigation of Novel Tip Leakage Mitigation Methods for High Pressure Turbine Blades

NASA Technical Reports Server (NTRS)

Ibrahim, Mounir; Gupta, Abhinav; Shyam, Vikram

2014-01-01

This paper presents preliminary findings on a possible approach to reducing tip leakage losses. In this paper a computational study was conducted on the Energy Efficient Engine (EEE) High Pressure Turbine (HPT) rotor tip geometry using the commercial numerical solver ANSYS FLUENT. The flow solver was validated against aerodynamic data acquired in the NASA Transonic Turbine Blade Cascade facility. The scope of the ongoing study is to computationally investigate how the tip leakage and overall blade losses are affected by (1) injection from the tip near the pressure side, (2) injection from the tip surface at the camber line, and (3) injection from the tip surface into the tip separation bubble. The objective is to identify the locations on the tip surface at which to place appropriately configured blowing keeping in mind the film cooling application of tip blowing holes. The validation was conducted at Reynolds numbers of 85,000, 343,000, and 685,000 and at engine realistic flow conditions. The coolant injection simulations were conducted at a Reynolds number of 343,000 based on blade chord and inlet velocity and utilized the SST turbulence model in FLUENT. The key parameters examined are the number of jets, jet angle and jet location. A coolant to inlet pressure ratio of 1.0 was studied for angles of +30 deg, -30 deg, and 90 deg to the local free stream on the tip. For the 3 hole configuration, 3 holes spaced 3 hole diameters apart with length to diameter ratio of 1.5 were used. A simulation including 11 holes along the entire mean camber line is also presented (30 deg toward suction side). In addition, the effect of a single hole is also compared to a flat tip with no injection. The results provide insight into tip flow control methods and can be used to guide further investigation into tip flow control. As noted in past research it is concluded that reducing leakage flow is not necessarily synonymous with reducing losses due to leakage.

Computational Investigation of Novel Tip Leakage Mitigation Methods for High Pressure Turbine Blades

NASA Technical Reports Server (NTRS)

Ibrahim, Mounir; Gupta, Abhinav; Shyam, Vikram

2014-01-01

This paper presents preliminary findings on a possible approach to reducing tip leakage losses. In this paper a computational study was conducted on the EEE (Energy Efficient Engine) HPT (High Pressure Turbine) rotor tip geometry using the commercial numerical solver ANSYS FLUENT. The flow solver was validated against aerodynamic data acquired in the NASA Transonic Turbine Blade Cascade facility. The scope of the ongoing study is to computationally investigate how the tip leakage and overall blade losses are affected by 1. injection from the tip near the pressure side, 2. injection from the tip surface at the camber line, and 3. injection from the tip surface into the tip separation bubble. The objective is to identify the locations on the tip surface at which to place appropriately configured blowing keeping in mind the film cooling application of tip blowing holes. The validation was conducted at Reynolds numbers of 85,000, 343,000 and 685,000 and at engine realistic flow conditions. The coolant injection simulations were conducted at a Reynolds number of 343,000 based on blade chord and inlet velocity and utilized the SST turbulence model in FLUENT. The key parameters examined are the number of jets, jet angle and jet location. A coolant to inlet pressure ratio of 1.0 was studied for angles of +30 deg., -30 deg. and 90 deg. to the local free stream on the tip. For the 3 hole configuration, 3 holes spaced 3 hole diameters apart with length to diameter ratio of 1.5 were used. A simulation including 11 holes along the entire mean camber line is also presented (30 degrees toward suction side). In addition, the effect of a single hole is also compared to a flat tip with no injection. The results provide insight into tip flow control methods and can be used to guide further investigation into tip flow control. As noted in past research it is concluded that reducing leakage flow is not necessarily synonymous with reducing losses due to leakage.

Hover performance tests of full scale variable geometry rotors

NASA Technical Reports Server (NTRS)

Rorke, J. B.

1976-01-01

Full scale whirl tests were conducted to determine the effects of interblade spatial relationships and pitch variations on the hover performance and acoustic signature of a 6-blade main rotor system. The variable geometry rotor (VGR) variations from the conventional baseline were accomplished by: (1) shifting the axial position of alternate blades by one chord-length to form two tip path planes; and (2) varying the relative azimuthal spacing from the upper rotor to the lagging hover rotor in four increments from 25.2 degrees to 62.1 degrees. For each of these four configurations, the differential collective pitch between upper and lower rotors was set at + or - 1 deg, 0 deg and -1 deg. Hover performance data for all configurations were acquired at blade tip Mach numbers of 0.523 and 0.45. Acoustic data were recorded at all test conditions, but analyzed only at 0 deg differential pitch at the higher rotor speed. The VGR configurations tested demonstrated improvements in thrust at constant power as high as 6 percent. Reductions of 3 PNdb in perceived noise level and of 4 db in blade passage frequency noise level were achieved at the higher thrust levels. Consistent correlation exists between performance and acoustic improvements. For any given azimuth spacing, performance was consistently better for the differential pitch condition of + or - 1 degree, i.e. with the upper rotor pitch one degree higher than the lower rotor.

Measurements of Tip Vortices from a Full-Scale UH-60A Rotor by Retro- Reflective Background Oriented Schlieren and Stereo Photogrammetry

NASA Technical Reports Server (NTRS)

Schairer, Edward; Kushner, Laura K.; Heineck, James T.

2013-01-01

Positions of vortices shed by a full-scale UH-60A rotor in forward flight were measured during a test in the National Full- Scale Aerodynamics Complex at NASA Ames Research Center. Vortices in a region near the tip of the advancing blade were visualized from two directions by Retro-Reflective Background-Oriented Schlieren (RBOS). Correspondence of points on the vortex in the RBOS images from both cameras was established using epipolar geometry. The object-space coordinates of the vortices were then calculated from the image-plane coordinates using stereo photogrammetry. One vortex from the tip of the blade that had most recently passed was visible in most of the data. The visibility of the vortices was greatest at high thrust and low advance ratios. At these favorable conditions, vortices from the most recent passages of all four blades were detected. The vortex positions were in good agreement with PIV data for a case where PIV measurements were also made. RBOS and photogrammetry provided measurements of the angle at which each vortex passed through the PIV plane.

Helicopter far-field acoustic levels as a function of reduced main-rotor advancing blade-tip Mach number

NASA Technical Reports Server (NTRS)

Mueller, Arnold W.; Smith, Charles D.; Lemasurier, Philip

1990-01-01

During the design of a helicopter, the weight, engine, rotor speed, and rotor geometry are given significant attention when considering the specific operations for which the helicopter will be used. However, the noise radiated from the helicopter and its relationship to the design variables is currently not well modeled with only a limited set of full-scale field test data to study. In general, limited field data have shown that reduced main-rotor advancing blade-tip Mach numbers result in reduced far-field noise levels. The status of a recent helicopter noise research project is reviewed. It is designed to provide flight experimental data which may be used to further understand helicopter main-rotor advancing blade-tip Mach number effects on far-field acoustic levels. Preliminary results are presented relative to tests conducted with a Sikorsky S-76A helicopter operating with both the rotor speed and the flight speed as the control variable. The rotor speed was operated within the range of 107 to 90 percent NR at nominal forward speeds of 35, 100, and 155 knots.

Measurement of noise and its correlation to performance and geometry of small aircraft propellers

NASA Astrophysics Data System (ADS)

Štorch, Vít; Nožička, Jiří; Brada, Martin; Gemperle, Jiří; Suchý, Jakub

2016-03-01

A set of small model and UAV propellers is measured both in terms of aerodynamic performance and acoustic noise under static conditions. Apart from obvious correlation of noise to tip speed and propeller diameter the influence of blade pitch, blade pitch distribution, efficiency and shape of the blade is sought. Using the measured performance data a computational model for calculation of aerodynamic noise of propellers will be validated. The range of selected propellers include both propellers designed for nearly static conditions and propellers that are running at highly offdesign conditions, which allows to investigate i.e. the effect of blade stall on both noise level and performance results.

MPT Prediction of Aircraft-Engine Fan Noise

NASA Technical Reports Server (NTRS)

Connell, Stuart D.

2004-01-01

A collection of computer programs has been developed that implements a procedure for predicting multiple-pure-tone (MPT) noise generated by fan blades of an aircraft engine (e.g., a turbofan engine). MPT noise arises when the fan is operating with supersonic relative tip Mach No. Under this flow condition, there is a strong upstream running shock. The strength and position of this shock are very sensitive to blade geometry variations. For a fan where all the blades are identical, the primary tone observed upstream of the fan will be the blade passing frequency. If there are small variations in geometry between blades, then tones below the blade passing frequency arise MPTs. Stagger angle differences as small as 0.1 can give rise to significant MPT. It is also noted that MPT noise is more pronounced when the fan is operating in an unstarted mode. Computational results using a three-dimensional flow solver to compute the complete annulus flow with non-uniform fans indicate that MPT noise can be estimated in a relatively simple way. Hence, once the effect of a typical geometry variation of one blade in an otherwise uniform blade row is known, the effect of all the blades being different can be quickly computed via superposition. Two computer programs that were developed as part of this work are used in conjunction with a user s computational fluid dynamics (CFD) code to predict MPT spectra for a fan with a specified set of geometric variations: (1) The first program ROTBLD reads the users CFD solution files for a single blade passage via an API (Application Program Interface). There are options to replicate and perturb the geometry with typical variations stagger, camber, thickness, and pitch. The multi-passage CFD solution files are then written in the user s file format using the API. (2) The second program SUPERPOSE requires two input files: the first is the circumferential upstream pressure distribution extracted from the CFD solution on the multi-passage mesh, the second file defines the geometry variations of each blade in a complete fan. Superposition is used to predict the spectra resulting from the geometric variations.

Turbine blade squealer tip rail with fence members

DOE Office of Scientific and Technical Information (OSTI.GOV)

Little, David A

2012-11-20

A turbine blade includes an airfoil, a blade tip section, a squealer tip rail, and a plurality of chordally spaced fence members. The blade tip section includes a blade tip floor located at an end of the airfoil distal from the root. The blade tip floor includes a pressure side and a suction side joined together at chordally spaced apart leading and trailing edges of the airfoil. The squealer tip rail extends radially outwardly from the blade tip floor adjacent to the suction side and extends from a first location adjacent to the airfoil trailing edge to a second locationmore » adjacent to the airfoil leading edge. The fence members are located between the airfoil leading and trailing edges and extend radially outwardly from the blade tip floor and axially from the squealer tip rail toward the pressure side.« less

The Relationship between Appendage Geometry and Propeller Blade Unsteady Forces.

DTIC Science & Technology

1987-11-01

unsteady thrust and torque for a given propeller geometry. The results indicate that unsteady force reduction can be obtained by modification of the flow ... unsteady force calculation methods available are: 1) quasi-steady using uniform flow ; 2) quasi-steady using lifting-line theory; 3) two-dimensional... experimental data and the calculated unsteady forces that both the flow field near the body surface and behind the appendage tip must be

Computational investigation of flow control by means of tubercles on Darrieus wind turbine blades

NASA Astrophysics Data System (ADS)

Sevinç, K.; Özdamar, G.; Şentürk, U.; Özdamar, A.

2015-09-01

This work presents the current status of the computational study of the boundary layer control of a vertical axis wind turbine blade by modifying the blade geometry for use in wind energy conversion. The control method is a passive method which comprises the implementation of the tubercle geometry of a humpback whale flipper onto the leading edge of the blades. The baseline design is an H-type, three-bladed Darrieus turbine with a NACA 0015 cross-section. Finite-volume based software ANSYS Fluent was used in the simulations. Using the optimum control parameters for a NACA 634-021 profile given by Johari et al. (2006), turbine blades were modified. Three dimensional, unsteady, turbulent simulations for the blade were conducted to look for a possible improvement on the performance. The flow structure on the blades was investigated and flow phenomena such as separation and stall were examined to understand their impact on the overall performance. For a tip speed ratio of 2.12, good agreement was obtained in the validation of the baseline model with a relative error in time- averaged power coefficient of 1.05%. Modified turbine simulations with a less expensive but less accurate turbulence model yielded a decrease in power coefficient. Results are shown comparatively.

Design, analysis, optimization and control of rotor tip flows

NASA Astrophysics Data System (ADS)

Maesschalck, Cis Guy M. De

Developments in turbomachinery focus on efficiency and reliability enhancements, while reducing the production costs. In spite of the many noteworthy experimental and numerical investigations over the past decades, the turbine tip design presents numerous challenges to the engine manufacturers, and remains the primary factor defining the machine durability for the periodic removal of the turbine components during overhaul. Due to the hot gases coming from the upstream combustion chamber, the turbine blades are subjected to temperatures far above the metal creep temperature, combined with severe thermal stresses induced within the blade material. Inadequate designs cause early tip burnouts leading to considerable performance degradations, or even a catastrophic turbine failure. Moreover, the leakage spillage, nowadays often exceeding the transonic regime, generates large aerodynamic penalties which are responsible for about one third of the turbine losses. In this view, the current doctoral research exploits the potential through the modification and optimization of the blade tip shape as a means to control the tip leakage flow aerodynamics and manage the heat load distribution over the blade profile to improve the turbine efficiency and durability. Three main design strategies for unshrouded turbine blade tips were analyzed and optimized: tight running clearances, blade tip contouring and the use of complex squealer-like geometries. The altered overtip flow physics and heat transfer characteristics were simulated for tight gap sizes as low as 0.5% down to 0.1% of the blade height, occurring during engine transients and soon to be expected due to recent developments in active clearance control strategies. The potential of fully 3D contoured blade top surfaces, allowing to adapt the profile locally to the changing flow conditions throughout the camberline, is quantified. First adopting a quasi-3D approach and subsequently using a full 3D optimization. For the industrial rub-safe squealer profiles featuring cavities separated by upstanding rims, a topology-like multi-objective 3D optimization strategy is used to identify so far undiscovered, optimal blade tip profiles. Furthermore, the additional potential of the widely adopted shroud coolant injection just upstream of the rotor blade is examined. Specifically, the possibility of combining the beneficial effect of the purge flow in the overtip region while minimizing the detrimental influence on the upper passage vortex is explored. Eventually, a high-speed rotating turbine facility at the von Karman Institute was redesigned, allowing simultaneous testing of multiple distinct blade (tip) profiles mounted in separate sectors around the rotor annulus. Important considerations related with the balancing and precise clearance design are highlighted, arising from the complexity of such rainbow-rotor configuration. Moreover, approaches are described to integrate Reynolds-Averaged Navier-Stokes simulations to a priori estimate the errors induced by the finite spatial sampling and inherent limited sensor bandwidth. This research effort provided new insights into the overtip flow topology and aerothermal characteristics, identified new design strategies to create future turbines with enhanced aerodynamic efficiencies and reduced thermal loads, and paved the way for an elaborate experimental validation in a rotating turbine facility, at engine-matched conditions.

Optimization of an Active Twist Rotor Blade Planform for Improved Active Response and Forward Flight Performance

NASA Technical Reports Server (NTRS)

Sekula, Martin K; Wilbur, Matthew L.

2014-01-01

A study was conducted to identify the optimum blade tip planform for a model-scale active twist rotor. The analysis identified blade tip design traits which simultaneously reduce rotor power of an unactuated rotor while leveraging aeromechanical couplings to tailor the active response of the blade. Optimizing the blade tip planform for minimum rotor power in forward flight provided a 5 percent improvement in performance compared to a rectangular blade tip, but reduced the vibration control authority of active twist actuation by 75 percent. Optimizing for maximum blade twist response increased the vibration control authority by 50 percent compared to the rectangular blade tip, with little effect on performance. Combined response and power optimization resulted in a blade tip design which provided similar vibration control authority to the rectangular blade tip, but with a 3.4 percent improvement in rotor performance in forward flight.

Advanced turbine blade tip seal system

NASA Technical Reports Server (NTRS)

Zelahy, J. W.

1981-01-01

An advanced blade/shroud system designed to maintain close clearance between blade tips and turbine shrouds and at the same time, be resistant to environmental effects including high temperature oxidation, hot corrosion, and thermal cycling is described. Increased efficiency and increased blade life are attained by using the advanced blade tip seal system. Features of the system include improved clearance control when blade tips preferentially wear the shrouds and a superior single crystal superalloy tip. The tip design, joint location, characterization of the single crystal tip alloy, the abrasive tip treatment, and the component and engine test are among the factors addressed. Results of wear testing, quality control plans, and the total manufacturing cycle required to fully process the blades are also discussed.

Predesign study for a modern 4-bladed rotor for RSRA

NASA Technical Reports Server (NTRS)

Davis, S. J.

1981-01-01

The feasibility of providing a modern four-bladed rotor for flight research testing on a rotor system aircraft was evaluated. The capabilities of a state of the art rotor system and the contributions of key design parameters to these capabilities were investigated. Three candidate rotors were examined: the UH-60A BLACK HAWK rotor with and without root extenders and the H-3 composite blade rotor. It was concluded that the technical/cost objectives could best be accomplished using the basic BLACK HAWK rotor (i.e. without root extenders). Further, the availability of three existing sets of blade tip of varying design, together with a demonstrated capability for altering airfoil geometry should provide early research information on important design variables at reduced cost.

A transport model for the deterministic stresses associated with turbomachinery blade row interactions

NASA Astrophysics Data System (ADS)

van de Wall, Allan George

The unsteady process resulting from the interaction of upstream vortical structures with a downstream blade row in turbomachines can have a significant impact on the machine efficiency. A transport model assuming incompressible flow and using linear theory was developed to take this process into account in the computation of time-average multistage turbomachinery flows. The upstream vortical structures are transported by the mean flow of the downstream blade row, redistributing the time-average unsteady kinetic energy (Uke ) associated with the incoming disturbance. The model was applied to compressor and turbine geometry. For compressors, the Uke associated with upstream 2-D wakes and 3-D tip clearance flows is reduced as a result of the interaction with a downstream blade row. This reduction results from inviscid effects as well as viscous effects and reduces the loss associated with the upstream disturbance. Any disturbance passing through a compressor blade row results in a smaller loss than if the disturbance was mixed-out prior to entering the blade row. For turbines, the Uke associated with upstream 2-D wakes and 3-D tip clearance flows are significantly amplified by inviscid effects as a result of the interaction with a downstream turbine blade row. Viscous effects act to reduce the amplification of the Uke by inviscid effects but results in a substantial loss. Any disturbance passing through a turbine blade row results in a larger loss than if the disturbance was mixedout prior to entering the blade row.

Navier-Stokes analysis of an oxidizer turbine blade with tip clearance

NASA Technical Reports Server (NTRS)

Gibeling, Howard J.; Sabnis, Jayant S.

1992-01-01

The Gas Generator Oxidizer Turbine (GGOT) Blade is being analyzed by various investigators under the NASA MSFC sponsored Turbine Stage Technology Team design effort. The present work concentrates on the tip clearance region flow and associated losses; however, flow details for the passage region are also obtained in the simulations. The present calculations simulate the rotor blade row in a rotating reference frame with the appropriate coriolis and centrifugal acceleration terms included in the momentum equation. The upstream computational boundary is located about one axial chord from the blade leading edge. The boundary conditions at this location were determined by using a Euler analysis without the vanes to obtain approximately the same flow profiles at the rotor as were obtained with the Euler stage analysis including the vanes. Inflow boundary layer profiles are then constructed assuming the skin friction coefficient at both the hub and the casing. The downstream computational boundary is located about one axial chord from the blade trailing edge, and the circumferentially averaged static pressure at this location was also obtained from the Euler analysis. Results were obtained for the 3-D baseline GGOT geometry at the full scale design Reynolds number. Details of the clearance region flow behavior and blade pressure distributions were computed. The spanwise variation in blade loading distributions are shown, and circumferentially averaged spanwise distributions of total pressure, total temperature, Mach number, and flow angle are shown at several axial stations. The spanwise variation of relative total pressure loss shows a region of high loss in the region near the casing. Particle traces in the near tip region show vortical behavior of the fluid which passes through the clearance region and exits at the downstream edge of the gap.

Flowfield and heat transfer past an unshrouded gas turbine blade tip with different shapes

NASA Astrophysics Data System (ADS)

Liu, Jian-Jun; Li, Peng; Zhang, Chao; An, Bai-Tao

2013-04-01

This paper describes the numerical investigations of flow and heat transfer in an unshrouded turbine rotor blade of a heavy duty gas turbine with four tip configurations. By comparing the calculated contours of heat transfer coefficients on the flat tip of the HP turbine rotor blade in the GE-E3 aircraft engine with the corresponding experimental data, the κ-ω turbulence model was chosen for the present numerical simulations. The inlet and outlet boundary conditions for the turbine rotor blade are specified as the real gas turbine, which were obtained from the 3D full stage simulations. The rotor blade and the hub endwall are rotary and the casing is stationary. The influences of tip configurations on the tip leakage flow and blade tip heat transfer were discussed. It's showed that the different tip configurations changed the leakage flow patterns and the pressure distributions on the suction surface near the blade tip. Compared with the flat tip, the total pressure loss caused by the leakage flow was decreased for the full squealer tip and pressure side squealer tip, while increased for the suction side squealer tip. The suction side squealer tip results in the lowest averaged heat transfer coefficient on the blade tip compared to the other tip configurations.

Aeroelastic Stability of Modern Bearingless Rotors: A Parametric Investigation

NASA Technical Reports Server (NTRS)

Nguyen, Khanh Q.

1994-01-01

The University of Maryland Advanced Rotorcraft Code (UMARC) is utilized to study the effects of blade design parameters on the aeroelastic stability of an isolated modern bearingless rotor blade in hover. The McDonnell Douglas Advanced Rotor Technology (MDART) Rotor is the baseline rotor investigated. Results indicate that kinematic pitch-lag coupling introduced through the control system geometry and the damping levels of the shear lag dampers strongly affect the hover inplane damping of the baseline rotor blade. Hub precone, pitchcase chordwise stiffness, and blade fundamental torsion frequency have small to moderate influence on the inplane damping, while blade pre-twist and placements of blade fundamental flapwise and chord-wise frequencies have negligible effects. A damperless configuration with a leading edge pitch-link, 15 deg of pitch-link cant angle, and reduced pitch-link stiffness is shown to be stable with an inplane damping level in excess of 2.7 percent critical at the full hover tip speed.

Smart helicopter rotor with active blade tips

NASA Astrophysics Data System (ADS)

Bernhard, Andreas Paul Friedrich

2000-10-01

The smart active blade tip (SABT) rotor is an on-blade rotor vibration reduction system, incorporating active blade tips that can be independently pitched with respect to the main blade. The active blade tip rotor development included an experimental test program culminating in a Mach scale hover test, and a parallel development of a coupled, elastic actuator and rotor blade analysis for preliminary design studies and hover performance prediction. The experimental testing focussed on a small scale rotor on a bearingless Bell-412 hub. The fabricated Mach-scale active-tip rotor has a diameter of 1.524 m, a blade chord of 76.2 mm and incorporated a 10% span active tip. The nominal operating speed is 2000 rpm, giving a tip Mach number of 0.47. The blade tips are driven by a novel piezo-induced bending-torsion coupled actuator beam, located spanwise in the hollow mid-cell of the main rotor blade. In hover at 2000 rpm, at 2 deg collective, and for an actuation of 125 Vrms, the measured blade tip deflection at the first four rotor harmonics is between +/-1.7 and +/-2.8 deg, increasing to +/-5.3 deg at 5/rev with resonant amplification. The corresponding oscillatory amplitude of the rotor thrust coefficient is between 0.7 · 10-3 and 1.3 · 10-1 at the first four rotor harmonics, increasing to 2.1 · 10-3 at 5/rev. In general, the experimental blade tip frequency response and corresponding rotor thrust response are well captured by the analysis. The flexbeam root flap bending moment is predicted in trend, but is significantly over-estimated. The blade tips did not deflect as expected at high collective settings, because of the blade tip shaft locking up in the bearing. This is caused by the high flap bending moment on the blade tip shaft. Redesign of the blade tip shaft assembly and bearing support is identified as the primary design improvement for future research. The active blade tip rotor was also used as a testbed for the evaluation of an adaptive neural-network based control algorithm. Effective background vibration reduction of an intentional 1/rev hover imbalance was demonstrated. The control algorithm also showed the capability to generate desired multi-frequency control loads on the hub, based on artificial signal injection into the vibration measurement. The research program demonstrates the technical feasibility of the active blade tip concept for vibration reduction and warrants further investigation in terms of closed loop forward flight tests in the windtunnel and full scale design studies.

Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade - Part 1: Experimental Measurements. Part 1; Experimental Measurements

NASA Technical Reports Server (NTRS)

Bunker, Ronald S.; Wetzel, Todd G.; Rigby, David L.; Reddy, D. R. (Technical Monitor)

2000-01-01

A combined experimental and computational study has been performed to investigate the detailed heat transfer coefficient distributions within a complex blade trailing edge passage. The experimental measurements are made using a steady liquid crystal thermography technique applied to one major side of the passage. The geometry of the trailing edge passage is that of a two-pass serpentine circuit with a sharp 180-degree turning region at the tip. The upflow channel is split by interrupted ribs into two major subchannels, one of which is turbulated. This channel has an average aspect ratio of roughly 14:1. The spanwise extent of the channel geometry includes both area convergence from root to tip, as well as taper towards the trailing edge apex. The average section Reynolds numbers tested in this upflow channel range from 55,000 to 98,000. The tip section contains a turning vane near the extreme comer. The downflow channel has an aspect ratio of about 5:1, and also includes convergence and taper. Turbulators of varying sizes are included in this channel also. Both detailed heat transfer and pressure distribution measurements are presented. The pressure measurements are incorporated into a flow network model illustrating the major loss contributors.

Numerical Investigation of the Interaction between Mainstream and Tip Shroud Leakage Flow in a 2-Stage Low Pressure Turbine

NASA Astrophysics Data System (ADS)

Jia, Wei; Liu, Huoxing

2014-06-01

The pressing demand for future advanced gas turbine requires to identify the losses in a turbine and to understand the physical mechanisms producing them. In low pressure turbines with shrouded blades, a large portion of these losses is generated by tip shroud leakage flow and associated interaction. For this reason, shroud leakage losses are generally grouped into the losses of leakage flow itself and the losses caused by the interaction between leakage flow and mainstream. In order to evaluate the influence of shroud leakage flow and related losses on turbine performance, computational investigations for a 2-stage low pressure turbine is presented and discussed in this paper. Three dimensional steady multistage calculations using mixing plane approach were performed including detailed tip shroud geometry. Results showed that turbines with shrouded blades have an obvious advantage over unshrouded ones in terms of aerodynamic performance. A loss mechanism breakdown analysis demonstrated that the leakage loss is the main contributor in the first stage while mixing loss dominates in the second stage. Due to the blade-to-blade pressure gradient, both inlet and exit cavity present non-uniform leakage injection and extraction. The flow in the exit cavity is filled with cavity vortex, leakage jet attached to the cavity wall and recirculation zone induced by main flow ingestion. Furthermore, radial gap and exit cavity size of tip shroud have a major effect on the yaw angle near the tip region in the main flow. Therefore, a full calculation of shroud leakage flow is necessary in turbine performance analysis and the shroud geometric features need to be considered during turbine design process.

Online monitoring of dynamic tip clearance of turbine blades in high temperature environments

NASA Astrophysics Data System (ADS)

Han, Yu; Zhong, Chong; Zhu, Xiaoliang; Zhe, Jiang

2018-04-01

Minimized tip clearance reduces the gas leakage over turbine blade tips and improves the thrust and efficiency of turbomachinery. An accurate tip clearance sensor, measuring the dynamic clearances between blade tips and the turbine case, is a critical component for tip clearance control. This paper presents a robust inductive tip clearance sensor capable of monitoring dynamic tip clearances of turbine machines in high-temperature environments and at high rotational speeds. The sensor can also self-sense the temperature at a blade tip in situ such that temperature effect on tip clearance measurement can be estimated and compensated. To evaluate the sensor’s performance, the sensor was tested for measuring the tip clearances of turbine blades under various working temperatures ranging from 700 K to 1300 K and at turbine rotational speeds ranging from 3000 to 10 000 rpm. The blade tip clearance was varied from 50 to 2000 µm. The experiment results proved that the sensor can accurately measure the blade tip clearances with a temporal resolution of 10 µm. The capability of accurately measuring the tip clearances at high temperatures (~1300 K) and high turbine rotation speeds (~30 000 rpm), along with its compact size, makes it promising for online monitoring and active control of blade tip clearances of high-temperature turbomachinery.

77 FR 20518 - Airworthiness Directives; Agusta S.p.A. Helicopters

Federal Register 2010, 2011, 2012, 2013, 2014

2012-04-05

... intended to prevent loss of the blade tip weight, loss of a blade, and subsequent loss of control of the... airworthy blade. This AD is prompted by incidents where a blade tip weight separated from a blade in flight... a blade tip weight separating from a blade in flight, and the subsequent investigation showed that...

Cold-air performance of a 12.766-centimeter-tip-diameter axial-flow cooled turbine. 3: Effect of rotor tip clearance on overall performance of a solid blade configuration

NASA Technical Reports Server (NTRS)

Haas, J. E.; Kofskey, M. G.

1977-01-01

Two tip clearance configurations, one with a recess in the casing and the other with a reduced rotor blade height, were investigated at design equivalent speed over a range of tip clearance from about 2.0 to 5.0 percent of the stator blade height. The optimum configuration with a recess in the casing was the one where the rotor tip diameter was equal to the stator tip diameter (zero blade extension). For this configuration there was an approximate 1.5 percent decrease in total efficiency for an increase in tip clearance of 1 percent of stator blade height. For the reduced blade height configurations there was an approximate 2.0 percent decrease in total efficiency for an increase in tip clearance of 1 percent of stator blade height.

Method to improve the blade tip-timing accuracy of fiber bundle sensor under varying tip clearance

NASA Astrophysics Data System (ADS)

Duan, Fajie; Zhang, Jilong; Jiang, Jiajia; Guo, Haotian; Ye, Dechao

2016-01-01

Blade vibration measurement based on the blade tip-timing method has become an industry-standard procedure. Fiber bundle sensors are widely used for tip-timing measurement. However, the variation of clearance between the sensor and the blade will bring a tip-timing error to fiber bundle sensors due to the change in signal amplitude. This article presents methods based on software and hardware to reduce the error caused by the tip clearance change. The software method utilizes both the rising and falling edges of the tip-timing signal to determine the blade arrival time, and a calibration process suitable for asymmetric tip-timing signals is presented. The hardware method uses an automatic gain control circuit to stabilize the signal amplitude. Experiments are conducted and the results prove that both methods can effectively reduce the impact of tip clearance variation on the blade tip-timing and improve the accuracy of measurements.

Aerodynamics and Heat Transfer Studies of Parameters Specific to the IGCC-Requirements: Endwall Contouring, Leading Edge and Blade Tip Ejection under Rotating Turbine Conditions

DOE Office of Scientific and Technical Information (OSTI.GOV)

Schobeiri, Meinhard; Han, Je-Chin

2014-09-30

This report deals with the specific aerodynamics and heat transfer problematic inherent to high pressure (HP) turbine sections of IGCC-gas turbines. Issues of primary relevance to a turbine stage operating in an IGCC-environment are: (1) decreasing the strength of the secondary flow vortices at the hub and tip regions to reduce (a), the secondary flow losses and (b), the potential for end wall deposition, erosion and corrosion due to secondary flow driven migration of gas flow particles to the hub and tip regions, (2) providing a robust film cooling technology at the hub and that sustains high cooling effectiveness lessmore » sensitive to deposition, (3) investigating the impact of blade tip geometry on film cooling effectiveness. The document includes numerical and experimental investigations of above issues. The experimental investigations were performed in the three-stage multi-purpose turbine research facility at the Turbomachinery Performance and Flow Research Laboratory (TPFL), Texas A&M University. For the numerical investigations a commercial Navier-Stokes solver was utilized.« less

Wind-tunnel test of an articulated helicopter rotor model with several tip shapes

NASA Technical Reports Server (NTRS)

Berry, J. D.; Mineck, R. E.

1980-01-01

Six interchangeable tip shapes were tested: a square (baseline) tip, an ogee tip, a subwing tip, a swept tip, a winglet tip, and a short ogee tip. In hover at the lower rotational speeds the swept, ogee, and short ogee tips had about the same torque coefficient, and the subwing and winglet tips had a larger torque coefficient than the baseline square tip blades. The ogee and swept tip blades required less torque coefficient at lower rotational speeds and roughly equivalent torque coefficient at higher rotational speeds compared with the baseline square tip blades in forward flight. The short ogee tip required higher torque coefficient at higher lift coefficients than the baseline square tip blade in the forward flight test condition.

RANS Simulation (Virtual Blade Model [VBM]) of Single Lab Scaled DOE RM1 MHK Turbine

DOE Data Explorer

Javaherchi, Teymour; Stelzenmuller, Nick; Aliseda, Alberto; Seydel, Joseph

2014-04-15

Attached are the .cas and .dat files for the Reynolds Averaged Navier-Stokes (RANS) simulation of a single lab-scaled DOE RM1 turbine implemented in ANSYS FLUENT CFD-package. The lab-scaled DOE RM1 is a re-design geometry, based of the full scale DOE RM1 design, producing same power output as the full scale model, while operating at matched Tip Speed Ratio values at reachable laboratory Reynolds number (see attached paper). In this case study the flow field around and in the wake of the lab-scaled DOE RM1 turbine is simulated using Blade Element Model (a.k.a Virtual Blade Model) by solving RANS equations coupled with k-\\omega turbulence closure model. It should be highlighted that in this simulation the actual geometry of the rotor blade is not modeled. The effect of turbine rotating blades are modeled using the Blade Element Theory. This simulation provides an accurate estimate for the performance of device and structure of it's turbulent far wake. Due to the simplifications implemented for modeling the rotating blades in this model, VBM is limited to capture details of the flow field in near wake region of the device. The required User Defined Functions (UDFs) and look-up table of lift and drag coefficients are included along with the .cas and .dat files.

Extended Glauert tip correction to include vortex rollup effects

DOE Office of Scientific and Technical Information (OSTI.GOV)

Maniaci, David; Schmitz, Sven

Wind turbine loads predictions by blade-element momentum theory using the standard tip-loss correction have been shown to over-predict loading near the blade tip in comparison to experimental data. This over-prediction is theorized to be due to the assumption of light rotor loading, inherent in the standard tip-loss correction model of Glauert. A higher- order free-wake method, WindDVE, is used to compute the rollup process of the trailing vortex sheets downstream of wind turbine blades. Results obtained serve an exact correction function to the Glauert tip correction used in blade-element momentum methods. Lastly, it is found that accounting for the effectsmore » of tip vortex rollup within the Glauert tip correction indeed results in improved prediction of blade tip loads computed by blade-element momentum methods.« less

Extended Glauert tip correction to include vortex rollup effects

DOE PAGES

Maniaci, David; Schmitz, Sven

2016-10-03

Wind turbine loads predictions by blade-element momentum theory using the standard tip-loss correction have been shown to over-predict loading near the blade tip in comparison to experimental data. This over-prediction is theorized to be due to the assumption of light rotor loading, inherent in the standard tip-loss correction model of Glauert. A higher- order free-wake method, WindDVE, is used to compute the rollup process of the trailing vortex sheets downstream of wind turbine blades. Results obtained serve an exact correction function to the Glauert tip correction used in blade-element momentum methods. Lastly, it is found that accounting for the effectsmore » of tip vortex rollup within the Glauert tip correction indeed results in improved prediction of blade tip loads computed by blade-element momentum methods.« less

The design and development of transonic multistage compressors

NASA Technical Reports Server (NTRS)

Ball, C. L.; Steinke, R. J.; Newman, F. A.

1988-01-01

The development of the transonic multistage compressor is reviewed. Changing trends in design and performance parameters are noted. These changes are related to advances in compressor aerodynamics, computational fluid mechanics and other enabling technologies. The parameters normally given to the designer and those that need to be established during the design process are identified. Criteria and procedures used in the selection of these parameters are presented. The selection of tip speed, aerodynamic loading, flowpath geometry, incidence and deviation angles, blade/vane geometry, blade/vane solidity, stage reaction, aerodynamic blockage, inlet flow per unit annulus area, stage/overall velocity ratio, and aerodynamic losses are considered. Trends in these parameters both spanwise and axially through the machine are highlighted. The effects of flow mixing and methods for accounting for the mixing in the design process are discussed.

Numerical analysis of the blade tip-timing signal of a fiber bundle sensor probe

NASA Astrophysics Data System (ADS)

Guo, Haotian; Duan, Fajie; Cheng, Zhonghai

2015-03-01

Blade tip-timing is the most effective method for online blade vibration measurement of large rotating machines like turbine engines. Fiber bundle sensors are utilized in tip-timing system to measure the arrival time of the blade. The model of the tip-timing signal of the fiber bundle sensor is established. Experiments are conducted and the results are in concordance with the model established. The rising speed of the tip-timing signal is analyzed. To minimize the tip-timing error, the effects of the clearance change between the sensor and the blade and the deflection of the tip surface are analyzed. Simulation results indicate that the variable gain amplifier, which amplifies the signals to a similar level, can eliminate the measurement error caused by the variation of the clearance between the sensor and blade. Increasing the clearance between the sensor and blade can reduce the measurement error introduced by deflection of the tip surface.

Materials for advanced turbine engines. Volume 1: Advanced blade tip seal system

NASA Technical Reports Server (NTRS)

Zelahy, J. W.; Fairbanks, N. P.

1982-01-01

Project 3, the subject of this technical report, was structured toward the successful engine demonstration of an improved-efficiency, long-life, tip-seal system for turbine blades. The advanced tip-seal system was designed to maintain close operating clearances between turbine blade tips and turbine shrouds and, at the same time, be resistant to environmental effects including high-temperature oxidation, hot corrosion, and thermal cycling. The turbine blade tip comprised an environmentally resistant, activated-diffussion-bonded, monocrystal superalloy combined with a thin layer of aluminium oxide abrasive particles entrapped in an electroplated NiCr matrix. The project established the tip design and joint location, characterized the single-crystal tip alloy and abrasive tip treatment, and established the manufacturing and quality-control plans required to fully process the blades. A total of 171 blades were fully manufactured, and 100 were endurance and performance engine-tested.

Air-Cooled Turbine Blades with Tip Cap For Improved Leading-Edge Cooling

NASA Technical Reports Server (NTRS)

Calvert, Howard F.; Meyer, Andre J., Jr.; Morgan, William C.

1959-01-01

An investigation was conducted in a modified turbojet engine to determine the cooling characteristics of the semistrut corrugated air- cooled turbine blade and to compare and evaluate a leading-edge tip cap as a means for improving the leading-edge cooling characteristics of cooled turbine blades. Temperature data were obtained from uncapped air-cooled blades (blade A), cooled blades with the leading-edge tip area capped (blade B), and blades with slanted corrugations in addition to leading-edge tip caps (blade C). All data are for rated engine speed and turbine-inlet temperature (1660 F). A comparison of temperature data from blades A and B showed a leading-edge temperature reduction of about 130 F that could be attributed to the use of tip caps. Even better leading-edge cooling was obtained with blade C. Blade C also operated with the smallest chordwise temperature gradients of the blades tested, but tip-capped blade B operated with the lowest average chordwise temperature. According to a correlation of the experimental data, all three blade types 0 could operate satisfactorily with a turbine-inlet temperature of 2000 F and a coolant flow of 3 percent of engine mass flow or less, with an average chordwise temperature limit of 1400 F. Within the range of coolant flows investigated, however, only blade C could maintain a leading-edge temperature of 1400 F for a turbine-inlet temperature of 2000 F.

RANS Simulation (Rotating Reference Frame Model [RRF]) of Single Lab-Scaled DOE RM1 MHK Turbine

DOE Data Explorer

Javaherchi, Teymour; Stelzenmuller, Nick; Aliseda, Alberto; Seydel, Joseph

2014-04-15

Attached are the .cas and .dat files for the Reynolds Averaged Navier-Stokes (RANS) simulation of a single lab-scaled DOE RM1 turbine implemented in ANSYS FLUENT CFD-package. The lab-scaled DOE RM1 is a re-design geometry, based of the full scale DOE RM1 design, producing same power output as the full scale model, while operating at matched Tip Speed Ratio values at reachable laboratory Reynolds number (see attached paper). In this case study taking advantage of the symmetry of lab-scaled DOE RM1 geometry, only half of the geometry is models using (Single) Rotating Reference Frame model [RRF]. In this model RANS equations, coupled with k-\\omega turbulence closure model, are solved in the rotating reference frame. The actual geometry of the turbine blade is included and the turbulent boundary layer along the blade span is simulated using wall-function approach. The rotation of the blade is modeled by applying periodic boundary condition to sets of plane of symmetry. This case study simulates the performance and flow field in the near and far wake of the device at the desired operating conditions. The results of these simulations were validated against in-house experimental data. Please see the attached paper.

Advanced ceramic material for high temperature turbine tip seals

NASA Technical Reports Server (NTRS)

Vogan, J. W.; Solomon, N. G.; Stetson, A. R.

1980-01-01

Forty-one material systems were evaluated for potential use in turbine blade tip seal applications at 1370 C. Both ceramic blade tip inserts and abradable ceramic tip shoes were tested. Hot gas erosion, impact resistance, thermal stability, and dynamic rub performance were the criteria used in rating the various materials. Silicon carbide and silicon nitride were used, both as blade tips and abradables. The blade tip inserts were fabricated by hot pressing while low density and honeycomb abradables were sintered or reaction bonded.

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2010 CFR

2010-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2011 CFR

2011-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2012 CFR

2012-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2013 CFR

2013-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

14 CFR Appendix J to Part 36 - Alternative Noise Certification Procedure for Helicopters Under Subpart H Having a Maximum...

Code of Federal Regulations, 2014 CFR

2014-01-01

... produce the same advancing blade tip Mach number as associated with the reference conditions; (i) Advancing blade tip Mach number (MAT) is defined as the ratio of the arithmetic sum of blade tip rotational... the reference advancing blade tip Mach number. The adjusted reference airspeed shall be maintained...

A review of turbine blade tip heat transfer.

PubMed

Bunker, R S

2001-05-01

This paper presents a review of the publicly available knowledge base concerning turbine blade tip heat transfer, from the early fundamental research which laid the foundations of our knowledge, to current experimental and numerical studies utilizing engine-scaled blade cascades and turbine rigs. Focus is placed on high-pressure, high-temperature axial-turbine blade tips, which are prevalent in the majority of today's aircraft engines and power generating turbines. The state of our current understanding of turbine blade tip heat transfer is in the transitional phase between fundamentals supported by engine-based experience, and the ability to a priori correctly predict and efficiently design blade tips for engine service.

On the development of a magnetoresistive sensor for blade tip timing and blade tip clearance measurement systems

NASA Astrophysics Data System (ADS)

Tomassini, R.; Rossi, G.; Brouckaert, J.-F.

2014-05-01

The accurate control of the gap between static and rotating components is vital to preserve the mechanical integrity and ensure a correct functioning of any rotating machinery. Moreover, tip leakage above the airfoil tip results in relevant aerodynamic losses. One way to measure and to monitor blade tip gaps is by the so-called Blade Tip Clearance (BTC) technique. Another fundamental phenomenon to control in the turbomachines is the vibration of the blades. For more than half a century, this has been performed by installing strain gauges on the blades and using telemetry to transmit the signals. The Blade Tip Timing (BTT) technique, (i.e. measuring the blade time of arrival from the casing at different angular locations with proximity sensors) is currently being adopted by all manufacturers as a replacement for the classical strain gauge technique because of its non-intrusive character. This paper presents a novel magnetoresistive sensor for blade tip timing and blade tip clearance systems, which offers high temporal and high spatial resolution simultaneously. The sensing element adopted is a Wheatstone bridge of Permalloy elements. The principle of the sensor is based on the variation of magnetic field at the passage of ferromagnetic objects. Two different configurations have been realized, a digital and an analogue sensor. Measurements of tip clearance have been performed in an high speed compressor and the calibration curve is reported. Measurements of blade vibration have been carried out in a dedicated calibration bench; results are presented and discussed. The magnetoresistive sensor is characterized by high repeatability, low manufacturing costs and measurement accuracy in line with the main probes used in turbomachinery testing. The novel sensor has great potential and is capable of fulfilling the requirements for a simultaneous BTC and BTT measurement system.

Evaluation of urethane and carbide-tipped blades on wheel-supported snow plows.

DOT National Transportation Integrated Search

1997-01-01

The objective of this study was to evaluate the performance of urethane and carbide-tipped snow plow blades on wheel supported plows. Their performance was compared to that of VDOT's standard blade arrangement: carbide-tipped blades on plows without ...

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

NASA Technical Reports Server (NTRS)

Azad, Gm S.; Han, Je-Chin; Boyle, Robert J.

2000-01-01

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.

Analysis of internal flow of J85-13 multistage compressor

NASA Technical Reports Server (NTRS)

Hager, R. D.

1977-01-01

Interstage data recorded on a J85-13 engine were used to analyze the internal flow of the compressor. Measured pressures and temperatures were used as input to a streamline analysis program to calculate the velocity diagrams at the inlet and outlet of each blade row. From the velocity diagrams and blade geometry, selected blade-element performance parameters were calculated. From the detailed analysis it is concluded that the compressor is probably hub critical (stall initiates at the hub) in the latter stages for the design speed conditions. As a result, the casing treatment over the blade tips has little or no effect on stall margin at design speed. Radial inlet distortion did not appear to change the flow in the stages that control stall because of the rapid attenuation of the distortion within the compressor.

Correlation of Puma airloads: Lifting-line and wake calculation

NASA Technical Reports Server (NTRS)

Bousman, William G.; Young, Colin; Gilbert, Neil; Toulmay, Francois; Johnson, Wayne; Riley, M. J.

1989-01-01

A cooperative program undertaken by organizations in the United States, England, France, and Australia has assessed the strengths and weaknesses of four lifting-line/wake methods and three CFD methods by comparing their predictions with the data obtained in flight trials of a research Puma. The Puma was tested in two configurations: a mixed bladed rotor with instrumented rectangular tip blades, and a configuration with four identical swept tip blades. The results are examined of the lifting-line predictions. The better lifting-line methods show good agreement with lift at the blade tip for the configuration with four swept tips; the moment is well predicted at 0.92 R, but deteriorates outboard. The predictions for the mixed bladed rotor configuration range from fair to good. The lift prediction is better for the swept tip blade than for the rectangular tip blade, but the reasons for this cannot be determined because of the unmodeled effects of the mixed bladed rotor.

Effect of rotor tip clearance and configuration on overall performance of a 12.77-centimeter tip diameter axial-flow turbine

NASA Technical Reports Server (NTRS)

Haas, J. E.; Kofskey, M. G.

1978-01-01

The rotor tip clearance was obtained by use of a recess in the casing above the rotor blades and also by use of a reduced blade height. For the recessed casing configuration, the optimum rotor blade height was found to be the one where the rotor tip diameter was equal to the stator tip diameter. The tip clearance loss associated with this optimum recessed casing configuration was less than that for the reduced blade height configuration.

Effects of Double-Leakage Tip Clearance Flow on the Performance of a Compressor Stage with a Large Rotor Tip Gap

NASA Technical Reports Server (NTRS)

Hah, Chunill

2016-01-01

Effects of a large rotor tip gap on the performance of a one and half stage axial compressor are investigated in detail with a numerical simulation based on LES and available PIV data. The current paper studies the main flow physics, including why and how the loss generation is increased with the large rotor tip gap. The present study reveals that when the tip gap becomes large, tip clearance fluid goes over the tip clearance core vortex and enters into the next blade's tip gap, which is called double-leakage tip clearance flow. As the tip clearance flow enters into the adjacent blade's tip gap, a vortex rope with a lower pressure core is generated. This vortex rope breaks up the tip clearance core vortex of the adjacent blade, resulting in a large additional mixing. This double-leakage tip clearance flow occurs at all operating conditions, from design flow to near stall condition, with the large tip gap for the current compressor stage. The double-leakage tip clearance flow, its interaction with the tip clearance core vortex of the adjacent blade, and the resulting large mixing loss are the main flow mechanism of the large rotor tip gap in the compressor. When the tip clearance is smaller, flow near the end wall follows more closely with the main passage flow and this double-leakage tip clearance flow does not happen near the design flow condition for the current compressor stage. When the compressor with a large tip gap operates at near stall operation, a strong vortex rope is generated near the leading edge due to the double-leakage flow. Part of this vortex separates from the path of the tip clearance core vortex and travels from the suction side of the blade toward the pressure side of the blade. This vortex is generated periodically at near stall operation with a large tip gap. As the vortex travels from the suction side to the pressure side of the blade, a large fluctuation of local pressure forces blade vibration. Nonsynchronous blade vibration occurs due to this vortex as the frequency of this vortex generation is not the same as the rotor. The present investigation confirms that this vortex is a part of separated tip clearance vortex, which is caused by the double-leakage tip clearance flow.

RANS Simulation (Virtual Blade Model [VBM]) of Array of Three Coaxial Lab Scaled DOE RM1 MHK Turbine with 5D Spacing

DOE Data Explorer

Javaherchi, Teymour

2016-06-08

Attached are the .cas and .dat files along with the required User Defined Functions (UDFs) and look-up table of lift and drag coefficients for the Reynolds Averaged Navier-Stokes (RANS) simulation of three coaxially located lab-scaled DOE RM1 turbine implemented in ANSYS FLUENT CFD-package. The lab-scaled DOE RM1 is a re-design geometry, based of the full scale DOE RM1 design, producing same power output as the full scale model, while operating at matched Tip Speed Ratio values at reachable laboratory Reynolds number (see attached paper). In this case study the flow field around and in the wake of the lab-scaled DOE RM1 turbines in a coaxial array is simulated using Blade Element Model (a.k.a Virtual Blade Model) by solving RANS equations coupled with k-\\omega turbulence closure model. It should be highlighted that in this simulation the actual geometry of the rotor blade is not modeled. The effect of turbine rotating blades are modeled using the Blade Element Theory. This simulation provides an accurate estimate for the performance of each device and structure of their turbulent far wake. The results of these simulations were validated against the developed in-house experimental data. Simulations for other turbine configurations are available upon request.

Acoustic and aerodynamic performance of a 1.5-pressure-ratio, 1.83-meter (6 ft) diameter fan stage for turbofan engines (QF-2)

NASA Technical Reports Server (NTRS)

Woodward, R. P.; Lucas, J. G.; Balombin, J. R.

1977-01-01

The fan was externally driven by an electric motor. Design features for low-noise generation included the elimination of inlet guide vanes, long axial spacing between the rotor and stator blade rows, and the selection of blade-vane numbers to achieve duct-mode cutoff. The fan QF-2 results were compared with those of another full-scale fan having essentially identical aerodynamic design except for nozzle geometry and the direction of rotation. The fan QF-2 aerodynamic results were also compared with those obtained from a 50.8 cm rotor-tip-diameter model of the reverse rotation fan QF-2 design. Differences in nozzle geometry other than exit area significantly affected the comparison of the results of the full-scale fans.

Local measurement and numerical modeling of mass/heat transfer from a turbine blade in a linear cascade with tip clearance

NASA Astrophysics Data System (ADS)

Jin, Peitong

2000-11-01

Local mass/heat transfer measurements from the turbine blade near-tip and the tip surfaces are performed using the naphthalene sublimation technique. The experiments are conducted in a linear cascade consisting of five high-pressure blades with a central test-blade configuration. The incoming flow conditions are close to those of the gas turbine engine environment (boundary layer displacement thickness is about 0.01 of chord) with an exit Reynolds number of 6.2 x 105. The effects of tip clearance level (0.86%--6.90% of chord), mainstream Reynolds number and turbulence intensity (0.2 and 12.0%) are investigated. Two methods of flow visualization---oil and lampblack, laser light sheet smoke wire---as well as static pressure measurement on the blade surface are used to study the tip leakage flow and vortex in the cascade. In addition, numerical modeling of the flow and heat transfer processes in the linear cascade with different tip clearances is conducted using commercial software incorporating advanced turbulence models. The present study confirms many important results on the tip leakage flow and vortex from the literature, contributes to the current understanding in the effects of tip leakage flow and vortex on local heat transfer from the blade near-tip and the tip surfaces, and provides detailed local and average heat/mass transfer data applicable to turbine blade tip cooling design.

Transient airload computer analysis for simulating wind induced impulsive noise conditions of a hovering helicopter rotor

NASA Technical Reports Server (NTRS)

Hall, G. F.

1975-01-01

A numerical analysis was developed to determine the airloads on helicopter rotors operating under near-hovering flight conditions capable of producing impulsive noise. A computer program was written in which the solutions for the rotor tip vortex geometry, inflow, aeroelastic response, and airloads are solved in a coupled manner at sequential time steps, with or without the influence of an imposed steady ambient wind or transient gust. The program was developed for future applications in which predicted airloads would be incorporated in an acoustics analysis to attempt to predict and analyze impulsive noise (blade slap). The analysis was applied to a hovering full-scale rotor for which impulsive noise was recorded in the presence of ambient wind. The predicted tip vortex coordinates are in reasonable agreement with the test data, and the blade airload solutions converged to a periodic behavior for an imposed steady ambient wind conditions.

Acoustic measurements of a full-scale rotor with four tip shapes. Volume 2: Appendices C, D, E and F

NASA Technical Reports Server (NTRS)

Mosher, M.

1984-01-01

A full scale helicopter with four different blade tip geometries is tested in a 40- by 80-foot wind tunnel. Performance, loads, and noise are measured. The four tip shapes tested were rectangular, tapered, swept, and swept/tapered. The noise data include measurements of the sound pressure levels in decibels dB, decibels adjusted dBA, and tone-corrected PNdB, for all of the conditions tested. Also included are the detailed measurements, 1/3 octave spectra and time histories for some selected data, and plots of dBA as function of test condition. Some performance measurements are included to aid interpretation of the noise data.

Advanced optical blade tip clearance measurement system

NASA Technical Reports Server (NTRS)

Ford, M. J.; Honeycutt, R. E.; Nordlund, R. E.; Robinson, W. W.

1978-01-01

An advanced electro-optical system was developed to measure single blade tip clearances and average blade tip clearances between a rotor and its gas path seal in an operating gas turbine engine. This system is applicable to fan, compressor, and turbine blade tip clearance measurement requirements, and the system probe is particularly suitable for operation in the extreme turbine environment. A study of optical properties of blade tips was conducted to establish measurement system application limitations. A series of laboratory tests was conducted to determine the measurement system's operational performance characteristics and to demonstrate system capability under simulated operating gas turbine environmental conditions. Operational and environmental performance test data are presented.

UWB Wind Turbine Blade Deflection Sensing for Wind Energy Cost Reduction.

PubMed

Zhang, Shuai; Jensen, Tobias Lindstrøm; Franek, Ondrej; Eggers, Patrick C F; Olesen, Kim; Byskov, Claus; Pedersen, Gert Frølund

2015-08-12

A new application of utilizing ultra-wideband (UWB) technology to sense wind turbine blade deflections is introduced in this paper for wind energy cost reduction. The lower UWB band of 3.1-5.3 GHz is applied. On each blade, there will be one UWB blade deflection sensing system, which consists of two UWB antennas at the blade root and one UWB antenna at the blade tip. The detailed topology and challenges of this deflection sensing system are addressed. Due to the complexity of the problem, this paper will first realize the on-blade UWB radio link in the simplest case, where the tip antenna is situated outside (and on the surface of) a blade tip. To investigate this case, full-blade time-domain measurements are designed and conducted under different deflections. The detailed measurement setups and results are provided. If the root and tip antenna locations are properly selected, the first pulse is always of sufficient quality for accurate estimations under different deflections. The measured results reveal that the blade tip-root distance and blade deflection can be accurately estimated in the complicated and lossy wireless channels around a wind turbine blade. Some future research topics on this application are listed finally.

Evaluation of helicopter noise due to b blade-vortex interaction for five tip configurations. [conducted in the Langley V/STOL tunnel

NASA Technical Reports Server (NTRS)

Hoad, D. R.

1979-01-01

The effect of tip shape modification on blade vortex interaction induced helicopter blade slap noise was investigated. Simulated flight and descent velocities which have been shown to produce blade slap were tested. Aerodynamic performance parameters of the rotor system were monitored to ensure properly matched flight conditions among the tip shapes. The tunnel was operated in the open throat configuration with treatment to improve the acoustic characteristics of the test chamber. Four promising tips were used along with a standard square tip as a baseline configuration. A detailed acoustic evaluation on the same rotor system of the relative applicability of the various tip configurations for blade slap noise reduction is provided.

A new aeroelastic model for composite rotor blades with straight and swept tips

NASA Technical Reports Server (NTRS)

Yuan, Kuo-An; Friedmann, Peretz P.; Venkatesan, Comandur

1992-01-01

An analytical model for predicting the aeroelastic behavior of composite rotor blades with straight and swept tips is presented. The blade is modeled by beam type finite elements along the elastic axis. A single finite element is used to model the swept tip. The nonlinear equations of motion for the finite element model are derived using Hamilton's principle and based on a moderate deflection theory and accounts for: arbitrary cross-sectional shape, pretwist, generally anisotropic material behavior, transverse shears and out-of-plane warping. Numerical results illustrating the effects of tip sweep, anhedral and composite ply orientation on blade aeroelastic behavior are presented. Tip sweep can induce aeroelastic instability by flap-twist coupling. Tip anhedral causes lag-torsion and flap-axial couplings, however, its effects on blade stability is less pronounced than the effect due to sweep. Composite ply orientation has a substantial effect on blade stability.

Rotor blade system with reduced blade-vortex interaction noise

NASA Technical Reports Server (NTRS)

Leishman, John G. (Inventor); Han, Yong Oun (Inventor)

2005-01-01

A rotor blade system with reduced blade-vortex interaction noise includes a plurality of tube members embedded in proximity to a tip of each rotor blade. The inlets of the tube members are arrayed at the leading edge of the blade slightly above the chord plane, while the outlets are arrayed at the blade tip face. Such a design rapidly diffuses the vorticity contained within the concentrated tip vortex because of enhanced flow mixing in the inner core, which prevents the development of a laminar core region.

Evaluation of Aeroelastically Tailored Small Wind Turbine Blades Final Project Report

DOE Office of Scientific and Technical Information (OSTI.GOV)

Griffin, Dayton A.

2005-09-29

Evaluation of Aeroelastically Tailored Small Wind Turbine Blades Final Report Global Energy Concepts, LLC (GEC) has performed a conceptual design study concerning aeroelastic tailoring of small wind turbine blades. The primary objectives were to evaluate ways that blade/rotor geometry could be used to enable cost-of-energy reductions by enhancing energy capture while constraining or mitigating blade costs, system loads, and related component costs. This work builds on insights developed in ongoing adaptive-blade programs but with a focus on application to small turbine systems with isotropic blade material properties and with combined blade sweep and pre-bending/pre-curving to achieve the desired twist coupling.more » Specific goals of this project are to: (A) Evaluate and quantify the extent to which rotor geometry can be used to realize load-mitigating small wind turbine rotors. Primary aspects of the load mitigation are: (1) Improved overspeed safety affected by blades twisting toward stall in response to speed increases. (2) Reduced fatigue loading affected by blade twisting toward feather in response to turbulent gusts. (B) Illustrate trade-offs and design sensitivities for this concept. (C) Provide the technical basis for small wind turbine manufacturers to evaluate this concept and commercialize if the technology appears favorable. The SolidWorks code was used to rapidly develop solid models of blade with varying shapes and material properties. Finite element analyses (FEA) were performed using the COSMOS code modeling with tip-loads and centripetal accelerations. This tool set was used to investigate the potential for aeroelastic tailoring with combined planform sweep and pre-curve. An extensive matrix of design variables was investigated, including aerodynamic design, magnitude and shape of planform sweep, magnitude and shape of blade pre-curve, material stiffness, and rotor diameter. The FEA simulations resulted in substantial insights into the structural response of these blades. The trends were used to identify geometries and rotor configurations that showed the greatest promise for achieving beneficial aeroelastic response. The ADAMS code was used to perform complete aeroelastic simulations of selected rotor configurations; however, the results of these simulations were not satisfactory. This report documents the challenges encountered with the ADAMS simulations and presents recommendations for further development of this concept for aeroelastically tailored small wind turbine blades.« less

Aeroelastic considerations for torsionally soft rotors

NASA Technical Reports Server (NTRS)

Mantay, W. R.; Yeager, W. T., Jr.

1985-01-01

A research study was initiated to systematically determine the impact of selected blade tip geometric parameters on conformable rotor performance and loads characteristics. The model articulated rotors included baseline and torsionally soft blades with interchangeable tips. Seven blade tip designs were evaluated on the baseline rotor and six tip designs were tested on the torsionally soft blades. The designs incorporated a systemmatic variation in geometric parameters including sweep, taper, and anhedral. The rotors were evaluated in the NASA Langley Transonic Dynamics Tunnel at several advance ratios, lift and propulsive force values, and tip Mach numbers. A track sensitivity study was also conducted at several advance ratios for both rotors. Based on the test results, tip parameter variations generated significant rotor performance and loads differences for both baseline and torsionally soft blades.

Comparison of NACA 6-series and 4-digit airfoils for Darrieus wind turbines

NASA Astrophysics Data System (ADS)

Migliore, P. G.

1983-08-01

The aerodynamic efficiency of Darrieus wind turbines as effected by blade airfoil geometry was investigated. Analysis was limited to curved-bladed machines having rotor solidities of 7-21 percent and operating at a Reynolds number of 3 x 10 to the 6th. Ten different airfoils, having thickness-to-chord ratios of 12, 15, and 18 percent, were studied. Performance estimates were made using a blade element/momentum theory approach. Results indicated that NACA 6-series airfoils yield peak power coefficients as great as NACA 4-digit airfoils and have broader and flatter power coefficient-tip speed ratio curves. Sample calculations for an NACA 63(2)-015 airfoil showed an annual energy output increase of 17-27 percent, depending on rotor solidity, compared to an NACA 0015 airfoil.

UWB Wind Turbine Blade Deflection Sensing for Wind Energy Cost Reduction

PubMed Central

Zhang, Shuai; Jensen, Tobias Lindstrøm; Franek, Ondrej; Eggers, Patrick C. F.; Olesen, Kim; Byskov, Claus; Pedersen, Gert Frølund

2015-01-01

A new application of utilizing ultra-wideband (UWB) technology to sense wind turbine blade deflections is introduced in this paper for wind energy cost reduction. The lower UWB band of 3.1–5.3 GHz is applied. On each blade, there will be one UWB blade deflection sensing system, which consists of two UWB antennas at the blade root and one UWB antenna at the blade tip. The detailed topology and challenges of this deflection sensing system are addressed. Due to the complexity of the problem, this paper will first realize the on-blade UWB radio link in the simplest case, where the tip antenna is situated outside (and on the surface of) a blade tip. To investigate this case, full-blade time-domain measurements are designed and conducted under different deflections. The detailed measurement setups and results are provided. If the root and tip antenna locations are properly selected, the first pulse is always of sufficient quality for accurate estimations under different deflections. The measured results reveal that the blade tip-root distance and blade deflection can be accurately estimated in the complicated and lossy wireless channels around a wind turbine blade. Some future research topics on this application are listed finally. PMID:26274964

A study of the rotor wake of a small-scale rotor model in forward flight using laser light sheet flow visualization with comparisons to analytical models

NASA Technical Reports Server (NTRS)

Ghee, Terence A.; Elliott, Joe W.

1992-01-01

An experimental investigation was conducted in the 14 by 22 ft subsonic tunnel at NASA Langley Research Center to quantify the rotor wake behind a scale model helicopter rotor in forward flight (mu = 0.15 and 0.23) at one thrust level (C sub T = 0.0064). The rotor system used in the present test consisted of a four-bladed, fully articulated hub and utilized blades of rectangular planform with a NACA-0012 airfoil section. A laser light sheet, seeded with propylene glycol smoke, was used to visualize the flow in planes parallel and perpendicular to the freestream flow. Quantitative measurements of vortex location, vertical skew angle, and vortex particle void radius were obtained for vortices in the flow; convective velocities were obtained for blade tip vortices. Comparisons were made between the experimental results and the wake geometry generated by computational predictions. The results of these comparisons show that the interaction between wake vortex structures is an important consideration for correctly predicting the wake geometry.

Aeroelastic modeling of composite rotor blades with straight and swept tips

NASA Technical Reports Server (NTRS)

Yuan, Kuo-An; Friedmann, Peretz P.; Venkatesan, Comandur

1992-01-01

This paper presents an analytical study of the aeroelastic behavior of composite rotor blades with straight and swept tips. The blade is modeled by beam type finite elements. A single finite element is used to model the swept tip. The nonlinear equations of motion for the FEM are derived using Hamilton's principle and based on a moderate deflection theory and accounts for: arbitrary cross-sectional shape, pretwist, generally anisotropic material behavior, transverse shears and out-of-plane warping. Numerical results illustrating the effects of tip sweep, anhedral and composite ply orientation on blade aeroelastic behavior are presented. It is shown that composite ply orientation has a substantial effect on blade stability. At low thrust conditions, certain ply orientations can cause instability in the lag mode. The flap-torsion coupling associated with tip sweep can also induce aeroelastic instability in the blade. This instability can be removed by appropriate ply orientation in the composite construction. These results illustrate the inherent potential for aeroelastic tailoring present in composite rotor blades with swept tips, which still remains to be exploited in the design process.

Aeroelastic considerations for torsionally soft rotors

NASA Technical Reports Server (NTRS)

Mantay, W. R.; Yeager, W. T., Jr.

1986-01-01

A research study was initiated to systematically determine the impact of selected blade tip geometric parameters on conformable rotor performance and loads characteristics. The model articulated rotors included baseline and torsionally soft blades with interchangeable tips. Seven blade tip designs were evaluated on the baseline rotor and six tip designs were tested on the torsionally soft blades. The designs incorporated a systemmatic variation in geometric parameters including sweep, taper, and anhedral. The rotors were evaluated in the NASA Langley Transonic Dynamics Tunnel at several advance ratios, lift and propulsive force values, and tip Mach numbers. A track sensitivity study was also conducted at several advance ratios for both rotors. Based on the test results, tip parameter variations generated significant rotor performance and loads differences for both baseline and torsionally soft blades. Azimuthal variation of elastic twist generated by variations in the tip parameters strongly correlated with rotor performance and loads, but the magnitude of advancing blade elastic twist did not. In addition, fixed system vibratory loads and rotor track for potential conformable rotor candidates appears very sensitive to parametric rotor changes.

Axial compressor blade design for desensitization of aerodynamic performance and stability to tip clearance

NASA Astrophysics Data System (ADS)

Erler, Engin

Tip clearance flow is the flow through the clearance between the rotor blade tip and the shroud of a turbomachine, such as compressors and turbines. This flow is driven by the pressure difference across the blade (aerodynamic loading) in the tip region and is a major source of loss in performance and aerodynamic stability in axial compressors of modern aircraft engines. An increase in tip clearance, either temporary due to differential radial expansion between the blade and the shroud during transient operation or permanent due to engine wear or manufacturing tolerances on small blades, increases tip clearance flow and results in higher fuel consumption and higher risk of engine surge. A compressor design that can reduce the sensitivity of its performance and aerodynamic stability to tip clearance increase would have a major impact on short and long-term engine performance and operating envelope. While much research has been carried out on improving nominal compressor performance, little had been done on desensitization to tip clearance increase beyond isolated observations that certain blade designs such as forward chordwise sweep, seem to be less sensitive to tip clearance size increase. The current project aims to identify through a computational study the flow features and associated mechanisms that reduces sensitivity of axial compressor rotors to tip clearance size and propose blade design strategies that can exploit these results. The methodology starts with the design of a reference conventional axial compressor rotor followed by a parametric study with variations of this reference design through modification of the camber line and of the stacking line of blade profiles along the span. It is noted that a simple desensitization method would be to reduce the aerodynamic loading of the blade tip which would reduce the tip clearance flow and its proportional contribution to performance loss. However, with the larger part of the work on the flow done in this region, this approach would entail a nominal performance penalty. Therefore, the chosen rotor design philosophy aims to keep the spanwise loading constant to avoid trading performance for desensitization. The rotor designs that resulted from this exercise are simulated in ANSYS CFX at different tip clearance sizes. The change in their performance with respect to tip clearance size (sensitivity) is compared both on an integral level in terms of pressure ratio and adiabatic efficiency, as well as on a detailed level in terms of aerodynamic losses and blockage associated with tip clearance flow. The sensitivity of aerodynamic stability is evaluated either directly through the simulations of the rotor characteristics up to the stall point (expensive in time and resources) for a few designs or indirectly through the position of the interface between the incoming and tip clearance flow with respect to the rotor leading edge plane. The latter approach is based on a generally observed stall criteria in modern axial compressors. The rotor designs are then assessed according to their sensitivity in comparison to that of the reference rotor design to detect features that can explain the trend in sensitivity to tip clearance size. These features can then be validated and the associated flow mechanisms explained through numerical simulations and modelling. Analysis of the database from the rotor parametric study shows that the observed trend in sensitivity cannot be explained by the shifting of the aerodynamic loading along the blade chord, as initially hypothesized based on the literature review. Instead, two flow features are found to reduce sensitivity of performance and stability to tip clearance, namely an increase in incoming meridional momentum in the tip region and a reduction/elimination of double leakage flow. Double leakage flow is the flow that exits the tip clearance of one blade and proceeds into the clearance of the adjacent blade rather than convecting downstream out of the local blade passage. These flow features are isolated and validated based on the reference rotor design through changes in the inlet total pressure condition to alter incoming flow momentum and blade number count to change double leakage rate. In terms of flow mechanism, double leakage is shown to be detrimental to performance and stability, and its proportional increase with tip clearance size explains the sensitivity increase in the presence of double leakage and, conversely, the desensitization effect of reducing or eliminating double leakage. The increase in incoming meridional momentum in the tip region reduces sensitivity to tip clearance through its reduction of double leakage as well as through improved mixing with tip clearance flow, as demonstrated by an analytical model without double leakage flow. The above results imply that any blade design strategy that exploits the two desensitizing flow features would reduce the performance and stability sensitivity to tip clearance size. The increase of the incoming meridional momentum can be achieved through forward chordwise sweep of the blade. The reduction of double leakage without changing blade pitch can be obtained by decreasing the blade stagger angle in the tip region. Examples of blade designs associated with these strategies are shown through CFX simulations to be successful in reducing sensitivity to tip clearance size.

75 FR 26889 - Airworthiness Directives; Arrow Falcon Exporters, Inc. (previously Utah State University) et al...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-05-13

... the forward tip weight retention block (tip block) or aft tip closure (tip closure), loss of the blade...) forward tip weight retention block (tip block) and the aft tip closure (tip closure) for adhesive bond... prevent loss of a tip block or tip closure, loss of a blade, and subsequent loss of control of the...

Large eddy simulation of tip-leakage flow in an axial flow fan

NASA Astrophysics Data System (ADS)

Park, Keuntae; Choi, Haecheon; Choi, Seokho; Sa, Yongcheol; Kwon, Oh-Kyoung

2016-11-01

An axial flow fan with a shroud generates a complicated tip-leakage flow by the interaction of the axial flow with the fan blades and shroud near the blade tips. In this study, large eddy simulation is performed for tip-leakage flow in a forward-swept axial flow fan inside an outdoor unit of an air-conditioner, operating at the design condition of the Reynolds number of 547,000 based on the radius of blade tip and the tip velocity. A dynamic global model is used for a subgrid-scale model, and an immersed boundary method in a non-inertial reference frame is adopted. The present simulation clearly reveals the generation and evolution of tip-leakage vortex near the blade tip by the leakage flow. At the inception of the leakage vortex near the leading edge of the suction-side of the blade tip, the leakage vortex is composed of unsteady multiple vortices containing high-frequency fluctuations. As the leakage vortex develops downstream along a slant line toward the following blade, large and meandering movements of the leakage vortex are observed. Thus low-frequency broad peaks of velocity and pressure occur near the pressure surface. Supported by the KISTI Supercomputing Center (KSC-2016-C3-0027).

Aeroelastic behavior of composite rotor blades with swept tips

NASA Technical Reports Server (NTRS)

Yuan, Kuo-An; Friedmann, Peretz P.; Venkatesan, Comandur

1992-01-01

This paper presents an analytical study of the aeroelastic behavior of composite rotor blades with straight and swept tips. The blade is modeled by beam type finite elements. A single finite element is used to model the swept tip. The nonlinear equations of motion for the finite element model are derived using Hamilton's principle and based on a moderate deflection theory and accounts for: arbitrary cross-sectional shape, pretwist, generally anisotropic material behavior, transverse shears and out-of-plane warping. Numerical results illustrating the effects of tip sweep, anhedral and composite ply orientation on blade aeroelastic behavior are presented. It is shown that composite ply orientation has a substantial effect on blade stability. At low thrust conditions, certain ply orientations can cause instability in the lag mode. The flap-torsion coupling associated with tip sweep can also induce aeroelastic instability in the blade. This instability can be removed by appropriate ply orientation in the composite construction.

Multi-objective shape optimization of runner blade for Kaplan turbine

NASA Astrophysics Data System (ADS)

Semenova, A.; Chirkov, D.; Lyutov, A.; Chemy, S.; Skorospelov, V.; Pylev, I.

2014-03-01

Automatic runner shape optimization based on extensive CFD analysis proved to be a useful design tool in hydraulic turbomachinery. Previously the authors developed an efficient method for Francis runner optimization. It was successfully applied to the design of several runners with different specific speeds. In present work this method is extended to the task of a Kaplan runner optimization. Despite of relatively simpler blade shape, Kaplan turbines have several features, complicating the optimization problem. First, Kaplan turbines normally operate in a wide range of discharges, thus CFD analysis of each variant of the runner should be carried out for several operation points. Next, due to a high specific speed, draft tube losses have a great impact on the overall turbine efficiency, and thus should be accurately evaluated. Then, the flow in blade tip and hub clearances significantly affects the velocity profile behind the runner and draft tube behavior. All these features are accounted in the present optimization technique. Parameterization of runner blade surface using 24 geometrical parameters is described in details. For each variant of runner geometry steady state three-dimensional turbulent flow computations are carried out in the domain, including wicket gate, runner, draft tube, blade tip and hub clearances. The objectives are maximization of efficiency in best efficiency and high discharge operation points, with simultaneous minimization of cavitation area on the suction side of the blade. Multiobjective genetic algorithm is used for the solution of optimization problem, requiring the analysis of several thousands of runner variants. The method is applied to optimization of runner shape for several Kaplan turbines with different heads.

A Method to Further Reduce the Perceived Noise of Low Tip Speed Fans

NASA Technical Reports Server (NTRS)

Dittmar, James H.

2000-01-01

The use of low tip speed, high bypass ratio fans is a method for reducing the noise of turbofan jet engines. These fans typically have a low number of rotor blades and a number of stator vanes sufficient to achieve cut-off of the blade passing tone. Their perceived noise levels are typically dominated by broadband noise caused by the rotor wake turbulence - stator interaction mechanism. A 106 bladed, 1100 ft/sec takeoff tip speed fan, the Alternative Low Noise Fan, has been tested and shown to have reduced broadband noise. This reduced noise is believed to be the result of the high rotor blade number. Although this fan with 106 blades would not be practical with materials as they exist today, a fan with 50 or so blades could be practically realized. A noise estimate has indicated that such a 50 bladed, low tip speed fan could be 2 to 3 EPNdB quieter than an 18 bladed fan. If achieved, this level of noise reduction would be significant and points to the use of a high blade number, low tip speed fan as a possible configuration for reduced fan noise.

The effect of tip vortex structure on helicopter noise due to blade/vortex interaction

NASA Technical Reports Server (NTRS)

Wolf, T. L.; Widnall, S. E.

1978-01-01

A potential cause of helicopter impulsive noise, commonly called blade slap, is the unsteady lift fluctuation on a rotor blade due to interaction with the vortex trailed from another blade. The relationship between vortex structure and the intensity of the acoustic signal is investigated. The analysis is based on a theoretical model for blade/vortex interaction. Unsteady lift on the blades due to blade/vortex interaction is calculated using linear unsteady aerodynamic theory, and expressions are derived for the directivity, frequency spectrum, and transient signal of the radiated noise. An inviscid rollup model is used to calculate the velocity profile in the trailing vortex from the spanwise distribution of blade tip loading. A few cases of tip loading are investigated, and numerical results are presented for the unsteady lift and acoustic signal due to blade/vortex interaction. The intensity of the acoustic signal is shown to be quite sensitive to changes in tip vortex structure.

Rotor blade assembly having internal loading features

DOE Office of Scientific and Technical Information (OSTI.GOV)

Soloway, Daniel David

Rotor blade assemblies and wind turbines are provided. A rotor blade assembly includes a rotor blade having exterior surfaces defining a pressure side, a suction side, a leading edge and a trailing edge each extending between a tip and a root, the rotor blade defining a span and a chord, the exterior surfaces defining an interior of the rotor blade. The rotor blade assembly further includes a loading assembly, the loading assembly including a weight disposed within the interior and movable generally along the span of the rotor blade, the weight connected to a rotor blade component such that movementmore » of the weight towards the tip causes application of a force to the rotor blade component by the weight. Centrifugal force due to rotation of the rotor blade biases the weight towards the tip.« less

Prediction of Unshsrouded Rotor Blade Tip Heat Transfer

NASA Technical Reports Server (NTRS)

Ameri, A. A.; Steinthorsson, E.

1994-01-01

The rate of heat transfer on the tip of a turbine rotor blade and on the blade surface in the vicinity of the tip, was successfully predicted. The computations were performed with a multiblock computer code which solves the Reynolds Averaged Navier-Stokes equations using an efficient multigrid method. The case considered for the present calculations was the Space Shuttle Main Engine (SSME) high pressure fuel side turbine. The predictions of the blade tip heat transfer agreed reasonably well with the experimental measurements using the present level of grid refinement. On the tip surface, regions with high rate of heat transfer was found to exist close to the pressure side and suction side edges. Enhancement of the heat transfer was also observed on the blade surface near the tip. Further comparison of the predictions was performed with results obtained from correlations based on fully developed channel flow.

Performance and Vibratory Loads Data From a Wind-Tunnel Test of a Model Helicopter Main-Rotor Blade With a Paddle-Type Tip

NASA Technical Reports Server (NTRS)

Yeager, William T., Jr.; Noonan, Kevin W.; Singleton, Jeffrey D.; Wilbur, Matthew L.; Mirick, Paul H.

1997-01-01

An investigation was conducted in the Langley Transonic Dynamics Tunnel to obtain data to permit evaluation of paddle-type tip technology for possible use in future U.S. advanced rotor designs. Data was obtained for both a baseline main-rotor blade and a main-rotor blade with a paddle-type tip. The baseline and paddle-type tip blades were compared with regard to rotor performance, oscillatory pitch-link loads, and 4-per-rev vertical fixed-system loads. Data was obtained in hover and forward flight over a nominal range of advance ratios from 0.15 to 0.425. Results indicate that the paddle-type tip offers no performance improvements in either hover or forward flight. Pitch-link oscillatory loads for the paddle-type tip are higher than for the baseline blade, whereas 4-per-rev vertical fixed-system loads are generally lower.

An asymptotic theory of supersonic propeller noise

NASA Technical Reports Server (NTRS)

Envia, Edmane

1992-01-01

A theory for predicting the noise field of supersonic propellers with realistic blade geometries is presented. The theory, which utilizes a large-blade-count approximation, provides an efficient formula for predicting the radiation of sound from all three sources of propeller noise. Comparisons with a full numerical integration indicate that the levels predicted by this formula are quite accurate. Calculations also show that, for high speed propellers, the noise radiated by the Lighthill quadrupole source is rather substantial when compared with the noise radiated by the blade thickness and loading sources. Results from a preliminary application of the theory indicate that the peak noise level generated by a supersonic propeller initially increases with increasing tip helical Mach number, but is eventually reaches a plateau and does not increase further. The predicted trend shows qualitative agreement with the experimental observations.

Characterization of Aeromechanics Response and Instability in Fans, Compressors, and Turbine Blades

NASA Technical Reports Server (NTRS)

Tan, Choon S.

2003-01-01

This study investigated the effect of interaction between tip clearance flow, steady and unsteady upstream wakes in rotor and stator blade rows in terms of blade forced response. In a stator blade row, the interaction of steady wakes in the upstream rotor frame with the stator imply a blade forced response whose spectrum contains the Blade passing frequency (BPF) and its harmonics, with a decaying amplitude as the frequency increases. When the incoming wakes are unsteady, however, the spectrum of blade excitation exhibits unexpectedly amplified high frequencies due to the modulation of BPF with the fluctuation frequency. In a rotor blade row, a tip flow instability has been demonstrated with a frequency (TVF) equal to 0.45 times the Blade Passing frequency corresponding to a reduced frequency (F(sub c) (sup +)) of 0.7. Under uniform inlet flow conditions, the frequency and spatial content of the tip flow region have been characterized. The disturbance TVF was the dominant disturbance in the flow field and was found to imply variations of the pressure coefficient of more than 30% on the blade tip (between 35% to 90% chord) and in the rotor-generated wake (from 75% to 100% hub-to-tip position). In an attempt to better understand the origin of the instability, the structure of the tip flow has also been analyzed. The interface between the tip flow region and the core flow has been found to have periodical wave-like flow patterns which proceed downstream at a speed of approximately 0.42 times the core flow speed at a frequency corresponding to TVF. A list of conclusions derived from these interactions is presented.

Blade resonance parameter identification based on tip-timing method without the once-per revolution sensor

NASA Astrophysics Data System (ADS)

Guo, Haotian; Duan, Fajie; Zhang, Jilong

2016-01-01

Blade tip-timing is the most effective method for blade vibration online measurement of turbomachinery. In this article a synchronous resonance vibration measurement method of blade based on tip-timing is presented. This method requires no once-per revolution sensor which makes it more generally applicable in the condition where this sensor is difficult to install, especially for the high-pressure rotors of dual-rotor engines. Only three casing mounted probes are required to identify the engine order, amplitude, natural frequency and the damping coefficient of the blade. A method is developed to identify the blade which a tip-timing data belongs to without once-per revolution sensor. Theoretical analyses of resonance parameter measurement are presented. Theoretic error of the method is investigated and corrected. Experiments are conducted and the results indicate that blade resonance parameter identification is achieved without once-per revolution sensor.

Characteristics of tip-leakage flow in an axial fan

NASA Astrophysics Data System (ADS)

Park, Keuntae; Choi, Haecheon; Choi, Seokho; Sa, Yongcheol

2014-11-01

An axial fan with a shroud generates complicated vortical structures by the interaction of the axial flow with the fan blades and shroud near the blade tips. Large eddy simulation (LES) is performed for flow through a forward-swept axial fan, operating at the design condition of Re = 547,000 based on the radius of blade tip and the tip velocity. A dynamic global model (Lee et al. 2010) is used for a subgrid-scale model, and an immersed boundary method in a non-inertial reference frame (Kim & Choi 2006) is adopted for the present simulation. It is found that two vortical structures are formed near the blade tip: the main tip leakage vortex (TLV) and the auxiliary TLV. The main TLV is initiated near the leading edge, develops downstream, and impinges on the pressure surface of the next blade, where the pressure fluctuations and turbulence intensity become high. On the other hand, the auxiliary TLV is initiated at the aft part of the blade but is relatively weak such that it merges with the main TLV. Supported by the KISTI Supercomputing Center (KSC-2014-C2-014).

Tracking Blade Tip Vortices for Numerical Flow Simulations of Hovering Rotorcraft

NASA Technical Reports Server (NTRS)

Kao, David L.

2016-01-01

Blade tip vortices generated by a helicopter rotor blade are a major source of rotor noise and airframe vibration. This occurs when a vortex passes closely by, and interacts with, a rotor blade. The accurate prediction of Blade Vortex Interaction (BVI) continues to be a challenge for Computational Fluid Dynamics (CFD). Though considerable research has been devoted to BVI noise reduction and experimental techniques for measuring the blade tip vortices in a wind tunnel, there are only a handful of post-processing tools available for extracting vortex core lines from CFD simulation data. In order to calculate the vortex core radius, most of these tools require the user to manually select a vortex core to perform the calculation. Furthermore, none of them provide the capability to track the growth of a vortex core, which is a measure of how quickly the vortex diffuses over time. This paper introduces an automated approach for tracking the core growth of a blade tip vortex from CFD simulations of rotorcraft in hover. The proposed approach offers an effective method for the quantification and visualization of blade tip vortices in helicopter rotor wakes. Keywords: vortex core, feature extraction, CFD, numerical flow visualization

An experimental study of the compressor rotor blade boundary layer

NASA Technical Reports Server (NTRS)

Pouagare, M.; Lakshminarayana, B.; Galmes, J. M.

1984-01-01

The three-dimensional turbulent boundary layer developing on a rotor blade of an axial flow compressor was measured using a miniature 'x' configuration hot-wire probe. The measurements were carried out at nine radial locations on both surfaces of the blade at various chordwise locations. The data derived includes streamwise and radial mean velocities and turbulence intensities. The validity of conventional velocity profiles such as the 'power law profile' for the streamwise profile, and Mager and Eichelbrenner's for the radial profile, is examined. A modification to Mager's crossflow profile is proposed. Away from the blade tip, the streamwise component of the blade boundary layer seems to be mainly influenced by the streamwise pressure gradient. Near the tip of the blade, the behavior of the blade boundary layer is affected by the tip leakage flow and the annulus wall boundary layer. The 'tangential blockage' due to the blade boundary layer is derived from the data. The profile losses are found to be less than that of an equivalent cascade, except in the tip region of the blade.

An acoustic experimental and theoretical investigation of single disc propellers

NASA Technical Reports Server (NTRS)

Bumann, Elizabeth A.; Korkan, Kenneth D.

1989-01-01

An experimental study of the acoustic field associated with two, three, and four blade propeller configurations with a blade root angle of 50 deg was performed in the Texas A&M University 5 ft. x 6 ft. acoustically-insulated subsonic wind tunnel. A waveform analysis package was utilized to obtain experimental acoustic time histories, frequency spectra, and overall sound pressure level (OASPL) and served as a basis for comparison to the theoretical acoustic compact source theory of Succi (1979). Valid for subsonic tip speeds, the acoustic analysis replaced each blade by an array of spiraling point sources which exhibited a unique force vector and volume. The computer analysis of Succi was modified to include a propeller performance strip analysis which used a NACA 4-digit series airfoil data bank to calculate lift and drag for each blade segment given the geometry and motion of the propeller. Theoretical OASPL predictions were found to moderately overpredict experimental values for all operating conditions and propeller configurations studied.

Acoustic measurements of a full-scale rotor with four tip shapes. Volume 1: Text, appendices A and B

NASA Technical Reports Server (NTRS)

Mosher, M.

1984-01-01

A full-scale helicopter with four different blade-tip geometries was tested in the 40- by 80-foot wind tunnel at Ames Research Center. Performance, loads, and noise were measured. The four tip shapes tested were rectangular, tapered, swept, and swept-tapered. Noise measurements from that test are presented in the form of tables and plots. The noise data include measurements of the sound pressure level in dB, dBA, and tone-corrected PNdB, for all of the conditions tested. Detailed measurements, 1/3-octave spectra and time-histories for some selected data are included as well as plots of dBA as function of test condition. Some performance measurements are given to aid interpretation of the noise data.

WT - WIND TUNNEL PERFORMANCE ANALYSIS

NASA Technical Reports Server (NTRS)

Viterna, L. A.

1994-01-01

WT was developed to calculate fan rotor power requirements and output thrust for a closed loop wind tunnel. The program uses blade element theory to calculate aerodynamic forces along the blade using airfoil lift and drag characteristics at an appropriate blade aspect ratio. A tip loss model is also used which reduces the lift coefficient to zero for the outer three percent of the blade radius. The application of momentum theory is not used to determine the axial velocity at the rotor plane. Unlike a propeller, the wind tunnel rotor is prevented from producing an increase in velocity in the slipstream. Instead, velocities at the rotor plane are used as input. Other input for WT includes rotational speed, rotor geometry, and airfoil characteristics. Inputs for rotor blade geometry include blade radius, hub radius, number of blades, and pitch angle. Airfoil aerodynamic inputs include angle at zero lift coefficient, positive stall angle, drag coefficient at zero lift coefficient, and drag coefficient at stall. WT is written in APL2 using IBM's APL2 interpreter for IBM PC series and compatible computers running MS-DOS. WT requires a CGA or better color monitor for display. It also requires 640K of RAM and MS-DOS v3.1 or later for execution. Both an MS-DOS executable and the source code are provided on the distribution medium. The standard distribution medium for WT is a 5.25 inch 360K MS-DOS format diskette in PKZIP format. The utility to unarchive the files, PKUNZIP, is also included. WT was developed in 1991. APL2 and IBM PC are registered trademarks of International Business Machines Corporation. MS-DOS is a registered trademark of Microsoft Corporation. PKUNZIP is a registered trademark of PKWare, Inc.

A Blade Tip Timing Method Based on a Microwave Sensor

PubMed Central

Zhang, Jilong; Duan, Fajie; Niu, Guangyue; Jiang, Jiajia; Li, Jie

2017-01-01

Blade tip timing is an effective method for blade vibration measurements in turbomachinery. This method is increasing in popularity because it is non-intrusive and has several advantages over the conventional strain gauge method. Different kinds of sensors have been developed for blade tip timing, including optical, eddy current and capacitance sensors. However, these sensors are unsuitable in environments with contaminants or high temperatures. Microwave sensors offer a promising potential solution to overcome these limitations. In this article, a microwave sensor-based blade tip timing measurement system is proposed. A patch antenna probe is used to transmit and receive the microwave signals. The signal model and process method is analyzed. Zero intermediate frequency structure is employed to maintain timing accuracy and dynamic performance, and the received signal can also be used to measure tip clearance. The timing method uses the rising and falling edges of the signal and an auto-gain control circuit to reduce the effect of tip clearance change. To validate the accuracy of the system, it is compared experimentally with a fiber optic tip timing system. The results show that the microwave tip timing system achieves good accuracy. PMID:28492469

An optical fiber bundle sensor for tip clearance and tip timing measurements in a turbine rig.

PubMed

García, Iker; Beloki, Josu; Zubia, Joseba; Aldabaldetreku, Gotzon; Illarramendi, María Asunción; Jiménez, Felipe

2013-06-05

When it comes to measuring blade-tip clearance or blade-tip timing in turbines, reflective intensity-modulated optical fiber sensors overcome several traditional limitations of capacitive, inductive or discharging probe sensors. This paper presents the signals and results corresponding to the third stage of a multistage turbine rig, obtained from a transonic wind-tunnel test. The probe is based on a trifurcated bundle of optical fibers that is mounted on the turbine casing. To eliminate the influence of light source intensity variations and blade surface reflectivity, the sensing principle is based on the quotient of the voltages obtained from the two receiving bundle legs. A discrepancy lower than 3% with respect to a commercial sensor was observed in tip clearance measurements. Regarding tip timing measurements, the travel wave spectrum was obtained, which provides the average vibration amplitude for all blades at a particular nodal diameter. With this approach, both blade-tip timing and tip clearance measurements can be carried out simultaneously. The results obtained on the test turbine rig demonstrate the suitability and reliability of the type of sensor used, and suggest the possibility of performing these measurements in real turbines under real working conditions.

A Blade Tip Timing Method Based on a Microwave Sensor.

PubMed

Zhang, Jilong; Duan, Fajie; Niu, Guangyue; Jiang, Jiajia; Li, Jie

2017-05-11

Blade tip timing is an effective method for blade vibration measurements in turbomachinery. This method is increasing in popularity because it is non-intrusive and has several advantages over the conventional strain gauge method. Different kinds of sensors have been developed for blade tip timing, including optical, eddy current and capacitance sensors. However, these sensors are unsuitable in environments with contaminants or high temperatures. Microwave sensors offer a promising potential solution to overcome these limitations. In this article, a microwave sensor-based blade tip timing measurement system is proposed. A patch antenna probe is used to transmit and receive the microwave signals. The signal model and process method is analyzed. Zero intermediate frequency structure is employed to maintain timing accuracy and dynamic performance, and the received signal can also be used to measure tip clearance. The timing method uses the rising and falling edges of the signal and an auto-gain control circuit to reduce the effect of tip clearance change. To validate the accuracy of the system, it is compared experimentally with a fiber optic tip timing system. The results show that the microwave tip timing system achieves good accuracy.

An Optical Fiber Bundle Sensor for Tip Clearance and Tip Timing Measurements in a Turbine Rig

PubMed Central

García, Iker; Beloki, Josu; Zubia, Joseba; Aldabaldetreku, Gotzon; Illarramendi, María Asunción; Jiménez, Felipe

2013-01-01

When it comes to measuring blade-tip clearance or blade-tip timing in turbines, reflective intensity-modulated optical fiber sensors overcome several traditional limitations of capacitive, inductive or discharging probe sensors. This paper presents the signals and results corresponding to the third stage of a multistage turbine rig, obtained from a transonic wind-tunnel test. The probe is based on a trifurcated bundle of optical fibers that is mounted on the turbine casing. To eliminate the influence of light source intensity variations and blade surface reflectivity, the sensing principle is based on the quotient of the voltages obtained from the two receiving bundle legs. A discrepancy lower than 3% with respect to a commercial sensor was observed in tip clearance measurements. Regarding tip timing measurements, the travel wave spectrum was obtained, which provides the average vibration amplitude for all blades at a particular nodal diameter. With this approach, both blade-tip timing and tip clearance measurements can be carried out simultaneously. The results obtained on the test turbine rig demonstrate the suitability and reliability of the type of sensor used, and suggest the possibility of performing these measurements in real turbines under real working conditions. PMID:23739163

Numerical investigation of the unsteady tip leakage flow and rotating stall inception in a transonic compressor

NASA Astrophysics Data System (ADS)

Zhang, Yanfeng; Lu, Xingen; Chu, Wuli; Zhu, Junqiang

2010-08-01

It is well known that tip leakage flow has a strong effect on the compressor performance and stability. This paper reports on a numerical investigation of detailed flow structures in an isolated transonic compressor rotor-NASA Rotor 37 at near stall and stalled conditions aimed at improving understanding of changes in 3D tip leakage flow structures with rotating stall inception. Steady and unsteady 3D Navier-Stokes analyses were conducted to investigate flow structures in the same rotor. For steady analysis, the predicted results agree well with the experimental data for the estimation of compressor rotor global performance. For unsteady flow analysis, the unsteady flow nature caused by the breakdown of the tip leakage vortex in blade tip region in the transonic compressor rotor at near stall condition has been captured with a single blade passage. On the other hand, the time-accurate unsteady computations of multi-blade passage at near stall condition indicate that the unsteady breakdown of the tip leakage vortex triggered the short length-scale — spike type rotating stall inception at blade tip region. It was the forward spillage of the tip leakage flow at blade leading edge resulting in the spike stall inception. As the mass flow ratio is decreased, the rotating stall cell was further developed in the blade passage.

Hover and Wind-Tunnel Testing of Shrouded Rotors for Improved Micro Air Vehicle Design

DTIC Science & Technology

2008-01-01

and the shroud surface pressure distributions. The uniformity of the wake was improved by the presence of the shrouds and by decreasing the blade tip...213 3.35 Effect of blade tip clearance on shrouded-rotor exit-plane wake profiles215 3.36 Effects of changing blade tip clearance on induced...Wright [139] developed a vortex wake model for heavily loaded ducted fans, in which the “inner vortex sheets [shed from the blades ] move at a different

Chem-Braze Abradable Seal Attachment

DTIC Science & Technology

1980-05-01

bonding system for attaching sintered abradable seals such as FELTMETAL® to titanium -, steel- and nickel-base compressor blade tip-shrouds has been... blade tip-shrouds was developed. The improved Chem-Braze system incorporates glycerin as an inhibitor to prevent premature evaporation which prolongs...compressor blade tip-shrouds using the improved Chem-Braze system compared to attachment with gold-nickel braze. p. p. FORM . . yn

High Temperature Investigations into an Active Turbine Blade Tip Clearance Control Concept

NASA Technical Reports Server (NTRS)

Taylor, Shawn; Steinetz, Bruce M.; Oswald, Jay J.

2007-01-01

System studies have shown the benefits of reducing blade tip clearances in modern turbine engines. Minimizing blade tip clearances throughout the engine will contribute materially to meeting NASA s Ultra-Efficient Engine Technology (UEET) turbine engine project goals. NASA GRC is examining two candidate approaches including rub-avoidance and regeneration which are explained in subsequent slides.

High Temperature Investigations into an Active Turbine Blade Tip Clearance Control Concept

NASA Technical Reports Server (NTRS)

Taylor, Shawn C.; Steinetz, Bruce; Oswald, Jay J.

2008-01-01

System studies have shown the benefits of reducing blade tip clearances in modern turbine engines. Minimizing blade tip clearances throughout the engine will contribute materially to meeting NASA s Ultra-Efficient Engine Technology (UEET) turbine engine project goals. NASA GRC is examining two candidate approaches including rub-avoidance and regeneration which are explained in subsequent slides.

Effect of Crystal Orientation on Analysis of Single-Crystal, Nickel-Based Turbine Blade Superalloys

NASA Technical Reports Server (NTRS)

Swanson, G. R.; Arakere, N. K.

2000-01-01

High-cycle fatigue-induced failures in turbine and turbopump blades is a pervasive problem. Single-crystal nickel turbine blades are used because of their superior creep, stress rupture, melt resistance, and thermomechanical fatigue capabilities. Single-crystal materials have highly orthotropic properties making the position of the crystal lattice relative to the part geometry a significant and complicating factor. A fatigue failure criterion based on the maximum shear stress amplitude on the 24 octahedral and 6 cube slip systems is presented for single-crystal nickel superalloys (FCC crystal). This criterion greatly reduces the scatter in uniaxial fatigue data for PWA 1493 at 1,200 F in air. Additionally, single-crystal turbine blades used in the Space Shuttle main engine high pressure fuel turbopump/alternate turbopump are modeled using a three-dimensional finite element (FE) model. This model accounts for material orthotrophy and crystal orientation. Fatigue life of the blade tip is computed using FE stress results and the failure criterion that was developed. Stress analysis results in the blade attachment region are also presented. Results demonstrate that control of crystallographic orientation has the potential to significantly increase a component's resistance to fatigue crack growth without adding additional weight or cost.

Influence of blade tip rounding on tip leakage vortex cavitation of axial flow pump

NASA Astrophysics Data System (ADS)

Wu, S. Q.; Shi, W. D.; Zhang, D. S.; Yao, J.; Cheng, C.

2013-12-01

Tip leakage flow in axial flow pumps is mainly caused by the tip clearance, which is the main cause of tip leakage vortex cavitation and blade tip cavitation erosion. In order to improve tip clearance flow and reduce TLV cavitation, four schemes were adopted to the round blade tip. These are: no tip rounding, one time tip clearance tip rounding, two times tip clearance tip rounding, four times tip clearance tip rounding. Using SST k-ω turbulence model and Zwart cavitation model in CFX software, this simulation obtained four kinds of inner flow field results. The numerical results indicated that with the increase of r*, NPSHc gradually increased and the cavitation performance reduced. However, corner vortex was eliminated so that cavitation in gap was restrained. But TLV vorticity increased and cavitation's range here had a little expansion. Combined with the research of this paper and the different analyses of four schemes, we recommend adopting the two times of the tip clearance rounding.

Comparison of calculated and measured pressures on straight and swept-tip model rotor blades

NASA Technical Reports Server (NTRS)

Tauber, M. E.; Chang, I. C.; Caughey, D. A.; Phillipe, J. J.

1983-01-01

Using the quasi-steady, full potential code, ROT22, pressures were calculated on straight and swept tip model helicopter rotor blades at advance ratios of 0.40 and 0.45, and into the transonic tip speed range. The calculated pressures were compared with values measured in the tip regions of the model blades. Good agreement was found over a wide range of azimuth angles when the shocks on the blade were not too strong. However, strong shocks persisted longer than predicted by ROT22 when the blade was in the second quadrant. Since the unsteady flow effects present at high advance ratios primarily affect shock waves, the underprediction of shock strengths is attributed to the simplifying, quasi-steady, assumption made in ROT22.

System Developed for Real-Time Blade-Flutter Monitoring in the Wind Tunnel

NASA Technical Reports Server (NTRS)

Kurkov, Anatole P.; Dhadwal, Harbans S.; Radzikowski, mark; Strukov, Dmitri

2005-01-01

A real-time system has been developed to monitor flutter vibrations in turbomachinery. The system is designed for continuous processing of blade tip timing data at a rate of 10 MB/sec. A USB 2.0 interface provides uninterrupted real-time processing of the data, and the blade-tip arrival times are measured with a 50-MHz oscillator and a 24-bit pipelined architecture counter. The input stage includes a glitch catcher, which reduces the probability of detecting a ghost blade to negligible levels. A graphical user interface provides online interrogation of any blade tip from any light probe sensor. Alternatively, data from all blades and all sensors can be superimposed into a single composite scatter plot displaying the vibration amplitude of each blade.

Computation of the tip vortex flowfield for advanced aircraft propellers

NASA Technical Reports Server (NTRS)

Tsai, Tommy M.; Dejong, Frederick J.; Levy, Ralph

1988-01-01

The tip vortex flowfield plays a significant role in the performance of advanced aircraft propellers. The flowfield in the tip region is complex, three-dimensional and viscous with large secondary velocities. An analysis is presented using an approximate set of equations which contains the physics required by the tip vortex flowfield, but which does not require the resources of the full Navier-Stokes equations. A computer code was developed to predict the tip vortex flowfield of advanced aircraft propellers. A grid generation package was developed to allow specification of a variety of advanced aircraft propeller shapes. Calculations of the tip vortex generation on an SR3 type blade at high Reynolds numbers were made using this code and a parametric study was performed to show the effect of tip thickness on tip vortex intensity. In addition, calculations of the tip vortex generation on a NACA 0012 type blade were made, including the flowfield downstream of the blade trailing edge. Comparison of flowfield calculations with experimental data from an F4 blade was made. A user's manual was also prepared for the computer code (NASA CR-182178).

Propeller noise caused by blade tip radial forces

NASA Technical Reports Server (NTRS)

Hanson, D. B.

1986-01-01

New experimental evidence which indicates the presence of leading edge and tip edge vortex flow on Prop-Fans is examined, and performance and noise consequences are addressed. It was shown that the tip edge vortex is a significant noise source, particularly for unswept Prop-Fan blades. Preliminary calculations revealed that the addition of the tip side edge source to single rotation Prop-Fans during take off conditions improved the agreement between experiment and theory at blade passing frequency. At high-speed conditions such as the Prop-Fan cruise point, the tip loading effect tends to cancel thickness noise.

Turbine blade with contoured chamfered squealer tip

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lee, Ching-Pang

2014-12-30

A squealer tip formed from a pressure side tip wall and a suction side tip wall extending radially outward from a tip of the turbine blade is disclosed. The pressure and suction side tip walls may be positioned along the pressure sidewall and the suction sidewall of the turbine blade, respectively. The pressure side tip wall may include a chamfered leading edge with film cooling holes having exhaust outlets positioned therein. An axially extending tip wall may be formed from at least two outer linear surfaces joined together at an intersection forming a concave axially extending tip wall. The axiallymore » extending tip wall may include a convex inner surface forming a radially outer end to an inner cavity forming a cooling system. The cooling system may include one or more film cooling holes in the axially extending tip wall proximate to the suction sidewall, which promotes increased cooling at the pressure and suction sidewalls.« less

Large Eddy Simulation of Vertical Axis Wind Turbine wakes; Part I: from the airfoil performance to the very far wake

NASA Astrophysics Data System (ADS)

Chatelain, Philippe; Duponcheel, Matthieu; Caprace, Denis-Gabriel; Marichal, Yves; Winckelmans, Gregoire

2017-11-01

A vortex particle-mesh (VPM) method with immersed lifting lines has been developed and validated. Based on the vorticity-velocity formulation of the Navier-Stokes equations, it combines the advantages of a particle method and of a mesh-based approach. The immersed lifting lines handle the creation of vorticity from the blade elements and its early development. Large-eddy simulation (LES) of vertical axis wind turbine (VAWT) flows is performed. The complex wake development is captured in detail and over up to 15 diameters downstream: from the blades to the near-wake coherent vortices and then through the transitional ones to the fully developed turbulent far wake (beyond 10 rotor diameters). The statistics and topology of the mean flow are studied with respect to the VAWT geometry and its operating point. The computational sizes also allow insights into the detailed unsteady vortex dynamics and topological flow features, such as a recirculation region influenced by the tip speed ratio and the rotor geometry.

An analysis of blade vortex interaction aerodynamics and acoustics

NASA Technical Reports Server (NTRS)

Lee, D. J.

1985-01-01

The impulsive noise associated with helicopter flight due to Blade-Vortex Interaction, sometimes called blade slap is analyzed especially for the case of a close encounter of the blade-tip vortex with a following blade. Three parts of the phenomena are considered: the tip-vortex structure generated by the rotating blade, the unsteady pressure produced on the following blade during the interaction, and the acoustic radiation due to the unsteady pressure field. To simplify the problem, the analysis was confined to the situation where the vortex is aligned parallel to the blade span in which case the maximum acoustic pressure results. Acoustic radiation due to the interaction is analyzed in space-fixed coordinates and in the time domain with the unsteady pressure on the blade surface as the source of chordwise compact, but spanwise non-compact radiation. Maximum acoustic pressure is related to the vortex core size and Reynolds number which are in turn functions of the blade-tip aerodynamic parameters. Finally noise reduction and performance are considered.

An on-line calibration technique for improved blade by blade tip clearance measurement

NASA Astrophysics Data System (ADS)

Sheard, A. G.; Westerman, G. C.; Killeen, B.

A description of a capacitance-based tip clearance measurement system which integrates a novel technique for calibrating the capacitance probe in situ is presented. The on-line calibration system allows the capacitance probe to be calibrated immediately prior to use, providing substantial operational advantages and maximizing measurement accuracy. The possible error sources when it is used in service are considered, and laboratory studies of performance to ascertain their magnitude are discussed. The 1.2-mm diameter FM capacitance probe is demonstrated to be insensitive to variations in blade tip thickness from 1.25 to 1.45 mm. Over typical compressor blading the probe's range was four times the variation in blade to blade clearance encountered in engine run components.

A combined piezoelectric composite actuator and its application to wing/blade tips

NASA Astrophysics Data System (ADS)

Ha, Kwangtae

A novel combined piezoelectric-composite actuator configuration is proposed and analytically modeled in this work. The actuator is a low complexity, active compliant mechanism obtained by coupling a modified star cross sectional configuration composite beam with a helicoidal bimorph piezoelectric actuator coiled around it. This novel actuator is a good candidate as a hinge tension-torsion bar actuator for a helicopter rotor blade flap or blade tip and mirror rotational positioning. In the wing tip case, the tip deflection angle is different only according to the aerodynamic moment depending on the hinge position of the actuator along the chord and applied voltage because there is no centrifugal force. For an active blade tip subject to incompressible flow and 2D quasi steady airloads, its twist angle is related not only to aerodynamic moment and applied voltage but also to coupling terms, such as the trapeze effect and the tennis racquet effect. Results show the benefit of hinge position aft of the aerodynamic center, such that the blade tip response is amplified by airloads. Contrary to this effect, results also show that the centrifugal effects and inertial effect cause an amplitude reduction in the response. Summation of these effects determines the overall blade tip response. The results for a certain hinge position of Xh=1.5% chord aft of the quarter chord point proves that the tip deflection target design range of beta ∈ [-2,+2] can be achieved for all pitch angle configurations chosen.

Low-coherence interferometric tip-clearance probe

NASA Astrophysics Data System (ADS)

Kempe, Andreas; Schlamp, Stefan; Rösgen, Thomas; Haffner, Ken

2003-08-01

We propose an all-fiber, self-calibrating, economical probe that is capable of near-real-time, single-port, simultaneous blade-to-blade tip-clearance measurements with submillimeter accuracy (typically <100 μm, absolute) in the first stages of a gas turbine. Our probe relies on the interference between backreflected light from the blade tips during the 1-μs blade passage time and a frequency-shifted reference with variable time delay, making use of a low-coherence light source. A single optical fiber of arbitrary length connects the self-contained optics and electronics to the turbine.

A study of the noise radiation from four helicopter rotor blades. [tests in Ames 40 by 20 foot wind tunnel

NASA Technical Reports Server (NTRS)

Lee, A.; Mosher, M.

1978-01-01

Acoustic measurements were taken of a modern helicopter rotor with four blade tip shapes in the NASA Ames 40-by-80-Foot Wind Tunnel. The four tip shapes are: rectangular, swept, trapezoidal, and swept tapered in platform. Acoustic effects due to tip shape changes were studied based on the dBA level, peak noise pressure, and subjective rating. The swept tapered blade was found to be the quietest above an advancing tip Mach number of about 0.9, and the swept blade was the quietest at low speed. The measured high speed impulsive noise was compared with theoretical predictions based on thickness effects; good agreement was found.

On the development of a magnetoresistive sensor for blade tip timing and blade tip clearance measurement systems.

PubMed

Tomassini, R; Rossi, G; Brouckaert, J-F

2016-10-01

A simultaneous blade tip timing (BTT) and blade tip clearance (BTC) measurement system enables the determination of turbomachinery blade vibrations and ensures the monitoring of the existing running gaps between the blade tip and the casing. This contactless instrumentation presents several advantages compared to the well-known telemetry system with strain gauges, at the cost of a more complex data processing procedure. The probes used can be optical, capacitive, eddy current as well as microwaves, everyone with its dedicated electronics and many existing different signal processing algorithms. Every company working in this field has developed its own processing method and sensor technology. Hence, repeating the same test with different instrumentations, the answer is often different. Moreover, rarely it is possible to achieve reliability for in-service measurements. Developments are focused on innovative instrumentations and a common standard. This paper focuses on the results achieved using a novel magnetoresistive sensor for simultaneous tip timing and tip clearance measurements. The sensor measurement principle is described. The sensitivity to gap variation is investigated. In terms of measurement of vibrations, experimental investigations were performed at the Air Force Institute of Technology (ITWL, Warsaw, Poland) in a real aeroengine and in the von Karman Institute (VKI) R2 compressor rig. The advantages and limitations of the magnetoresistive probe for turbomachinery testing are highlighted.

On the development of a magnetoresistive sensor for blade tip timing and blade tip clearance measurement systems

NASA Astrophysics Data System (ADS)

Tomassini, R.; Rossi, G.; Brouckaert, J.-F.

2016-10-01

A simultaneous blade tip timing (BTT) and blade tip clearance (BTC) measurement system enables the determination of turbomachinery blade vibrations and ensures the monitoring of the existing running gaps between the blade tip and the casing. This contactless instrumentation presents several advantages compared to the well-known telemetry system with strain gauges, at the cost of a more complex data processing procedure. The probes used can be optical, capacitive, eddy current as well as microwaves, everyone with its dedicated electronics and many existing different signal processing algorithms. Every company working in this field has developed its own processing method and sensor technology. Hence, repeating the same test with different instrumentations, the answer is often different. Moreover, rarely it is possible to achieve reliability for in-service measurements. Developments are focused on innovative instrumentations and a common standard. This paper focuses on the results achieved using a novel magnetoresistive sensor for simultaneous tip timing and tip clearance measurements. The sensor measurement principle is described. The sensitivity to gap variation is investigated. In terms of measurement of vibrations, experimental investigations were performed at the Air Force Institute of Technology (ITWL, Warsaw, Poland) in a real aeroengine and in the von Karman Institute (VKI) R2 compressor rig. The advantages and limitations of the magnetoresistive probe for turbomachinery testing are highlighted.

Design of a 1500 Ft/Sec, Transonic, High-through-Flow, Single-Stage Axial-Flow Compressor with Low Hub/Tip Ratio

DTIC Science & Technology

1976-10-01

aerodynamic flow field pertaining to the design point is defined on twenty-one stream surfaces, and radial and meridional distributions of significant...full radial equilibrium analysis of the compressor flow field using the streamline curvature solution technique. Through a series of iterations, it...one can assume the blade geometry, solving for the equilibriwn flow field using specified relative flow aigles as input to the aerodynamic program. In

Coupled Aerodynamic-Thermal-Structural (CATS) Analysis

NASA Technical Reports Server (NTRS)

1995-01-01

Coupled Aerodynamic-Thermal-Structural (CATS) Analysis is a focused effort within the Numerical Propulsion System Simulation (NPSS) program to streamline multidisciplinary analysis of aeropropulsion components and assemblies. Multidisciplinary analysis of axial-flow compressor performance has been selected for the initial focus of this project. CATS will permit more accurate compressor system analysis by enabling users to include thermal and mechanical effects as an integral part of the aerodynamic analysis of the compressor primary flowpath. Thus, critical details, such as the variation of blade tip clearances and the deformation of the flowpath geometry, can be more accurately modeled and included in the aerodynamic analyses. The benefits of this coupled analysis capability are (1) performance and stall line predictions are improved by the inclusion of tip clearances and hot geometries, (2) design alternatives can be readily analyzed, and (3) higher fidelity analysis by researchers in various disciplines is possible. The goals for this project are a 10-percent improvement in stall margin predictions and a 2:1 speed-up in multidisciplinary analysis times. Working cooperatively with Pratt & Whitney, the Lewis CATS team defined the engineering processes and identified the software products necessary for streamlining these processes. The basic approach is to integrate the aerodynamic, thermal, and structural computational analyses by using data management and Non-Uniform Rational B-Splines (NURBS) based data mapping. Five software products have been defined for this task: (1) a primary flowpath data mapper, (2) a two-dimensional data mapper, (3) a database interface, (4) a blade structural pre- and post-processor, and (5) a computational fluid dynamics code for aerothermal analysis of the drum rotor. Thus far (1) a cooperative agreement has been established with Pratt & Whitney, (2) a Primary Flowpath Data Mapper has been prototyped and delivered to General Electric Aircraft Engines and Pratt & Whitney for evaluation, (3) a collaborative effort has been initiated with the National Institute of Standards and Testing to develop a Standard Data Access Interface, and (4) a blade tip clearance capability has been implemented into the Structural Airfoil Blade Engineering Routine (SABER) program. We plan to continue to develop the data mappers and data management tools. As progress is made, additional efforts will be made to apply these tools to propulsion system applications.

Tip cap for a turbine rotor blade

DOEpatents

Kimmel, Keith D

2014-03-25

A turbine rotor blade with a spar and shell construction, and a tip cap that includes a row of lugs extending from a bottom side that form dovetail grooves that engage with similar shaped lugs and grooves on a tip end of the spar to secure the tip cap to the spar against radial displacement. The lug on the trailing edge end of the tip cap is aligned perpendicular to a chordwise line of the blade in the trailing edge region in order to minimize stress due to the lugs wanting to bend under high centrifugal loads. A two piece tip cap with lugs at different angles will reduce the bending stress even more.

Study of blade clearance effects on centrifugal pumps

NASA Technical Reports Server (NTRS)

Hoshide, R. K.; Nielson, C. E.

1972-01-01

A program of analysis, design, fabrication, and testing has been conducted to develop and experimentally verify analytical models to predict the effects of impeller blade clearance on centrifugal pumps. The effect of tip clearance on pump efficiency, and the relationship between the head coefficient and torque loss with tip clearance was established. Analysis were performed to determine the cost variation in design, manufacture, and test that would occur between unshrouded and shrouded impellers. An impeller, representative of typical rocket engine impellers, was modified by removing its front shroud to permit variation of its blade clearances. It was tested in water with special instrumentation to provide measurements of blade surface pressures during operation. Pump performance data were obtained from tests at various impeller tip clearances. Blade pressure data were obtained at the nominal tip clearance. Comparisons of predicted and measured data are given.

Turbine blade tip gap reduction system

DOEpatents

Diakunchak, Ihor S.

2012-09-11

A turbine blade sealing system for reducing a gap between a tip of a turbine blade and a stationary shroud of a turbine engine. The sealing system includes a plurality of flexible seal strips extending from a pressure side of a turbine blade generally orthogonal to the turbine blade. During operation of the turbine engine, the flexible seal strips flex radially outward extending towards the stationary shroud of the turbine engine, thereby reducing the leakage of air past the turbine blades and increasing the efficiency of the turbine engine.

Turbine blade tip flow discouragers

DOEpatents

Bunker, Ronald Scott

2000-01-01

A turbine assembly comprises a plurality of rotating blade portions in a spaced relation with a stationery shroud. The rotating blade portions comprise a root section, a tip portion and an airfoil. The tip portion has a pressure side wall and a suction side wall. A number of flow discouragers are disposed on the blade tip portion. In one embodiment, the flow discouragers extend circumferentially from the pressure side wall to the suction side wall so as to be aligned generally parallel to the direction of rotation. In an alternative embodiment, the flow discouragers extend circumferentially from the pressure side wall to the suction side wall so as to be aligned at an angle in the range between about 0.degree. to about 60.degree. with respect to a reference axis aligned generally parallel to the direction of rotation. The flow discouragers increase the flow resistance and thus reduce the flow of hot gas flow leakage for a given pressure differential across the blade tip portion so as to improve overall turbine efficiency.

The Role of Free-Stream Turbulence on High Pressure Turbine Aero-Thermal Stage Interaction

NASA Astrophysics Data System (ADS)

Kopriva, James Earl

Turbulence plays an important role on the aero-thermal performance of modern aircraft engine High Pressure Turbines (HPT). The role of the vane wake and passage turbulence on the downstream blade flow field is an important consideration for both performance and durability. Obtaining measurements to fully characterize the flow field can be challenging and costly in an experimental facility. Advances in computational Fluid Dynamic (CFD) modeling and High Performance Computing (HPC) are providing opportunity to close these measurement gaps. In order for CFD to be adopted, methods need to be both accurate and efficient. Meshing approaches must also be able to resolve complex HPT geometry while maintaining quality adequate for scale-resolved simulations. Therefore, the accuracy of executing scale-resolved simulations with a second-order code on a mesh of prisms and tetrahedrals in Fluent is considered. Before execution of the HPT computational study, a building block approach is taken to gain quantified predictive performance in the modeling approach as well as understanding limitations in lower computational cost modeling approaches. The predictive capability for Reynolds Averaged Navier Stokes (RANS), Hybrid Large Eddy Simulation (LES), and wall-resolved LES turbulence modeling approaches are first assessed for a cylinder in cross-flow at a Reynolds number of 2580. The flow condition and simple geometry facilitate a quick turn-around for modeling assessment before moving the HPT vane study at high Reynolds and Mach number conditions. Modeling approaches are then assessed relative to the experimental measurements of Arts and Rouvroit (1992) on a pitch-line HPT uncooled vane at high Mach and Reynolds numbers conditions with low (0-6%) free-stream turbulence. The current unstructured second-order LES approach agrees with experimental data and is found to be within the equivalent experimental uncertainty when compared to the structured high-ordered solver FDL3DI. The unstructured Hybrid LES aero-thermal predictions are found to be in close agreement with LES predictions and 4 times more computationally efficient. A sliding mesh approach is then used to understand the complex HPT vane and blade stage aero-thermal interaction at 0 and 20% inlet turbulence. A HPT blade has been designed to pair with the uncooled vane of Arts and Rouvroit (1992) to evaluate the impact of passage turbulence and vane wake on the downstream blade boundary layer as well as wake formation and evolution. The learnings from the statistical 2D pitch-line stage simulations are applied to a 3D annular representation of the geometry including endwalls and blade tip clearance to demonstrate the impact of secondary flows on the overall aero-thermal performance. Compared to the 2D pitchline predictions, the vane and blade overall mass average relative total pressure loss for the 3D geometry increases by 73 and 107%, respectively. The blade loss is shown to be largely driven by the formation of the tip vortex. Hybrid LES predictions show that by increasing stage inlet turbulence by 20% results in up to a 40% increase for the surface heat flux on the vane. However, the impact of stage inlet turbulence is found to be secondary compared to the periodic unsteadiness generated by the vane wake on the downstream blade surface heat transfer and mixing.

Alternative blade materials for technical and ecological optimization of a hydraulic pressure machine

NASA Astrophysics Data System (ADS)

Schwyzer, Olivier; Saenger, Nicole

2016-11-01

The Hydraulic Pressure Machine (HPM) is an energy converter to exploit head differences between 0.5 and 2.5 m in small streams and irrigation canals. Previous investigations show that efficiencies above 60% are possible. Several case studies indicate good continuity for aquatic life (e.g. fish) and bed load for the technology. The technology is described as an economically and ecologically viable option for small scale hydropower generation. Primary goal of this research is to improve the HPM blade design regarding its continuity properties by maintaining good efficiency rates. This is done by modifying the blade tip and testing within a large physical model under laboratory condition. Blade tips from steel (conventional - reference case) and a combination of EPDM rubber and steel as sandwich construction (rubber, steel, rubber - adhesive layered) are tested and compared. Both materials reach similar values for hydraulic efficiency (approx. 58%) and mechanical power output (approx. 220 W). The variation of different gap sizes pointed out the importance of small clearance gaps to reach high efficiencies. For assessing the two blade tip materials regarding continuity for aquatic life, fish dummies were led through the wheel. Analysis of slow motion video of dummies hit by the blade show significant advantages for the EPDM blade tip. The EPDM rubber allows to bend and thus reduces the shock and the probability for cuts on the fish dummy. It was shown that blade tips from EPDM have certain advantages regarding continuity compared to standard blade tips from steel. No compromise regarding energy production had to be made. These results from the HPM can be transferred to breast shot water wheel and may be applied for new and retrofitting projects.

Methods and apparatus for twist bend coupled (TCB) wind turbine blades

DOEpatents

Moroz, Emilian Mieczyslaw; LeMieux, David Lawrence; Pierce, Kirk Gee

2006-10-10

A method for controlling a wind turbine having twist bend coupled rotor blades on a rotor mechanically coupled to a generator includes determining a speed of a rotor blade tip of the wind turbine, measuring a current twist distribution and current blade loading, and adjusting a torque of a generator to change the speed of the rotor blade tip to thereby increase an energy capture power coefficient of the wind turbine.

Evaluation of range and distortion tolerance for high Mach number transonic fan stages. Task 2: Performance of a 1500-foot-per-second tip speed transonic fan stage with variable geometry inlet guide vanes and stator

NASA Technical Reports Server (NTRS)

Bilwakesh, K. R.; Koch, C. C.; Prince, D. C.

1972-01-01

A 0.5 hub/tip radius ratio compressor stage consisting of a 1500 ft/sec tip speed rotor, a variable camber inlet guide vane and a variable stagger stator was designed and tested with undistorted inlet flow, flow with tip radial distortion, and flow with 90 degrees, one-per-rev, circumferential distortion. At the design speed and design IGV and stator setting the design stage pressure ratio was achieved at a weight within 1% of the design flow. Analytical results on rotor tip shock structure, deviation angle and part-span shroud losses at different operating conditions are presented. The variable geometry blading enabled efficient operation with adequate stall margin at the design condition and at 70% speed. Closing the inlet guide vanes to 40 degrees changed the speed-versus-weight flow relationship along the stall line and thus provided the flexibility of operation at off-design conditions. Inlet flow distortion caused considerable losses in peak efficiency, efficiency on a constant throttle line through design pressure ratio at design speed, stall pressure ratio, and stall margin at the 0 degrees IGV setting and high rotative speeds. The use of the 40 degrees inlet guide vane setting enabled partial recovery of the stall margin over the standard constant throttle line.

Unsteady Subsonic and Transonic Potential Flow over Helicopter Rotor Blades

NASA Technical Reports Server (NTRS)

Isom, M. P.

1974-01-01

Differential equations and boundary conditions for a rotor blade in forward flight, with subsonic or transonic tip Mach number, are derived. A variety of limiting flow regimes determined by different limits involving blade thickness ratio, aspect ratio, advance ratio and maximum tip Mach number is discussed. The transonic problem is discussed in some detail, and in particular the conditions that make this problem quasi-steady or essentially unsteady are determined. Asymptotic forms of equations and boundary conditions that are valid in an appropriately scaled region of the tip and an azimuthal sector on the advancing side are derived. The equations are then put in a form that is valid from the blade tip inboard through the strip theory region.

Numerical investigation for design and critical performance evaluation of a horizontal axis hydrokinetic turbine

NASA Astrophysics Data System (ADS)

Subhra Mukherji, Suchi; Banerjee, Arindam

2010-11-01

We will discuss findings from our numerical investigation on the hydrodynamic performance of horizontal axis hydrokinetic turbines (HAHkT) under different turbine geometries and flow conditions. Hydrokinetic turbines are a class of zero-head hydropower systems which utilizes kinetic energy of flowing water to drive a generator. However, such turbines very often suffer from low efficiency which is primarily controlled by tip-speed ratio, solidity, angle of attack and number of blades. A detailed CFD study was performed using two-dimensional and three dimensional numerical models to examine the effect of each of these parameters on the performance of small HAHkTs having power capacities <= 10 kW. The two-dimensional numerical results provide an optimum angle of attack that maximizes the lift as well as lift to drag ratio yielding maximum power output. However three-dimensional numerical studies estimate optimum turbine solidity and blade numbers that produces maximum power coefficient at a given tip speed ratio. In addition, simulations were also performed to observe the axial velocity deficit at the turbine rotor downstream for different tip-speed ratios to obtain both qualitative and quantitative details about stall delay phenomena and the energy loss suffered by the turbine under ambient flow condition.

Localized, Non-Harmonic Active Flap Motions for Low Frequency In-Plane Rotor Noise Reduction

DTIC Science & Technology

2012-05-01

rotating -frame cyclic variations, of two- per-rev or greater, to augment blade motions and blade airloads. Recent studies (Refs. 5, 6) have...Advancing tip Mach number MH Rotational (Hover) tip Mach number NM Noise metric, peak-to-peak value R Blade radius α...from a full-scale, 2,900 lb. gross weight, four-bladed S-434TM helicopter. The rotor head, blade cuffs , and swash-plate were production S

Tip Clearance Control Using Plasma Actuators

DTIC Science & Technology

2007-03-01

Clearance Control Using Plasma Actuators 4 posed by Denton (1993). A number of investigators have used partial shrouds, or " winglet " designs to...main molded blade with a span of 3.42 in., a removable molded blade segment with a span of 0.1875 in., and removable blade tip winglets made of glass...segment and the main blade to vary the distance between the blade end and the front wall of the cascade section. The winglets were machined using a

Successful Solutions to SSME/AT Development Turbine Blade Distress

NASA Technical Reports Server (NTRS)

Montgomery, Stuart K.

1999-01-01

As part of the High-Pressure Fuel Turbopump/Alternate Turbopump (HPFTP/AT) turbine blade development program, unique turbine blade design features were implemented to address 2nd stage turbine blade high cycle fatigue distress and improve turbine robustness. Features included the addition of platform featherseal dampers, asymmetric blade tip seal segments, gold plating of the blade attachments, and airfoil tip trailing edge modifications. Development testing shows these features have eliminated turbine blade high cycle fatigue distress and consequently these features are currently planned for incorporation to the flight configuration. Certification testing will begin in 1999. This presentation summarizes these features.

Comparison of calculated and measured velocities near the tip of a model rotor blade at transonic speeds

NASA Technical Reports Server (NTRS)

Tauber, M. E.; Owen, F. K.; Langhi, R. G.; Palmer, G. E.

1985-01-01

The ability of the ROT22 code to predict accurately the transonic flow field in the crucial region around and beyond the tip of a high speed rotor blade was assessed. The computations were compared with extensive laser velocimetry measurements made at zero advance ratio and tip Mach numbers of 0.85, 0.88, 0.90, and 0.95. The comparison between theory and experiment was made using 300 scans for the three orthogonal velocity components covering a volume having a height of over one blade chord, a width of nearly two chords, and a length ranging from about 1 to 1.6 chords, depending on the tip speeds. The good agreement between the calculated and measured velocities established the ability of the code to predict the off blade flow field at high tip speeds. This supplements previous comparisons where surface pressures were shown to be well predicted on two different tips at advance ratios to 0.45, especially at the critical 90 deg azimuth blade position. These results demonstrate that the ROT22 code can be used with confidence to predict the important tip region flow field including the occurrence, strength, and location of shock waves causing high drag and noise.

Shape design and CFD analysis on a 1MW-class horizontal axis tidal current turbine blade

NASA Astrophysics Data System (ADS)

Singh, P. M.; Choi, Y. D.

2013-12-01

This study aims to develop a 1MW-class horizontal axis tidal current turbine rotor blade which can be applied near the southwest island regions of South Korea. On the basis of actual tidal current conditions of southern region of Korea, configuration design of 1MW class turbine rotor blade is carried out by BEMT (Blade element momentum theory). The hydrodynamic performance including the lift and drag forces, is conducted with the variation of the angle of attack using an open source code of X-Foil. The purpose of the study is to study the shape of the hydrofoil used and how it affects the performance of the turbine. After a thorough study of many airfoils, a new hydrofoil is developed using the S814 and DU-91-W2- 250 airfoils, which show good performance for rough conditions. A combination of the upper and lower surface of the two hydrofoils is tested. Three dimensional models were developed and the optimized blade geometry is used for CFD (Computational Fluid Dynamics) analysis with hexahedral numerical grids. Power coefficient, pressure coefficient and velocity distributions are investigated according to Tip Speed Ratio by CFD analysis.

Tip cap for a rotor blade

NASA Technical Reports Server (NTRS)

Kofel, W. K.; Tuley, E. N.; Gay, C. H., Jr.; Troeger, R. E.; Sterman, A. P. (Inventor)

1983-01-01

A replaceable tip cap for attachment to the end of a rotor blade is described. The tip cap includes a plurality of walls defining a compartment which, if desired, can be divided into a plurality of subcompartments. The tip cap can include inlet and outlet holes in walls thereof to permit fluid communication of a cooling fluid there through. Abrasive material can be attached with the radially outer wall of the tip cap.

Users Guide for Smooth-Prop: A Program for Smoothing Propeller Tip Geometry

DTIC Science & Technology

2013-10-01

Research and Development Canada – Atlantic Technical Memorandum DRDC Atlantic TM 2013-179 October 2013 c© Her Majesty the Queen in Right of Canada...a window You can magnify any region of the Blade or Plane Intersection windows by clicking and dragging with the left mouse button. A red rectangle...display. Each slider is a long rectangle containing a smaller black box: see Fig. 9. The black box is a handle that can be dragged back and forth

A method to estimate weight and dimensions of large and small gas turbine engines

NASA Technical Reports Server (NTRS)

Onat, E.; Klees, G. W.

1979-01-01

A computerized method was developed to estimate weight and envelope dimensions of large and small gas turbine engines within + or - 5% to 10%. The method is based on correlations of component weight and design features of 29 data base engines. Rotating components were estimated by a preliminary design procedure which is sensitive to blade geometry, operating conditions, material properties, shaft speed, hub tip ratio, etc. The development and justification of the method selected, and the various methods of analysis are discussed.

Turbine blade tip clearance measurement using a skewed dual-beam fiber optic sensor

NASA Astrophysics Data System (ADS)

Ye, De-chao; Duan, Fa-jie; Guo, Hao-tian; Li, Yangzong; Wang, Kai

2012-08-01

Optimization and active control of the tip clearance of turbine blades has been identified as a key to improve fuel efficiency, reduce emission, and increase service life of the engine. However, reliable and real-time tip clearance measurement is difficult due to the adverse environmental conditions that are typically found in a turbine. We describe a dual-beam fiber optic measurement system that can measure the tip timing and tip clearance simultaneously. Because the tip timing information is used to calculate the tip clearance, the method is insensitive to the signal intensity variation caused by fluctuations in environmental conditions such as light source instability, contamination, and blade tip imperfection. The system was calibrated and tested using experimental rotors. The test results indicated a high resolution of 4.5 μm and measurement accuracy of ±20 μm over the rotation speed range of 2000 to 10,000 rpm.

Integrated circuit cooled turbine blade

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lee, Ching-Pang; Jiang, Nan; Um, Jae Y.

A turbine rotor blade includes at least two integrated cooling circuits that are formed within the blade that include a leading edge circuit having a first cavity and a second cavity and a trailing edge circuit that includes at least a third cavity located aft of the second cavity. The trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity. The last cavity is located along a trailing edge of the blade. A tip axial cooling channelmore » connects to the first cavity of the leading edge circuit and the penultimate cavity of the trailing edge circuit. At least one crossover hole connects the penultimate cavity to the last cavity substantially near the tip end of the blade.« less

Integrated Turbine Tip Clearance and Gas Turbine Engine Simulation

NASA Technical Reports Server (NTRS)

Chapman, Jeffryes W.; Kratz, Jonathan; Guo, Ten-Huei; Litt, Jonathan

2016-01-01

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.

On the inverse problem of blade design for centrifugal pumps and fans

NASA Astrophysics Data System (ADS)

Kruyt, N. P.; Westra, R. W.

2014-06-01

The inverse problem of blade design for centrifugal pumps and fans has been studied. The solution to this problem provides the geometry of rotor blades that realize specified performance characteristics, together with the corresponding flow field. Here a three-dimensional solution method is described in which the so-called meridional geometry is fixed and the distribution of the azimuthal angle at the three-dimensional blade surface is determined for blades of infinitesimal thickness. The developed formulation is based on potential-flow theory. Besides the blade impermeability condition at the pressure and suction side of the blades, an additional boundary condition at the blade surface is required in order to fix the unknown blade geometry. For this purpose the mean-swirl distribution is employed. The iterative numerical method is based on a three-dimensional finite element method approach in which the flow equations are solved on the domain determined by the latest estimate of the blade geometry, with the mean-swirl distribution boundary condition at the blade surface being enforced. The blade impermeability boundary condition is then used to find an improved estimate of the blade geometry. The robustness of the method is increased by specific techniques, such as spanwise-coupled solution of the discretized impermeability condition and the use of under-relaxation in adjusting the estimates of the blade geometry. Various examples are shown that demonstrate the effectiveness and robustness of the method in finding a solution for the blade geometry of different types of centrifugal pumps and fans. The influence of the employed mean-swirl distribution on the performance characteristics is also investigated.

Blade-to-Blade Variations in Shocks Upstream of Both a Forward-Swept and an Aft-Swept Fan

NASA Technical Reports Server (NTRS)

Podboy, Gary G.; Krupar, Martin J.

2006-01-01

Detailed laser Doppler velocimeter (LDV) flow field measurements were made upstream of 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 supersonic tip speeds. The blade-to-blade variations in the flows associated with these shocks are thought to be responsible for generating Multiple Pure Tone (MPT) noise. The measured blade-to-blade variations are documented in this report through a series of slideshows which show relative Mach number contours computed from the velocity measurements. Data are presented for the forward-swept fan operating at three speeds (corresponding to tip relative Mach numbers of 0.817, 1.074, and 1.189), and for the aft-swept fan operating at two (tip relative Mach numbers of 1.074 and 1.189). These LDV data illustrate how the perturbations in the upstream flow field created by the rotating blades vary with axial position, radial position and rotor speed. As expected, at the highest tested speed the forward-swept fan swallowed the shocks which occur in the tip region, whereas the aftswept fan did not. This resulted in a much smaller flow disturbance just upstream of the tip of the forward-swept fan. Nevertheless, further upstream the two fan flows were much more similar.

Rotor tip clearance effects on overall and blade-element performance of axial-flow transonic fan stage

NASA Technical Reports Server (NTRS)

Moore, R. D.

1982-01-01

The effects of tip clearance on the overall and blade-element performance of an axial-flow transonic fan stage are presented. The 50-centimeter-diameter fan was tested at four tip clearances (nonrotating) from 0.061 to 0.178 centimeter. The calculated radial growth of the blades was 0.040 centimeter at design conditions. The decrease in overall stage performance with increasing clearance is attributed to the loss in rotor performance. For the rotor the effects of clearance on performance parameters extended to about 70 percent of blade span from the tip. The stage still margin based on an assumed operating line decreased from 15.3 to 0 percent as the clearance increased from 0.061 to 0.178 centimeter.

Aeroelasticity and structural optimization of composite helicopter rotor blades with swept tips

NASA Technical Reports Server (NTRS)

Yuan, K. A.; Friedmann, P. P.

1995-01-01

This report describes the development of an aeroelastic analysis capability for composite helicopter rotor blades with straight and swept tips, and its application to the simulation of helicopter vibration reduction through structural optimization. A new aeroelastic model is developed in this study which is suitable for composite rotor blades with swept tips in hover and in forward flight. The hingeless blade is modeled by beam type finite elements. A single finite element is used to model the swept tip. Arbitrary cross-sectional shape, generally anisotropic material behavior, transverse shears and out-of-plane warping are included in the blade model. The nonlinear equations of motion, derived using Hamilton's principle, are based on a moderate deflection theory. Composite blade cross-sectbnal properties are calculated by a separate linear, two-dimensional cross section analysis. The aerodynamic loads are obtained from quasi-steady, incompressible aerodynamics, based on an implicit formulation. The trim and steady state blade aeroelastic response are solved in a fully coupled manner. In forward flight, where the blade equations of motion are periodic, the coupled trim-aeroelastic response solution is obtained from the harmonic balance method. Subsequently, the periodic system is linearized about the steady state response, and its stability is determined from Floquet theory.

Experimental performance and acoustic investigation of modern, counterrotating blade concepts

NASA Technical Reports Server (NTRS)

Hoff, G. E.

1990-01-01

The aerodynamic, acoustic, and aeromechanical performance of counterrotating blade concepts were evaluated both theoretically and experimentally. Analytical methods development and design are addressed. Utilizing the analytical methods which evolved during the conduct of this work, aerodynamic and aeroacoustic predictions were developed, which were compared to NASA and GE wind tunnel test results. The detailed mechanical design and fabrication of five different composite shell/titanium spar counterrotating blade set configurations are presented. Design philosophy, analyses methods, and material geometry are addressed, as well as the influence of aerodynamics, aeromechanics, and aeroacoustics on the design procedures. Blade fabrication and quality control procedures are detailed; bench testing procedures and results of blade integrity verification are presented; and instrumentation associated with the bench testing also is identified. Additional hardware to support specialized testing is described, as are operating blade instrumentation and the associated stress limits. The five counterrotating blade concepts were scaled to a tip diameter of 2 feet, so they could be incorporated into MPS (model propulsion simulators). Aerodynamic and aeroacoustic performance testing was conducted in the NASA Lewis 8 x 6 supersonic and 9 x 15 V/STOL (vertical or short takeoff and landing) wind tunnels and in the GE freejet anechoic test chamber (Cell 41) to generate an experimental data base for these counterrotating blade designs. Test facility and MPS vehicle matrices are provided, and test procedures are presented. Effects on performance of rotor-to-rotor spacing, angle-of-attack, pylon proximity, blade number, reduced-diameter aft blades, and mismatched rotor speeds are addressed. Counterrotating blade and specialized aeromechanical hub stability test results are also furnished.

Integrated fiber optic light probe: Measurement of static deflections in rotating turbomachinery

NASA Astrophysics Data System (ADS)

Dhadwal, Harbans S.; Mehmud, Ali; Khan, Romel; Kurkov, Anatole

1996-02-01

This paper describes the design, fabrication, and testing of an integrated fiber optic light probe system for monitoring blade tip deflections, vibrational modes, and changes in blade tip clearances in the compressor stage of rotating turbomachinery. The system comprises a set of integrated fiber optic light probes which are positioned to detect the passing blade tip at the leading and the trailing edges. In this configuration measurements of both blade vibrations and steady-state blade deflection can be obtained from the timing information provided by each light probe, which comprises an integrated fiber optic transmitting channel and a number of high numerical aperture receiving fibers, all mounted in the same cylindrical housing. A spatial resolution of 50 μm is possible with the integrated fiber optic technology, while keeping the outer diameter below 2.5 mm. Additionally, one fiber sensor provides a capability of monitoring changes in the blade tip clearance of the order of 10 μm. Measurements from a single stage compressor facility and an engine-fan rig in a 9 ft×15 ft subsonic wind tunnel are presented.

Comprehensive Structural Dynamic Analysis of the SSME/AT Fuel Pump First-Stage Turbine Blade

NASA Technical Reports Server (NTRS)

Brown, A. M.

1998-01-01

A detailed structural dynamic analysis of the Pratt & Whitney high-pressure fuel pump first-stage turbine blades has been performed to identify the cause of the tip cracking found in the turbomachinery in November 1997. The analysis was also used to help evaluate potential fixes for the problem. Many of the methods available in structural dynamics were applied, including modal displacement and stress analysis, frequency and transient response to tip loading from the first-stage Blade Outer Gas Seals (BOGS), fourier analysis, and shock spectra analysis of the transient response. The primary findings were that the BOGS tip loading is impulsive in nature, thereby exciting many modes of the blade that exhibit high stress at the tip cracking location. Therefore, a proposed BOGS count change would not help the situation because a clearly identifiable resonance situation does not exist. The recommendations for the resolution of the problem are to maintain the existing BOGS count, eliminate the stress concentration in the blade due to its geometric design, and reduce the applied load on the blade by adding shiplaps in the BOGS.

An Investigation into the Aerodynamics Surrounding Vertical-Axis Wind Turbines

NASA Astrophysics Data System (ADS)

Parker, Colin M.

The flow surrounding a scaled model vertical-axis wind turbine (VAWT) at realistic operating conditions was studied. The model closely matches geometric and dynamic properties--tip-speed ratio and Reynolds number--of a full-size turbine. The flowfield is measured using particle imaging velocimetry (PIV) in the mid-plane upstream, around, and after (up to 4 turbine diameters downstream) the turbine, as well as a vertical plane behind the turbine. Ensemble-averaged results revealed an asymmetric wake behind the turbine, regardless of tip-speed ratio, with a larger velocity deficit for a higher tip-speed ratio. For the higher tip-speed ratio, an area of averaged flow reversal is present with a maximum reverse flow of -0.04Uinfinity. Phase-averaged vorticity fields--achieved by syncing the PIV system with the rotation of the turbine--show distinct structures form from each turbine blade. There are distinct differences in the structures that are shed into the wake for tip-speed ratios of 0.9, 1.3 and 2.2--switching from two pairs to a single pair of shed vortices--and how they convect into the wake--the middle tip-speed ratio vortices convect downstream inside the wake, while the high tip-speed ratio pair is shed into the shear layer of the wake. The wake structure is found to be much more sensitive to changes in tip-speed ratio than to changes in Reynolds number. The geometry of a turbine can influence tip-speed ratio, but the precise relationship among VAWT geometric parameters and VAWT wake characteristics remains unknown. Next, we characterize the wakes of three VAWTs that are geometrically similar except for the ratio of the turbine diameter (D), to blade chord (c), which was chosen to be D/c = 3, 6, and 9, for a fixed freestream Reynolds number based on the blade chord of Rec =16,000. In addition to two-component PIV and single-component constant temperature anemometer measurements are made at the horizontal mid-plane in the wake of each turbine. Hot-wire measurement locations are selected to coincide with the edge of the shear layer of each turbine wake, as deduced from the PIV data, which allows for an analysis of the frequency content of the wake due to vortex shedding by the turbine. Changing the tip-speed ratio leads to substantial wake variation possibly because changing the tip-speed ratio changes the dynamic solidity. In this work, we achieve a similar change in dynamic solidity by varying the D/c ratio and holding the tip-speed ratio constant. This change leads to very similar characteristic shifts in the wake, such as a greater blockage effect, including averaged flow reversal in the case of high dynamic solidity (D/c = 3). The phase-averaged vortex identification shows that both the blockage effect and the wake structures are similarly affected by a change in dynamic solidity. At lower dynamic solidity, pairs of vortices are shed into the wake directly downstream of the turbine. For all three models, a vortex chain is shed into the shear layer at the edge of the wake where the blade is processing into the freestream.

Investigation of rotor blade tip-vortex aerodynamics

NASA Technical Reports Server (NTRS)

Lewellen, W. S.

1971-01-01

Several aspects of the aerodynamics of rotor blade tip vortices are examined. Two particular categories are dealt with; (1) dynamic loads on a blade passing close to or intersecting a trailing vortex, and (2) the response of the trailing vortex core to changes in the flow. Results for both categories are in reasonable agreement with existing data, although lower pressure gradients were obtained than anticipated for category one. A correlation between trailing edge sweep angle at the tip and vortex core size was noted for category two.

Test Rig for Active Turbine Blade Tip Clearance Control Concepts: An Update

NASA Technical Reports Server (NTRS)

Taylor, Shawn; Steinetz, Bruce; Oswald, Jay; DeCastro, Jonathan; Melcher, Kevin

2006-01-01

The objective is to develop and demonstrate a fast-acting active clearance control system to improve turbine engine performance, reduce emissions, and increase service life. System studies have shown the benefits of reducing blade tip clearances in modern turbine engines. Minimizing blade tip clearances throughout the engine will contribute materially to meeting NASA's Ultra-Efficient Engine Technology (UEET) turbine engine project goals. NASA GRC is examining two candidate approaches including rub-avoidance and regeneration which are explained in subsequent slides.

Wind blade spar cap and method of making

DOEpatents

Mohamed, Mansour H [Raleigh, NC

2008-05-27

A wind blade spar cap for strengthening a wind blade including an integral, unitary three-dimensional woven material having a first end and a second end, corresponding to a root end of the blade and a tip end of the blade, wherein the material tapers in width from the first to the second end while maintaining a constant thickness and decreasing weight therebetween, the cap being capable of being affixed to the blade for providing increased strength with controlled variation in weight from the root end to the tip end based upon the tapered width of the material thereof. The present inventions also include the method of making the wind blade spar cap and a wind blade including the wind blade spar cap.

An asymptotic theory of supersonic propeller noise

NASA Technical Reports Server (NTRS)

Envia, Edmane

1992-01-01

A theory for predicting the noise field of a propeller with a realistic blade geometry is presented. The theory, which utilizes a large blade count approximation, provides an efficient formula for predicting the radiation of sound from all three sources of propeller noise. Comparisons with full numerical integration indicate that the noise levels predicted by this formula are quite accurate. Calculations based on this method also show that the radiation from the Lighthill quadrupole source is rather substantial when compared with thickness and loading noise for high speed propellers. A preliminary application of the theory to the problem of the sensitivity of the peak noise levels generated by a supersonic propeller to the variations in its tip helical Mach number has produced a trend that is in qualitative agreement with the experimental observations.

Numerical investigation of influence of tip leakage flow on secondary flow in transonic centrifugal compressor at design condition

NASA Astrophysics Data System (ADS)

Kaneko, Masanao; Tsujita, Hoshio

2015-04-01

In a centrifugal compressor, the leakage flow through the tip clearance generates the tip leakage vortex by the interaction with the main flow, and consequently makes the flow in the impeller passage more complex by the interaction with the passage vortex. In addition, the tip leakage vortex interacts with the shock wave on the suction surface near the blade tip in the transonic centrifugal compressor impeller. Therefore, the detailed examination for the influence of the tip leakage vortex becomes seriously important to improve the aerodynamic performance especially for the transonic centrifugal compressor. In this study, the flows in the transonic centrifugal compressor with and without the tip clearance at the design condition were analyzed numerically by using the commercial CFD code. The computed results revealed that the tip leakage vortex induced by the high loading at the blade tip around the leading edge affected the loss generation by the reduction or the suppression of the shock wave on the suction surface of the blade.

Energy dissipation in the blade tip region of an axial fan

NASA Astrophysics Data System (ADS)

Bizjan, B.; Milavec, M.; Širok, B.; Trenc, F.; Hočevar, M.

2016-11-01

A study of velocity and pressure fluctuations in the tip clearance flow of an axial fan is presented in this paper. Two different rotor blade tip designs were investigated: the standard one with straight blade tips and the modified one with swept-back tip winglets. Comparison of integral sound parameters indicates a significant noise level reduction for the modified blade tip design. To study the underlying mechanisms of the energy conversion and noise generation, a novel experimental method based on simultaneous measurements of local flow velocity and pressure has also been developed and is presented here. The method is based on the phase space analysis by the use of attractors, which enable more accurate identification and determination of the local flow structures and turbulent flow properties. Specific gap flow energy derived from the pressure and velocity time series was introduced as an additional attractor parameter to assess the flow energy distribution and dissipation within the phase space, and thus determines characteristic sources of the fan acoustic emission. The attractors reveal a more efficient conversion of the pressure to kinetic flow energy in the case of the modified (tip winglet) fan blade design, and also a reduction in emitted noise levels. The findings of the attractor analysis are in a good agreement with integral fan characteristics (efficiency and noise level), while offering a much more accurate and detailed representation of gap flow phenomena.

Fundamental Characterization of Spanwise Loading and Trailed Wake Vortices

DTIC Science & Technology

2016-07-01

the close interaction of the tip vortex with a following blade . Such vortex interactions are fundamental determinants of rotor performance, loads, and...wing loading distribution differs from a typical loading on a hovering rotor blade in that the maximum bound circulation occurs at the blade root...and not close to the tip; this is similar to a very highly twisted rotor blade , like a tilt-rotor, in hover. The wing-vortex interaction alters the

Full-Scale Wind-Tunnel Tests of a PCA-2 Autogiro Rotor

NASA Technical Reports Server (NTRS)

Wheatley, John B; Hood, Manley J

1935-01-01

This report presents the results of force tests on and air-flow surveys near PCA-2 autogiro rotor in the NACA full-scale wind tunnel. The force tests were made at three pitch settings and several rotor speeds; the effect of fairing protuberances on the rotor blade was determined. Induced downwash and yaw angles were determined at low tip-speed ratios in a plane 1 1/2 feet above the path of the blade tips. The results show that the maximum l/d of the rotor cannot be appreciably increased by increasing the blade pitch angle above about 4.5 degrees at the blade tip; that the protuberances on the blades cause more than 5 percent of the total rotor drag; and that the rotor center-of-pressure travel is very small.

Reasons for low aerodynamic performance of 13.5-centimeter-tip-diameter aircraft engine starter turbine

NASA Technical Reports Server (NTRS)

Haas, J. E.; Roelke, R. J.; Hermann, P.

1981-01-01

The reasons for the low aerodynamic performance of a 13.5 cm tip diameter aircraft engine starter turbine were investigated. Both the stator and the stage were evaluated. Approximately 10 percent improvement in turbine efficiency was obtained when the honeycomb shroud over the rotor blade tips was filled to obtain a solid shroud surface. Efficiency improvements were obtained for three rotor configurations when the shroud was filled. It is suggested that the large loss associated with the open honeycomb shroud is due primarily to energy loss associated with gas transportation as a result of the blade to blade pressure differential at the tip section.

Investigation of Unsteady Tip Clearance Flow in a Low-Speed One and Half Stage Axial Compressor with LES And PIV

NASA Technical Reports Server (NTRS)

Hah, Chunill; Hathaway, Michael; Katz, Joseph; Tan, David

2015-01-01

The primary focus of this paper is to investigate how a rotor's unsteady tip clearance flow structure changes in a low speed one and half stage axial compressor when the rotor tip gap size is increased from 0.5 mm (0.49% of rotor tip blade chord, 2% of blade span) to 2.4 mm (2.34% chord, 4% span) at the design condition are investigated. The changes in unsteady tip clearance flow with the 0.62 % tip gap as the flow rate is reduced to near stall condition are also investigated. A Large Eddy Simulation (LES) is applied to calculate the unsteady flow field at these three flow conditions. Detailed Stereoscopic PIV (SPIV) measurements of the current flow fields were also performed at the Johns Hopkins University in a refractive index-matched test facility which renders the compressor blades and casing optically transparent. With this setup, the unsteady velocity field in the entire flow domain, including the flow inside the tip gap, can be measured. Unsteady tip clearance flow fields from LES are compared with the PIV measurements and both LES and PIV results are used to study changes in tip clearance flow structures. The current study shows that the tip clearance vortex is not a single structure as traditionally perceived. The tip clearance vortex is formed by multiple interlaced vorticities. Therefore, the tip clearance vortex is inherently unsteady. The multiple interlaced vortices never roll up to form a single structure. When phased-averaged, the tip clearance vortex appears as a single structure. When flow rate is reduced with the same tip gap, the tip clearance vortex rolls further upstream and the tip clearance vortex moves further radially inward and away from the suction side of the blade. When the tip gap size is increased at the design flow condition, the overall tip clearance vortex becomes stronger and it stays closer to the blade suction side and the vortex core extends all the way to the exit of the blade passage. Measured and calculated unsteady flow fields inside the tip gap agree fairly well. Instantaneous velocity vectors inside the tip gap from both the PIV and LES do show flow separation and reattachment at the entrance of tip gap as some earlier studies suggested. This area at the entrance of tip gap flow (the pressure side of the blade) is confined very close to the rotor tip section. With a small tip gap (0.5mm), the gap flow looks like a simple two-dimensional channel flow with larger velocity near the casing for both flow rates. A small area with a sharp velocity gradient is observed just above the rotor tip. This strong shear layer is turned radially inward when it collides with the incoming flow and forms the core structure of the tip clearance vortex. When tip gap size is increased to 2.4 mm at the design operation, the radial profile of the tip gap flow changes drastically. With the large tip gap, the gap flow looks like a two-dimensional channel flow only near the casing. Near the rotor top section, a bigger region with very large shear and reversed flow is observed.

Study of casing treatment stall margin improvement phenomena. [for compressor rotor blade tips compressor blades rotating stalls

NASA Technical Reports Server (NTRS)

Prince, D. C., Jr.; Wisler, D. C.; Hilvers, D. E.

1974-01-01

The results of a program of experimental and analytical research in casing treatments over axial compressor rotor blade tips are presented. Circumferential groove, axial-skewed slot, and blade angle slot treatments were tested. These yielded, for reduction in stalling flow and loss in peak efficiency, 5.8% and 0 points, 15.3% and 2.0 points, and 15.0% and 1.2 points, respectively. These values are consistent with other experience. The favorable stalling flow situations correlated well with observations of higher-then-normal surface pressures on the rotor blade pressure surfaces in the tip region, and with increased maximum diffusions on the suction surfaces. Annular wall pressure gradients, especially in the 50-75% chord region, are also increased and blade surface pressure loadings are shifted toward the trailing edge for treated configurations. Rotor blade wakes may be somewhat thinner in the presence of good treatments, particularly under operating conditions close to the baseline stall.

System and method for improved rotor tip performance

NASA Technical Reports Server (NTRS)

Zientek, Thomas A. (Inventor); Bussom, Richard (Inventor); McVeigh, Michael A. (Inventor); Narducci, Robert P. (Inventor)

2007-01-01

The present invention discloses systems and methods for the performance enhancement of rotary wing aircraft through reduced torque, noise and vibration. In one embodiment, a system includes a sail having an aerodynamic shape positioned proximate to a tip of the rotor blade. An actuator may be configured to rotate the sail relative to the blade tip. a A control system receives information from a rotorcraft system and commands the actuator to rotate the sail to a predetermined favorable rotor blade operating condition. In another embodiment, a method includes configuring the rotorcraft in a selected flight condition, communicating input signals to a control system operable to position sails coupled to tips of blades of a rotor assembly, processing the input signals according to a constraint condition to generate sail positional information, and transferring the sail positional information to the sail.

Turbine-blade tip clearance and tip timing measurements using an optical fiber bundle sensor

NASA Astrophysics Data System (ADS)

Garcia, Iker; Beloki, Josu; Zubia, Joseba; Durana, Gaizka; Aldabaldetreku, Gotzon

2013-04-01

Traditional limitations of capacitive, inductive or discharging probe sensor for tip timing and tip clearance measurements are overcome by reflective intensity modulated optical fiber sensors. This paper presents the signals and results corresponding to a one stage turbine rig which rotor has 146 blades, obtained from a transonic wind-tunnel test. The probe is based on a trifurcated bundle of optical fibers that is mounted on turbine casing. It is composed of a central illuminating fiber that guides the light from a laser to the turbine blade, and two concentric rings of receiving fibers that collect the reflected light. Two photodetectors turn this reflected light signal from the receiving rings into voltage. The electrical signals are acquired and saved by a high-sample-rate oscilloscope. In tip clearance calculations the ratio of the signals provided by each ring of receiving fibers is evaluated and translated into distance. In the case of tip timing measurements, only one of the signals is considered to get the arrival time of the blade. The differences between the real and theoretical arrival times of the blades are used to obtain the deflections amplitude. The system provides the travelling wave spectrum, which presents the average vibration amplitude of the blades at a certain nodal diameter. The reliability of the results in the turbine rig testing facilities suggests the possibility of performing these measurements in real turbines under real working conditions.

Advanced ceramic material for high temperature turbine tip seals

NASA Technical Reports Server (NTRS)

Solomon, N. G.; Vogan, J. W.

1978-01-01

Ceramic material systems are being considered for potential use as turbine blade tip gas path seals at temperatures up to 1370 1/4 C. Silicon carbide and silicon nitride structures were selected for study since an initial analysis of the problem gave these materials the greatest potential for development into a successful materials system. Segments of silicon nitride and silicon carbide materials over a range of densities, processed by various methods, a honeycomb structure of silicon nitride and ceramic blade tip inserts fabricated from both materials by hot pressing were tested singly and in combination. The evaluations included wear under simulated engine blade tip rub conditions, thermal stability, impact resistance, machinability, hot gas erosion and feasibility of fabrication into engine components. The silicon nitride honeycomb and low-density silicon carbide using a selected grain size distribution gave the most promising results as rub-tolerant shroud liners. Ceramic blade tip inserts made from hot-pressed silicon nitride gave excellent test results. Their behavior closely simulated metal tips. Wear was similar to that of metals but reduced by a factor of six.

Enhanced Actuator Line Simulation of a Wind Turbine by including the Conservative Load at the Blade Tip

NASA Astrophysics Data System (ADS)

Herraez, Ivan; Micallef, Daniel; van Kuik, Gijs A. M.; Peinke, Joachim

2015-11-01

At the tip of wind turbine blades, the radial bound circulation is transformed into chordwise circulation just before being released as trailing vorticity, giving rise to the tip vortex. The force acting on the chordwise circulation contains a radial and a normal component with respect to the blade axis. This load does not contribute to the torque, so it is a conservative load. Due to this, it is disregarded in the engineering tools used for the design of wind turbines. However, as we demonstrated in a previous work, the conservative load might influence the trajectory of the tip vortex. In order to see how this affects the blade loads, in this research we perform large eddy simulations with an actuator line model where the conservative load has been included. The conservative load reduces the angle of attack in the tip region as a consequence of the modified tip vortex trajectory. This has a negative influence on the lift and the power output. We conclude that the accuracy of engineering design tools of wind turbines can be improved if the conservative load acting at the tip is considered.

Structural Health Monitoring on Turbine Engines Using Microwave Blade Tip Clearance Sensors

NASA Technical Reports Server (NTRS)

Woike, Mark; Abdul-Aziz, Ali; Clem, Michelle

2014-01-01

The ability to monitor the structural health of the rotating components, especially in the hot sections of turbine engines, is of major interest to aero community in improving engine safety and reliability. The use of instrumentation for these applications remains very challenging. It requires sensors and techniques that are highly accurate, are able to operate in a high temperature environment, and can detect minute changes and hidden flaws before catastrophic events occur. The National Aeronautics and Space Administration (NASA) has taken a lead role in the investigation of new sensor technologies and techniques for the in situ structural health monitoring of gas turbine engines. As part of this effort, microwave sensor technology has been investigated as a means of making high temperature non-contact blade tip clearance, blade tip timing, and blade vibration measurements for use in gas turbine engines. This paper presents a summary of key results and findings obtained from the evaluation of two different types of microwave sensors that have been investigated for use possible in structural health monitoring applications. The first is a microwave blade tip clearance sensor that has been evaluated on a large scale Axial Vane Fan, a subscale Turbofan, and more recently on sub-scale turbine engine like disks. The second is a novel microwave based blade vibration sensor that was also used in parallel with the microwave blade tip clearance sensors on the experiments with the sub-scale turbine engine disks.

Blade tip timing (BTT) uncertainties

NASA Astrophysics Data System (ADS)

Russhard, Pete

2016-06-01

Blade Tip Timing (BTT) is an alternative technique for characterising blade vibration in which non-contact timing probes (e.g. capacitance or optical probes), typically mounted on the engine casing (figure 1), and are used to measure the time at which a blade passes each probe. This time is compared with the time at which the blade would have passed the probe if it had been undergoing no vibration. For a number of years the aerospace industry has been sponsoring research into Blade Tip Timing technologies that have been developed as tools to obtain rotor blade tip deflections. These have been successful in demonstrating the potential of the technology, but rarely produced quantitative data, along with a demonstration of a traceable value for measurement uncertainty. BTT technologies have been developed under a cloak of secrecy by the gas turbine OEM's due to the competitive advantages it offered if it could be shown to work. BTT measurements are sensitive to many variables and there is a need to quantify the measurement uncertainty of the complete technology and to define a set of guidelines as to how BTT should be applied to different vehicles. The data shown in figure 2 was developed from US government sponsored program that bought together four different tip timing system and a gas turbine engine test. Comparisons showed that they were just capable of obtaining measurement within a +/-25% uncertainty band when compared to strain gauges even when using the same input data sets.

Structural health monitoring on turbine engines using microwave blade tip clearance sensors

NASA Astrophysics Data System (ADS)

Woike, Mark; Abdul-Aziz, Ali; Clem, Michelle

2014-04-01

The ability to monitor the structural health of the rotating components, especially in the hot sections of turbine engines, is of major interest to the aero community in improving engine safety and reliability. The use of instrumentation for these applications remains very challenging. It requires sensors and techniques that are highly accurate, are able to operate in a high temperature environment, and can detect minute changes and hidden flaws before catastrophic events occur. The National Aeronautics and Space Administration (NASA) has taken a lead role in the investigation of new sensor technologies and techniques for the in situ structural health monitoring of gas turbine engines. As part of this effort, microwave sensor technology has been investigated as a means of making high temperature non-contact blade tip clearance, blade tip timing, and blade vibration measurements for use in gas turbine engines. This paper presents a summary of key results and findings obtained from the evaluation of two different types of microwave sensors that have been investigated for possible use in structural health monitoring applications. The first is a microwave blade tip clearance sensor that has been evaluated on a large scale Axial Vane Fan, a subscale Turbofan, and more recently on sub-scale turbine engine like disks. The second is a novel microwave based blade vibration sensor that was also used in parallel with the microwave blade tip clearance sensors on the same experiments with the sub-scale turbine engine disks.

Direct Numerical Simulations of a Full Stationary Wind-Turbine Blade

NASA Astrophysics Data System (ADS)

Qamar, Adnan; Zhang, Wei; Gao, Wei; Samtaney, Ravi

2014-11-01

Direct numerical simulation of flow past a full stationary wind-turbine blade is carried out at Reynolds number, Re = 10,000 placed at 0 and 5 (degree) angle of attack. The study is targeted to create a DNS database for verification of solvers and turbulent models that are utilized in wind-turbine modeling applications. The full blade comprises of a circular cylinder base that is attached to a spanwise varying airfoil cross-section profile (without twist). An overlapping composite grid technique is utilized to perform these DNS computations, which permits block structure in the mapped computational space. Different flow shedding regimes are observed along the blade length. Von-Karman shedding is observed in the cylinder shaft region of the turbine blade. Along the airfoil cross-section of the blade, near body shear layer breakdown is observed. A long tip vortex originates from the blade tip region, which exits the computational plane without being perturbed. Laminar to turbulent flow transition is observed along the blade length. The turbulent fluctuations amplitude decreases along the blade length and the flow remains laminar regime in the vicinity of the blade tip. The Strouhal number is found to decrease monotonously along the blade length. Average lift and drag coefficients are also reported for the cases investigated. Supported by funding under a KAUST OCRF-CRG grant.

Loads and performance data from a wind-tunnel test of model articulated helicopter rotors with 2 different blade torsional stiffnesses

NASA Technical Reports Server (NTRS)

Yeager, W. T., Jr.; Mantay, W. R.

1983-01-01

A passive means of tailoring helicopter rotor blades to improve performance and reduce loads was evaluated. The parameters investigated were blade torsional stiffness, blade section camber, and distance between blade structural elastic axis and blade tip aerodynamic center. This offset was accomplished by sweeping the tip. The investigation was conducted at advance ratios of 0.20, 0.30, and 0.40. Data are presented without analysis; however, cross referencing of performance data and harmonic loads data may be useful to the analyst for validating aeroelastic theories and design methodologies as well as for evaluating passive aeroelastic tailoring or rotor blade parameters.

Flow Instability and Flow Control Scaling Laws

NASA Astrophysics Data System (ADS)

van Ness, Daniel; Corke, Thomas; Morris, Scott

2006-11-01

A flow instability that is receptive to perturbations is present in the tip clearance leakage flow over the tip of a turbine blade. This instability was investigated through the introduction of active flow control in the viscous flow field. Control was implemented in the form of a dielectric barrier discharge created by a weakly-ionized plasma actuation arrangement. The experimental setup consisted of a low-speed linear turbine cascade made up of an array of nine Pratt & Whitney ``PakB'' turbine blades. This idealized cascade configuration was used to examine the tip clearance leakage flow that exists within the low pressure turbine stage of a gas-turbine engine. The center blade of the cascade array had a variable tip clearance up to five percent chord. Reynolds numbers based on axial blade chord varied from 10^4 to 10^5. Multi-port pressure probe measurements, as well as Stereo Particle Image Velocimetry were used to document the dependence of the instability on the frequency and amplitude of flow control perturbations. Scaling laws based on the variation of blade tip clearance height and inflow conditions were investigated. These results permitted an improved understanding of the mechanism of flow instability.

Single-shot pressure-sensitive paint lifetime measurements on fast rotating blades using an optimized double-shutter technique

NASA Astrophysics Data System (ADS)

Weiss, Armin; Geisler, Reinhard; Schwermer, Till; Yorita, Daisuke; Henne, Ulrich; Klein, Christian; Raffel, Markus

2017-09-01

A pressure-sensitive paint (PSP) system is presented to measure global surface pressures on fast rotating blades. It is dedicated to solve the problem of blurred image data employing the single-shot lifetime method. The efficient blur reduction capability of an optimized double-shutter imaging technique is demonstrated omitting error-prone post-processing or laborious de-rotation setups. The system is applied on Mach-scaled DSA-9A helicopter blades in climb at various collective pitch settings and blade tip Mach and chord Reynolds numbers (M_{ {tip}} = 0.29-0.57; Re_{ {tip}} = 4.63-9.26 × 10^5). Temperature effects in the PSP are corrected by a theoretical approximation validated against measured temperatures using temperature-sensitive paint (TSP) on a separate blade. Ensemble-averaged PSP results are comparable to pressure-tap data on the same blade to within 250 Pa. Resulting pressure maps on the blade suction side reveal spatially high resolved flow features such as the leading edge suction peak, footprints of blade-tip vortices and evidence of laminar-turbulent boundary-layer (BL) transition. The findings are validated by a separately conducted BL transition measurement by means of TSP and numerical simulations using a 2D coupled Euler/boundary-layer code. Moreover, the principal ability of the single-shot technique to capture unsteady flow phenomena is stressed revealing three-dimensional pressure fluctuations at stall.

A Preliminary Study of a Propeller Powered by Gas Jets Issuing from the Blade Tips

DTIC Science & Technology

1946-11-01

ISSUING FROM THE BLADE TIPS By J. C. Sanders and N. D. Sanders Aircraft Engine Research Laboratory Cleveland, Ohio icaflit w<• w &£N •^5$" jm "^o*6w...propeller powered by Jets in the blade tips made by Roy in 1930 (reference 3) showed that this engine would be less efficient than;a reciprocating...development of the turbojet engine , which is .now of outstanding interest. The possibilities of the jet -operated propeller are re-exeroined and the

Investigation of Blade-row Flow Distributions in Axial-flow-compressor Stage Consisting of Guide Vanes and Rotor-blade Row

NASA Technical Reports Server (NTRS)

Mahoney, John J; Dugan, Paul D; Budinger, Raymond E; Goelzer, H Fred

1950-01-01

A 30-inch tip-diameter axial-flow compressor stage was investigated with and without rotor to determine individual blade-row performance, interblade-row effects, and outer-wall boundary-layer conditions. Velocity gradients at guide-vane outlet without rotor approximated design assumptions, when the measured variation of leaving angle was considered. With rotor in operation, Mach number and rotor-blade effects changed flow distribution leaving guide vanes and invalidated design assumption of radial equilibrium. Rotor-blade performance correlated interpolated two-dimensional results within 2 degrees, although tip stall was indicated in experimental and not two-dimensional results. Boundary-displacement thickness was less than 1.0 and 1.5 percent of passage height after guide vanes and after rotor, respectively, but increased rapidly after rotor when tip stall occurred.

Wind Tunnel Measurements of the Wake of a Full-Scale UH-60A Rotor in Forward Flight

NASA Technical Reports Server (NTRS)

Wadcock, Alan J.; Yamauchi, Gloria K.; Schairer, Edward T.

2013-01-01

A full-scale UH-60A rotor was tested in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel in May 2010. The test was designed to acquire a suite of measurements to validate state-of-the-art modeling tools. Measurements include blade airloads (from a single pressure-instrumented blade), blade structural loads (strain gages), rotor performance (rotor balance and torque measurements), blade deformation (stereo-photogrammetry), and rotor wake measurements (Particle Image Velocimetry (PIV) and Retro-reflective Backward Oriented Schlieren (RBOS)). During the test, PIV measurements of flow field velocities were acquired in a stationary cross-flow plane located on the advancing side of the rotor disk at approximately 90 deg rotor azimuth. At each test condition, blade position relative to the measurement plane was varied. The region of interest (ROI) was 4-ft high by 14-ft wide and covered the outer half of the blade radius. Although PIV measurements were acquired in only one plane, much information can be gleaned by studying the rotor wake trajectory in this plane, especially when such measurements are augmented by blade airloads and RBOS data. This paper will provide a comparison between PIV and RBOS measurements of tip vortex position and vortex filament orientation for multiple rotor test conditions. Blade displacement measurements over the complete rotor disk will also be presented documenting blade-to-blade differences in tip-path-plane and providing additional information for correlation with PIV and RBOS measurements of tip vortex location. In addition, PIV measurements of tip vortex core diameter and strength will be presented. Vortex strength will be compared with measurements of maximum bound circulation on the rotor blade determined from pressure distributions obtained from 235 pressure sensors distributed over 9 radial stations.

Ramp-integration technique for capacitance-type blade-tip clearance measurement

NASA Astrophysics Data System (ADS)

Sarma, Garimella R.; Barranger, John P.

The analysis of a proposed new technique for capacitance type blade tip clearance measurement is presented. The capacitance between the blade tip and a mounted capacitance electrode within a guard ring forms one of the feedback elements of a high speed operational amplifier. The differential equation governing the operational amplifier circuit is formulated and solved for two types of inputs to the amplifier - a constant voltage and a ramp. The resultant solution shows an output that contains a term that is proportional to the derivative of the product of the input voltage and the time constant of the feedback network. The blade tip clearance capacitance is obtained by subtracting the output of a balancing reference channel followed by integration. The proposed sampled data algorithm corrects for environmental effects and varying rotor speeds on-line, making the system suitable for turbine instrumentation. System requirements, block diagrams, and a typical application are included.

Ramp-integration technique for capacitance-type blade-tip clearance measurement

NASA Astrophysics Data System (ADS)

Sarma, G. R.; Barranger, J. P.

1986-05-01

The analysis of a proposed new technique for capacitance type blade tip clearance measurement is presented. The capacitance between the blade tip and a mounted capacitance electrode within a guard ring forms one of the feedback elements of a high speed operational amplifier. The differential equation governing the operational amplifier circuit is formulated and solved for two types of inputs to the amplifier - a constant voltage and a ramp. The resultant solutions shows an output that contains a term that is proportional to the derivative of the product of the input voltage and the time constant of the feedback network. The blade tip clearance capacitance is obtained by subtracting the output of a balancing reference channel followed by integration. The proposed sampled data algorithm corrects the environmental effects and varying rotor speeds on-line, making the system suitable for turbine instrumentation. System requirements, block diagrams, and typical application are included.

The effect of circumferential distortion on fan performance at two levels of blade loading

NASA Technical Reports Server (NTRS)

Hartmann, M. J.; Sanger, N. L.

1975-01-01

Single stage fans designed for two levels of pressure ratio or blade loading were subjected to screen-induced circumferential distortions of 90-degree extent. Both fan rotors were designed for a blade tip speed of 425 m/sec, blade solidity of 1.3 and a hub-to-tip radius ratio of 0.5. Circumferential measurements of total pressure, temperature, static pressure, and flow angle were obtained at the hub, mean and tip radii at five axial stations. Rotor loading level did not appear to have a significant influence on rotor response to distorted flow. Losses in overall pressure ratio due to distortion were most severe in the stator hub region of the more highly loaded stage. At the near stall operating condition tip and hub regions of (either) rotor demonstrated different response characteristics to the distorted flow. No effect of loading was apparent on interactions between rotor and upstream distorted flow fields.

Robust Optimization Design for Turbine Blade-Tip Radial Running Clearance using Hierarchically Response Surface Method

NASA Astrophysics Data System (ADS)

Zhiying, Chen; Ping, Zhou

2017-11-01

Considering the robust optimization computational precision and efficiency for complex mechanical assembly relationship like turbine blade-tip radial running clearance, a hierarchically response surface robust optimization algorithm is proposed. The distribute collaborative response surface method is used to generate assembly system level approximation model of overall parameters and blade-tip clearance, and then a set samples of design parameters and objective response mean and/or standard deviation is generated by using system approximation model and design of experiment method. Finally, a new response surface approximation model is constructed by using those samples, and this approximation model is used for robust optimization process. The analyses results demonstrate the proposed method can dramatic reduce the computational cost and ensure the computational precision. The presented research offers an effective way for the robust optimization design of turbine blade-tip radial running clearance.

Ramp-integration technique for capacitance-type blade-tip clearance measurement

NASA Technical Reports Server (NTRS)

Sarma, Garimella R.; Barranger, John P.

1986-01-01

The analysis of a proposed new technique for capacitance type blade tip clearance measurement is presented. The capacitance between the blade tip and a mounted capacitance electrode within a guard ring forms one of the feedback elements of a high speed operational amplifier. The differential equation governing the operational amplifier circuit is formulated and solved for two types of inputs to the amplifier - a constant voltage and a ramp. The resultant solution shows an output that contains a term that is proportional to the derivative of the product of the input voltage and the time constant of the feedback network. The blade tip clearance capacitance is obtained by subtracting the output of a balancing reference channel followed by integration. The proposed sampled data algorithm corrects for environmental effects and varying rotor speeds on-line, making the system suitable for turbine instrumentation. System requirements, block diagrams, and a typical application are included.

Ramp-integration technique for capacitance-type blade-tip clearance measurement

NASA Technical Reports Server (NTRS)

Sarma, G. R.; Barranger, J. P.

1986-01-01

The analysis of a proposed new technique for capacitance type blade tip clearance measurement is presented. The capacitance between the blade tip and a mounted capacitance electrode within a guard ring forms one of the feedback elements of a high speed operational amplifier. The differential equation governing the operational amplifier circuit is formulated and solved for two types of inputs to the amplifier - a constant voltage and a ramp. The resultant solutions shows an output that contains a term that is proportional to the derivative of the product of the input voltage and the time constant of the feedback network. The blade tip clearance capacitance is obtained by subtracting the output of a balancing reference channel followed by integration. The proposed sampled data algorithm corrects the environmental effects and varying rotor speeds on-line, making the system suitable for turbine instrumentation. System requirements, block diagrams, and typical application are included.

Impact behavior of filament wound graphite/epoxy fan blades

NASA Technical Reports Server (NTRS)

Bowles, K. J.

1978-01-01

The fabrication and impact tests of graphite/epoxy filament wound fan blades are discussed. Blades which were spin tested at tip speeds up to 305 meters per second retained their structural integrity. Two blades were each impacted with a 454 gram slice of a 908 gram simulated bird at a tip speed of 263 meters per second and impact angles of 22 and 32 deg. The impact tests were recorded with high-speed movie film. The blade which was impacted at 22 deg sustained some root delamination but remained intact. The 32 deg impact separated the blade from the root. No local damage other than leading edge debonding was observed for either blade. Results of a failure mode analysis are also discussed.

Laser anemometer measurements in a transonic axial-flow fan rotor

NASA Technical Reports Server (NTRS)

Strazisar, Anthony J.; Wood, Jerry R.; Hathaway, Michael D.; Suder, Kenneth L.

1989-01-01

Laser anemometer surveys were made of the 3-D flow field in NASA rotor 67, a low aspect ratio transonic axial-flow fan rotor. The test rotor has a tip relative Mach number of 1.38. The flowfield was surveyed at design speed at near peak efficiency and near stall operating conditions. Data is presented in the form of relative Mach number and relative flow angle distributions on surfaces of revolution at nine spanwise locations evenly spaced from hub to tip. At each spanwise location, data was acquired upstream, within, and downstream of the rotor. Aerodynamic performance measurements and detailed rotor blade and annulus geometry are also presented so that the experimental results can be used as a test case for 3-D turbomachinery flow analysis codes.

Reconstruction of a three-dimensional, transonic rotor flow field from holographic interferogram data

NASA Technical Reports Server (NTRS)

Kittleson, John K.; Yu, Yung H.

1987-01-01

Holographic interferometry and computerized aided tomography (CAT) are used to determine the transonic velocity field of a model rotor blade in hover. A pulsed ruby laser recorded 40 interferograms with a 2 ft dia view field near the model rotor blade tip operating at a tip Mach number of 0.90. After digitizing the interferograms and extracting the fringe order functions, the data are transferred to a CAT code. The CAT code then calculates the perturbation velocity in several planes above the blade surface. The values from the holography-CAT method compare favorably with previously obtained numerical computations in most locations near the blade tip. The results demonstrate the technique's potential for three dimensional transonic rotor flow studies.

Turbo machine tip clearance and vibration measurements using a fibre optic laser Doppler position sensor

NASA Astrophysics Data System (ADS)

Pfister, T.; Büttner, L.; Czarske, J.; Krain, H.; Schodl, R.

2006-07-01

This paper presents a novel fibre optic laser Doppler position sensor for single blade tip clearance and vibration measurements at turbo machines, which offers high temporal resolution and high position resolution simultaneously. The sensor principle is based on the generation of a measurement volume consisting of two superposed fan-like interference fringe systems with contrary fringe spacing gradients using wavelength division multiplexing. A flexible and robust measurement system with an all-passive fibre coupled measurement head has been realized employing diffractive and refractive optics. Measurements of tip clearance and rotor vibrations at a transonic centrifugal compressor performed during operation at up to 50 000 rpm (833 Hz) corresponding to 21.7 kHz blade frequency and 586 m s-1 blade tip velocity are presented. The results are in excellent agreement with those of capacitive probes. The mean uncertainty of the position measurement was around 20 µm and, thus, considerably better than for conventional tip clearance probes. Consequently, this sensor is capable of fulfilling the requirements for future active clearance control systems and has great potential for in situ and online tip clearance and vibration measurements at metallic and non-metallic turbine blades with high precision.

Light intensity of curved laryngoscope blades in Philadelphia emergency departments.

PubMed

Levitan, Richard M; Kelly, John J; Kinkle, William C; Fasano, Charles

2007-09-01

Laryngoscopy and tracheal intubation requires laryngeal exposure and illumination. The objective of this study is to assess variation in laryngoscope lights across different emergency departments (EDs). A convenience sample of 3 Mac #4 blade and handle pairs in each of 17 Philadelphia area EDs was tested with a digital light meter to derive the median lux at the distal tip. For each blade tested, we characterized blade design (American, English, or German) and light type (fiber-illuminated versus conventional bulb-on-blade) and measured light-to-tip distance. A total of 50 blades and handle pairs were tested (one ED had only 2 Mac #4 blades). American designs were the most common (38/50), followed by English (6/50) and German (3/50) designs. Three blades had hybrid design features and acrylic light-conducting fibers. Median luminance varied from 11 lux to 5,627 lux. The glass fiber-illuminated blades (n=13) produced greater luminance (median 1,205 lux; interquartile range [IQR] 726 to 2,176 lux) than bulb-on-blade designs (median 689 lux; IQR 290 to 906 lux). German fiber-illuminated blades produced the highest luminance (median 1,937 lux; IQR 1,453 to 3,782 lux). English bulb-on-blade designs produced more luminance (median 915 lux; IQR 745 to 1270 lux) than American (median 689 lux; IQR 269 to 807 lux). German and English blades had shorter light-to-tip distances (median 51 mm and 47 mm, respectively) than American blades (65 mm). Curved laryngoscope blades in different EDs have marked variation in light intensity. The contribution of luminance to laryngoscopy performance warrants investigation.

Turbine engine rotor health monitoring evaluation by means of finite element analyses and spin tests data

NASA Astrophysics Data System (ADS)

Abdul-Aziz, Ali; Woike, Mark R.; Clem, Michelle; Baaklini, George Y.

2014-04-01

Generally, rotating engine components undergo high centrifugal loading environment which subject them to various types of failure initiation mechanisms. Health monitoring of these components is a necessity and is often challenging to implement. This is primarily due to numerous factors including the presence of scattered loading conditions, flaw sizes, component geometry and materials properties, all which hinder the simplicity of applying health monitoring applications. This paper represents a summary work of combined experimental and analytical modeling that included data collection from a spin test experiment of a rotor disk addressing the aforementioned durability issues. It further covers presentation of results obtained from a finite element modeling study to characterize the structural durability of a cracked rotor as it relates to the experimental findings. The experimental data include blade tip clearance, blade tip timing and shaft displacement measurements. The tests were conducted at the NASA Glenn Research Center's Rotordynamics Laboratory, a high precision spin rig. The results are evaluated and examined to determine their significance on the development of a health monitoring system to pre-predict cracks and other anomalies and to assist in initiating a supplemental physics based fault prediction analytical model.

Observations from varying the lift and drag inputs to a noise prediction method for supersonic helical tip speed propellers

NASA Technical Reports Server (NTRS)

Dittmar, J. H.

1984-01-01

Previous comparisons between calculated and measured supersonic helical tip speed propeller noise show them to have different trends of peak blade passing tone versus helical tip Mach number. It was postulated that improvements in this comparison could be made first by including the drag force terms in the prediction and then by reducing the blade lift terms at the tip to allow the drag forces to dominate the noise prediction. Propeller hub to tip lift distributions were varied, but they did not yield sufficient change in the predicted lift noise to improve the comparison. This result indicates that some basic changes in the theory may be needed. In addition, the noise predicted by the drag forces did not exhibit the same curve shape as the measured data. So even if the drag force terms were to dominate, the trends with helical tip Mach number for theory and experiment would still not be the same. The effect of the blade shock wave pressure rise was approxmated by increasing the drag coefficient at the blade tip. Predictions using this shock wdave approximation did have a curve shape similar to the measured data. This result indicates that the shock pressure rise probably controls the noise at supersonic tip speed and that the linear prediction method can give the proper noise trend with Mach number.

Development and experimental characterization of a new non contact sensor for blade tip timing

NASA Astrophysics Data System (ADS)

Brouckaert, Jean-Francois; Marsili, Roberto; Rossi, Gianluca; Tomassini, Roberto

2012-06-01

Performances of blade tip timing measurement systems (BTT), recently used for non contact turbine blade vibration measurements, in terms of uncertainty and resolution are strongly affected by sensor characteristics. The sensors used for BTT generate pulses, to be used also for precise measurements of turbine blades time of arrival. All the literature on this measurement techniques do not address this problem in a clear way, defining the relevant dynamic and static sensor characteristics, fundamental for this application. Till now proximity sensors used are based on optical, capacitive, eddy current and microwave measuring principle. Also pressure sensors has been used. In this paper a new sensing principle is proposed. A proximity sensor based on magnetoresistive sensing element has been assembled end tested. A simple and portable test bench with variable speed, blade tip width, variable clearance was built and used in order to characterize the main sensor performances.

Holographic studies of shock waves within transonic fan rotors

NASA Technical Reports Server (NTRS)

Benser, W. A.; Bailey, E. E.; Gelder, T. F.

1974-01-01

NASA has funded two separate contracts to apply pulsed laser holographic interferometry to the detection of shock patterns in the outer span regions of high tip speed transonic rotors. The first holographic approach used ruby laser light reflected from a portion of the centerbody just ahead of the rotor. These holograms showed the bow wave patterns upstream of the rotor and the shock patterns just inside the blade row near the tip. The second holographic approach, on a different rotor, used light transmitted diagonally across the inlet annulus past the centerbody. This approach gave a more extensive view of the region bounded by the blade leading and trailing edges, by the part span shroud and by the blade tip. These holograms showed the passage shock emanating from the blade leading edge and a moderately strong conical shock originating at the intersection of the part span shroud leading edge and the blade suction surface.

Flow visualizations of perpendicular blade vortex interactions

NASA Technical Reports Server (NTRS)

Rife, Michael C.; Davenport, William J.

1992-01-01

Helium bubble flow visualizations have been performed to study perpendicular interaction of a turbulent trailing vortex and a rectangular wing in the Virginia Tech Stability Tunnel. Many combinations of vortex strength, vortex-blade separation (Z(sub s)) and blade angle of attack were studied. Photographs of representative cases are presented. A range of phenomena were observed. For Z(sub s) greater than a few percent chord the vortex is deflected as it passes the blade under the influence of the local streamline curvature and its image in the blade. Initially the interaction appears to have no influence on the core. Downstream, however, the vortex core begins to diffuse and grow, presumably as a consequence of its interaction with the blade wake. The magnitude of these effects increases with reduction in Z(sub s). For Z(sub s) near zero the form of the interaction changes and becomes dependent on the vortex strength. For lower strengths the vortex appears to split into two filaments on the leading edge of the blade, one passing on the pressure and one passing on the suction side. At higher strengths the vortex bursts in the vicinity of the leading edge. In either case the core of its remnants then rapidly diffuse with distance downstream. Increase in Reynolds number did not qualitatively affect the flow apart from decreasing the amplitude of the small low-frequency wandering motions of the vortex. Changes in wing tip geometry and boundary layer trip had very little effect.

Tabulation of data from the tip aerodynamics and acoustics test

NASA Technical Reports Server (NTRS)

Cross, Jeffrey L.; Tu, Wilson

1990-01-01

In a continuing effort to understand helicopter rotor tip aerodynamics and acoustics, researchers at Ames Research Center conducted a flight test. The test was performed using the NASA White Cobra and a set of highly instrumented blades. Tabular and graphic summaries of two data subsets from the Tip Aerodynamics and Acoustics Test are given. The data presented are for airloads, blade structural loads, blade vibrations, with summary tables of the aircraft states for each test point. The tabular data consist of the first 15 harmonics only, whereas the plots contain the entire measured frequency content.

Multiple piece turbine rotor blade

DOEpatents

Kimmel, Keith D.; Plank, William L.

2016-07-19

A spar and shell turbine rotor blade with a spar and a tip cap formed as a single piece, the spar includes a bottom end with dovetail or fir tree slots that engage with slots on a top end of a root section, and a platform includes an opening on a top surface for insertion of the spar in which a shell made from an exotic high temperature resistant material is secured between the tip cap and the platform. The spar is tapered to form thinner walls at the tip end to further reduce the weight and therefore a pulling force due to blade rotation. The spar and tip cap piece is made from a NiAL material to further reduce the weight and the pulling force.

Computational Aerodynamic Simulations of a 1215 ft/sec Tip Speed Transonic Fan System Model for Acoustic Methods Assessment and Development

NASA Technical Reports Server (NTRS)

Tweedt, Daniel L.

2014-01-01

Computational Aerodynamic simulations of a 1215 ft/sec tip speed transonic fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, low-noise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating points simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, which for this model did not include a split flow path with core and bypass ducts. As a result, it was only necessary to adjust fan rotational speed in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. Computed blade row flow fields at all fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the flow fields at all operating conditions reveals no excessive boundary layer separations or related secondary-flow problems.

Fan Noise Source Diagnostic Test: LDV Measured Flow Field Results

NASA Technical Reports Server (NTRS)

Podboy, Gary C.; Krupar, Martin J.; Hughes, Christopher E.; Woodward, Richard P.

2003-01-01

Results are presented of an experiment conducted to investigate potential sources of noise in the flow developed by two 22-in. diameter turbofan models. The R4 and M5 rotors that were tested were designed to operate at nominal take-off speeds of 12,657 and 14,064 RPMC, respectively. Both fans were tested with a common set of swept stators installed downstream of the rotors. Detailed measurements of the flows generated by the two were made using a laser Doppler velocimeter system. The wake flows generated by the two rotors are illustrated through a series of contour plots. These show that the two wake flows are quite different, especially in the tip region. These data are used to explain some of the differences in the rotor/stator interaction noise generated by the two fan stages. In addition to these wake data, measurements were also made in the R4 rotor blade passages. These results illustrate the tip flow development within the blade passages, its migration downstream, and (at high rotor speeds) its merging with the blade wake of the adjacent (following) blade. Data also depict the variation of this tip flow with tip clearance. Data obtained within the rotor blade passages at high rotational speeds illustrate the variation of the mean shock position across the different blade passages.

Optical Detection of Blade Flutter

NASA Technical Reports Server (NTRS)

Nieberding, W. C.; Pollack, J. L.

1977-01-01

Dynamic strain gages mounted on rotor blades are used as the primary instrumentation for detecting the onset of flutter and defining the vibratory mode and frequency. Optical devices are evaluated for performing the same measurements as well as providing supplementary information on the vibratory characteristics. Two separate methods are studied: stroboscopic imagery of the blade tip and photoelectric scanning of blade tip motion. Both methods give visual data in real time as well as video tape records. The optical systems are described, and representative results are presented. The potential of this instrumentation in flutter research is discussed.

Impact behavior of filament-wound graphite/epoxy fan blades. [foreign object damage to turbofan engines

NASA Technical Reports Server (NTRS)

Bowles, K. J.

1978-01-01

The fabrication and impact tests of graphite/epoxy filament wound fan blades are discussed. Blades which were spin tested at tip speeds up to 305 m/sec retained their structural integrity. Two blades were each impacted with a 454-g slice of a 908-g simulated bird at a tip speed of 263 deg and impact angles of 22 deg and 32 deg. The impact tests were recorded with high-speed movie film. The blade which was impacted at 22 deg sustained some root delamination but remained intact. The 32 deg impact separated the blade from the root. No local damage other than leading-edge debonding was observed for either blade. The results of a failure mode analysis are also discussed.

Vibration Based Crack Detection in a Rotating Disk. Part 2; Experimental Results

NASA Technical Reports Server (NTRS)

Gyekenyesi, Andrew L.; Sawicki, Jerzy T.; Martin, Richard E.; Haase, Wayne C.; Baaklini, George

2005-01-01

This paper describes the experimental results concerning the detection of a crack in a rotating disk. The goal was to utilize blade tip clearance and shaft vibration measurements to monitor changes in the system's center of mass and/or blade deformation behaviors. The concept of the approach is based on the fact that the development of a disk crack results in a distorted strain field within the component. As a result, a minute deformation in the disk's geometry as well as a change in the system's center of mass occurs. Here, a notch was used to simulate an actual crack. The vibration based experimental results failed to identify the existence of a notch when utilizing the approach described above, even with a rather large, circumferential notch (l.2 in.) located approximately mid-span on the disk (disk radius = 4.63 in. with notch at r = 2.12 in.). This was somewhat expected, since the finite element based results in Part 1 of this study predicted changes in blade tip clearance as well as center of mass shifts due to a notch to be less than 0.001 in. Therefore, the small changes incurred by the notch could not be differentiated from the mechanical and electrical noise of the rotor system. Although the crack detection technique of interest failed to identify the existence ofthe notch, the vibration data produced and captured here will be utilized in upcoming studies that will focus on different data mining techniques concerning damage detection in a disk.

Theoretical Evaluation of Methods of Cooling the Blades of Gas Turbines

NASA Technical Reports Server (NTRS)

Sanders, J. C.; Mendelson, Alexander

1947-01-01

A study was made of heat transfer in turbine blades and the effects on blade temperature of cooling the blade root and tip, changing the dimensions of the blades, raising the cycle temperatures, insulating with ceramics, and cooling by circulation of air or water through hollow blades.

Application of two-dimensional unsteady aerodynamic to a free-tip rotor response analysis

NASA Technical Reports Server (NTRS)

Yates, L.; Kumagai, H.

1985-01-01

The free-tip rotor utilizes a rotor blade tip which is structurally decoupled from the blade inboard section. The tip is free to pitch about its own pitch axis to respond to the local flow angularity changes. The tip also experiences the heaving motion due to the flapping of the rotor blade. For an airfoil in any pitching and heaving motion which can be expanded into a Fourier series, the lift and moment calculated by Theodoren's theory is simply the linear combination of the lift and moment calculated for each harmonic. These lift and moment are then used to determine the response of the free-tip rotor. A parametric study is performed to determine the effect of mechanical damping, mechanical spring, sweep, friction, and a constant control moment on the free-tip rotor response characteristics and the resulting azimuthal lift distributions. The results showed that the free-tip has the capability to suppress the oscillatory lift distribution around the azimuth and to eliminate a significant negative life peak on the advancing tip. This result agrees with the result of the previous analysis based on the steady aerodynamics.

Numerical Investigation on the Effects of Self-Excited Tip Flow Unsteadiness and Blade Row Interactions on the Performance Predictions of Low Speed and Transonic Compressor Rotors

NASA Astrophysics Data System (ADS)

Lee, Daniel H.

The impact blade row interactions can have on the performance of compressor rotors has been well documented. It is also well known that rotor tip clearance flows can have a large effect on compressor performance and stall margin and recent research has shown that tip leakage flows can exhibit self-excited unsteadiness at near stall conditions. However, the impact of tip leakage flow on the performance and operating range of a compressor rotor, relative to other important flow features such as upstream stator wakes or downstream potential effects, has not been explored. To this end, a numerical investigation has been conducted to determine the effects of self-excited tip flow unsteadiness, upstream stator wakes, and downstream blade row interactions on the performance prediction of low speed and transonic compressor rotors. Calculations included a single blade-row rotor configuration as well as two multi-blade row configurations: one where the rotor was modeled with an upstream stator and a second where the rotor was modeled with a downstream stator. Steady-state and time accurate calculations were performed using a RANS solver and the results were compared with detailed experimental data obtained in the GE Low Speed Research Compressor and the Notre Dame Transonic Rig at several operating conditions including near stall. Differences in the performance predictions between the three configurations were then used to determine the effect of the upstream stator wakes and the downstream blade row interactions. Results obtained show that for both the low speed and transonic research compressors used in this investigation time-accurate RANS analysis is necessary to accurately predict the stalling character of the rotor. Additionally, for the first time it is demonstrated that capturing the unsteady tip flow can have a larger impact on rotor performance predictions than adjacent blade row interactions.

Cruise noise of the SR-2 propeller model in a wind tunnel

NASA Technical Reports Server (NTRS)

Dittmar, James H.

1989-01-01

Noise data on the SR-2 model propeller were taken in the NASA Lewis Research Center 8- by 6-Foot Wind Tunnel. The maximum blade passing tone rises with increasing helical tip Mach number to a peak level at a helical tip Mach number of about 1.05; then it remains the same or decreases at higher helical tip Mach numbers. This behavior, which has been observed with other propeller models, points to the possibility of using higher propeller tip speeds to limit airplane cabin noise while maintaining high flight speed and efficiency. Noise comparisons of the straight-blade SR-2 propeller and the swept-blade SR-7A propeller showed that the tailored sweep of the SR-7A appears to be the cause of both lower peak noise levels and a slower noise increase with increasing helical tip Mach number.

Experimental quiet engine program

NASA Technical Reports Server (NTRS)

Cornell, W. G.

1975-01-01

Full-scale low-tip-speed fans, a full-scale high-tip-speed fan, scale model versions of fans, and two full-scale high-bypass-ratio turbofan engines, were designed, fabricated, tested, and evaluated. Turbine noise suppression was investigated. Preliminary design studies of flight propulsion system concepts were used in application studies to determine acoustic-economic tradeoffs. Salient results are as follows: tradeoff evaluation of fan tip speed and blade loading; systematic data on source noise characteristics and suppression effectiveness; documentation of high- and low-fan-speed aerodynamic and acoustic technology; aerodynamic and acoustic evaluation of acoustic treatment configurations, casing tip bleed, serrated and variable pitch rotor blades, leaned outlet guide vanes, slotted tip casings, rotor blade shape modifications, and inlet noise suppression; systematic evaluation of aerodynamic and acoustic effects; flyover noise projections of engine test data; turbine noise suppression technology development; and tradeoff evaluation of preliminary design high-fan-speed and low-fan-speed flight engines.

Towards hybrid mesh generation for realistic design environments

NASA Astrophysics Data System (ADS)

McMorris, Harlan Tom

Two different techniques that allow hybrid mesh generation to be easily used in the design environment are presented. The purpose of this research is to allow for hybrid meshes to be used during the design process where the geometry is being varied. The first technique, modular hybrid mesh generation, allows for the replacement of portions of a geometry with a new design shape. The mesh is maintained for the portions of that have not changed during the design process. A new mesh is generated for the new part of the geometry and this piece is added to the existing mesh. The new mesh must match the remaining portions of the geometry plus the element sizes must match exactly across the interfaces. The second technique, hybrid mesh movement, is used when the basic geometry remains the same with only small variations to portions of the geometry. These small variations include changing the cross-section of a wing, twisting a blade or changing the length of some portion of the geometry. The mesh for the original geometry is moved onto the new geometry one step at a time beginning with the curves of the surface, continuing with the surface mesh geometry and ending with the interior points of the mesh. The validity of the hybrid mesh is maintained by applying corrections to the motion of the points. Finally, the quality of the new hybrid mesh is improved to ensure that the new mesh maintains the quality of the original hybrid mesh. Applications of both design techniques are applied to typical example cases from the fields of turbomachinery, aerospace and offshore technology. The example test cases demonstrate the ability of the two techniques to reuse the majority of an existing hybrid mesh for typical design changes. Modular mesh generation is used to change the shape of piece of a seafloor pipeline geometry to a completely different configuration. The hybrid mesh movement technique is used to change the twist of a turbomachinery blade, the tip clearance of a rotor blade and to simulate the aeroelastic bending of a wing. Finally, the movement technique is applied to an offshore application where the solution for the original configuration is used as a starting point for solution for a new configuration. The application of both techniques show that the methods can be a powerful addition to the design environment and will facilitate a rapid turnaround when the design geometry changes.

RANS computations of tip vortex cavitation

NASA Astrophysics Data System (ADS)

Decaix, Jean; Balarac, Guillaume; Dreyer, Matthieu; Farhat, Mohamed; Münch, Cécile

2015-12-01

The present study is related to the development of the tip vortex cavitation in Kaplan turbines. The investigation is carried out on a simplified test case consisting of a NACA0009 blade with a gap between the blade tip and the side wall. Computations with and without cavitation are performed using a R ANS modelling and a transport equation for the liquid volume fraction. Compared with experimental data, the R ANS computations turn out to be able to capture accurately the development of the tip vortex. The simulations have also highlighted the influence of cavitation on the tip vortex trajectory.

Sound from a Two-Blade Propeller at Supersonic Tip Speeds

NASA Technical Reports Server (NTRS)

Hubbard, Harvey H; Lassiter, Leslie W

1952-01-01

Report presents the results of sound measurements at static conditions made for a two-blade 47-inch-diameter propeller in the tip Mach number range 0.75 to 1.30. For comparison, spectrums have been obtained at both subsonic and supersonic tip speeds. In addition, the measured data are compared with calculations by the theory of Gutin which has previously been found adequate for predicting the sound at subsonic tip speeds. Curves are presented from which the maximum over-all noise levels in free space may be estimated if the power, tip Mach number, and distance are known.

Dynamics of Isolated Tip Vortex Cavitation

NASA Astrophysics Data System (ADS)

Pennings, Pepijn; Bosschers, Johan; van Terwisga, Tom

2014-11-01

Performance of ship propellers and comfort levels in the surroundings are limited by various forms of cavitation. Amongst these forms tip vortex cavitation is one of the first appearing forms and is expected to be mainly responsible for the emission of broadband pressure fluctuations typically occurring between the 4th to the 7th blade passing frequency (approx. 40--70 Hz). These radiated pressure pulses are likely to excite parts of the hull structure resulting in a design compromise between efficiency and comfort. Insight is needed in the mechanism of acoustic emission from the oscillations by a tip vortex cavity. In the current experimental study the tip vortex cavity from a blade with an elliptic planform and sections based on NACA 662 - 415 with meanline a = 0 . 8 is observed using high speed shadowgraphy in combination with blade force and acoustic measurements. An analytic model describing three main cavity deformation modes is verified and used to explain the origin of a cavity eigenfrequency or ``vortex singing'' phenomenon observed by Maines and Arndt (1997) on the tip vortex cavity originating from the same blade. As no hydrodynamic sound originating from the tip vortex cavity was observed it is posed that a tip flow instability is essential for ``vortex singing.'' This research was funded by the Lloyd's Register Foundation as part of the International Institute for Cavitation Research.

Advanced blade tip seal system, volume 2

NASA Technical Reports Server (NTRS)

Zelahy, J. W.; Fairbanks, N. P.

1982-01-01

The results of the endurance and performance engine tests conducted on monocrystal/abrasive-tipped CF6-50 Stage 1 HPT blades fabricated in Task VII of MATE Project 3 are presented. Two engine tests are conducted. The endurance engine test is conducted for 1000 C cycles. The performance engine test is conducted on a variable cycle core engine. Posttest evaluation and analyses of the blades and shrouds included visual, dimensional, and destructive evaluations.

Computation of transonic flow about helicopter rotor blades

NASA Technical Reports Server (NTRS)

Arieli, R.; Tauber, M. E.; Saunders, D. A.; Caughey, D. A.

1986-01-01

An inviscid, nonconservative, three-dimensional full-potential flow code, ROT22, has been developed for computing the quasi-steady flow about a lifting rotor blade. The code is valid throughout the subsonic and transonic regime. Calculations from the code are compared with detailed laser velocimeter measurements made in the tip region of a nonlifting rotor at a tip Mach number of 0.95 and zero advance ratio. In addition, comparisons are made with chordwise surface pressure measurements obtained in a wind tunnel for a nonlifting rotor blade at transonic tip speeds at advance ratios from 0.40 to 0.50. The overall agreement between theoretical calculations and experiment is very good. A typical run on a CRAY X-MP computer requires about 30 CPU seconds for one rotor position at transonic tip speed.

Reflection plane tests of a wind turbine blade tip section with ailerons

NASA Technical Reports Server (NTRS)

Savino, J. M.; Nyland, T. W.; Birchenough, A. G.; Jordan, F. L.; Campbell, N. K.

1985-01-01

Tests were conducted in the NASA Langley 30 by 60 foot Wind Tunnel on a full scale 7.31 m (24 ft) long tip section of a wind turbine rotor blade. The blade tip section was built with ailerons on the trailing edge. The ailerons, which spanned a length of 6.1 m (20 ft), were designed so that two types could be evaluated: the plain and the balanced. The ailerons were hinged on the suction surface at the 0.62 X chord station behind the leading edge. The purpose of the tests was to measure the aerodynamic characteristics of the blade section for: an angle of attack range from 0 deg to 90 deg aileron deflections from 0 deg to -90 deg, and Reynolds numbers of 0.79 and 1.5 x 10 to the 6th power. These data were then used to determine which aileron configuration had the most desirable rotor control and aerodynamic braking characteristics. Tests were also run to determine the effects of vortex generators, leading edge roughness, and the gaps between the aileron sections on the lift, drag, and chordwise force coefficients of the blade tip section.

Rotorcraft acoustic radiation prediction based on a refined blade-vortex interaction model

NASA Astrophysics Data System (ADS)

Rule, John Allen

1997-08-01

The analysis of rotorcraft aerodynamics and acoustics is a challenging problem, primarily due to the fact that a rotorcraft continually flies through its own wake. The generation mechanism for a rotorcraft wake, which is dominated by strong, concentrated blade-tip trailing vortices, is similar to that in fixed wing aerodynamics. However, following blades encounter shed vortices from previous blades before they are swept downstream, resulting in sharp, impulsive loading on the blades. The blade/wake encounter, known as Blade-Vortex Interaction, or BVI, is responsible for a significant amount of vibratory loading and the characteristic rotorcraft acoustic signature in certain flight regimes. The present work addressed three different aspects of this interaction at a fundamental level. First, an analytical model for the prediction of trailing vortex structure is discussed. The model as presented is the culmination of a lengthy research effort to isolate the key physical mechanisms which govern vortex sheet rollup. Based on the Betz model, properties of the flow such as mass flux, axial momentum flux, and axial flux of angular momentum are conserved on either a differential or integral basis during the rollup process. The formation of a viscous central core was facilitated by the assumption of a turbulent mixing process with final vortex velocity profiles chosen to be consistent with a rotational flow mixing model and experimental observation. A general derivation of the method is outlined, followed by a comparison of model predictions with experimental vortex measurements, and finally a viscous blade drag model to account for additional effects of aerodynamic drag on vortex structure. The second phase of this program involved the development of a new formulation of lifting surface theory with the ultimate goal of an accurate, reduced order hybrid analytical/numerical model for fast rotorcraft load calculations. Currently, accurate rotorcraft airload analyses are limited by the massive computational power required to capture the small time scale events associated with BVI. This problem has two primary facets: accurate knowledge of the wake geometry, and accurate resolution of the impulsive loading imposed by a tip vortex on a blade. The present work addressed the second facet, providing a mathematical framework for solving the impulsive loading problem analytically, then asymptotically matching this solution to a low-resolution numerical calculation. A method was developed which uses continuous sheets of integrated boundary elements to model the lifting surface and wake. Special elements were developed to capture local behavior in high-gradient regions of the flow, thereby reducing the burden placed on the surrounding numerical method. Unsteady calculations for several classical cases were made in both frequency and time domain to demonstrate the performance of the method. Finally, a new unsteady, compressible boundary element method was applied to the problem of BVI acoustic radiation prediction. This numerical method, combined with the viscous core trailing vortex model, was used to duplicate the geometry and flight configuration of a detailed experimental BVI study carried out at NASA Ames Research Center. Blade surface pressure and near- and far-field acoustic radiation calculations were made. All calculations were shown to compare favorably with experimentally measured values. The linear boundary element method with non-linear corrections proved sufficient over most of the rotor azimuth, and particular in the region of the blade vortex interaction, suggesting that full non-linear CFD schemes are not necessary for rotorcraft noise prediction.

Holographic studies of shock waves within transonic fan rotors

NASA Technical Reports Server (NTRS)

Benser, W. A.; Bailey, E. E.; Gelder, T. F.

1973-01-01

Pulsed laser holographic interferometry has been applied to the detection of shock patterns in the outer span regions of high tip speed transonic rotors. The first holographic approach used ruby laser light reflected from a portion of the centerbody just ahead of the rotor. These holograms showed the bow wave patterns upstream of the rotor and the shock patterns just inside the blade row near the tip. Much of the region of interest was in the shadow of the blade leading edge and could not be visualized. The second holographic approach, on a different rotor, used light transmitted diagonally across the inlet annulus past the centerbody. This approach gave a more extensive view of the region bounded by the blade leading and trailing edges, by the part span shroud and by the blade tip. These holograms showed the passage shock emanating from the blade leading edge and a moderately strong conical shock originating at the intersection of the part span shroud leading edge and the blade suction surface. Reasonable details of the shock patterns were obtained from holograms which were made without extensive rig modifications.

Flowfield analysis of modern helicopter rotors in hover by Navier-Stokes method

NASA Technical Reports Server (NTRS)

Srinivasan, G. R.; Raghavan, V.; Duque, E. P. N.

1991-01-01

The viscous, three-dimensional, flowfields of UH60 and BERP rotors are calculated for lifting hover configurations using a Navier-Stokes computational fluid dynamics method with a view to understand the importance of planform effects on the airloads. In this method, the induced effects of the wake, including the interaction of tip vortices with successive blades, are captured as a part of the overall flowfield solution without prescribing any wake models. Numerical results in the form of surface pressures, hover performance parameters, surface skin friction and tip vortex patterns, and vortex wake trajectory are presented at two thrust conditions for UH60 and BERP rotors. Comparison of results for the UH60 model rotor show good agreement with experiments at moderate thrust conditions. Comparison of results with equivalent rectangular UH60 blade and BERP blade indicates that the BERP blade, with an unconventional planform, gives more thrust at the cost of more power and a reduced figure of merit. The high thrust conditions considered produce severe shock-induced flow separation for UH60 blade, while the BERP blade develops more thrust and minimal separation. The BERP blade produces a tighter tip vortex structure compared with the UH60 blade. These results and the discussion presented bring out the similarities and differences between the two rotors.

Wake Geometry Effects on Rotor Blade-Vortex Interaction Noise Directivity

NASA Technical Reports Server (NTRS)

Martin, R. M.; Marcolini, Michael A.; Splettstoesser, W. R.; Schultz, K.-J.

1990-01-01

Acoustic measurements from a model rotor wind tunnel test are presented which show that the directionality of rotor blade vortex interaction (BVI) noise is strongly dependent on the rotor advance ratio and disk attitude. A rotor free wake analysis is used to show that the general locus of interactions on the rotor disk is also strongly dependent on advance ratio and disk attitude. A comparison of the changing directionality of the BVI noise with changes in the interaction locations shows that the strongest noise radiation occurs in the direction of motion normal to the blade span at the time of interaction, for both advancing and retreating side BVI. For advancing side interactions, the BVI radiation angle down from the tip-path plane appears relatively insensitive to rotor operating condition and is typically between 40 and 55 deg below the disk. However, the azimuthal radiation direction shows a clear trend with descent speed, moving towards the right of the flight path with increasing descent speed. The movement of the strongest radiation direction is attributed to the movement of the interaction locations on the rotor disk with increasing descent speed.

Flow-driven rotor simulation of vertical axis tidal turbines: A comparison of helical and straight blades

NASA Astrophysics Data System (ADS)

Le, Tuyen Quang; Lee, Kwang-Soo; Park, Jin-Soon; Ko, Jin Hwan

2014-06-01

In this study, flow-driven rotor simulations with a given load are conducted to analyze the operational characteristics of a vertical-axis Darrieus turbine, specifically its self-starting capability and fluctuations in its torque as well as the RPM. These characteristics are typically observed in experiments, though they cannot be acquired in simulations with a given tip speed ratio (TSR). First, it is shown that a flow-driven rotor simulation with a two-dimensional (2D) turbine model obtains power coefficients with curves similar to those obtained in a simulation with a given TSR. 3D flowdriven rotor simulations with an optimal geometry then show that a helical-bladed turbine has the following prominent advantages over a straight-bladed turbine of the same size: an improvement of its self-starting capabilities and reduced fluctuations in its torque and RPM curves as well as an increase in its power coefficient from 33% to 42%. Therefore, it is clear that a flow-driven rotor simulation provides more information for the design of a Darrieus turbine than a simulation with a given TSR before experiments.

Turbine blade profile design method based on Bezier curves

NASA Astrophysics Data System (ADS)

Alexeev, R. A.; Tishchenko, V. A.; Gribin, V. G.; Gavrilov, I. Yu.

2017-11-01

In this paper, the technique of two-dimensional parametric blade profile design is presented. Bezier curves are used to create the profile geometry. The main feature of the proposed method is an adaptive approach of curve fitting to given geometric conditions. Calculation of the profile shape is produced by multi-dimensional minimization method with a number of restrictions imposed on the blade geometry.The proposed method has been used to describe parametric geometry of known blade profile. Then the baseline geometry was modified by varying some parameters of the blade. The numerical calculation of obtained designs has been carried out. The results of calculations have shown the efficiency of chosen approach.

Near wall cooling for a highly tapered turbine blade

DOEpatents

Liang, George [Palm City, FL

2011-03-08

A turbine blade having a pressure sidewall and a suction sidewall connected at chordally spaced leading and trailing edges to define a cooling cavity. Pressure and suction side inner walls extend radially within the cooling cavity and define pressure and suction side near wall chambers. A plurality of mid-chord channels extend radially from a radially intermediate location on the blade to a tip passage at the blade tip for connecting the pressure side and suction side near wall chambers in fluid communication with the tip passage. In addition, radially extending leading edge and trailing edge flow channels are located adjacent to the leading and trailing edges, respectively, and cooling fluid flows in a triple-pass serpentine path as it flows through the leading edge flow channel, the near wall chambers and the trailing edge flow channel.

Experimental and numerical study of the British Experimental Rotor Programme blade

NASA Technical Reports Server (NTRS)

Brocklehurst, Alan; Duque, Earl P. N.

1990-01-01

Wind-tunnel tests on the British Experimental Rotor Programme (BERP) tip are described, and the results are compared with computational fluid dynamics (CFD) results. The test model was molded using the Lynx-BERP blade tooling to provide a semispan, cantilever wing comprising the outboard 30 percent of the rotor blade. The tests included both surface-pressure measurements and flow visualization to obtain detailed information of the flow over the BERP tip for a range of angles of attack. It was observed that, outboard of the notch, favorable pressure gradients exist which ensure attached flow, and that the tip vortex also remains stable to large angles of attack. On the rotor, these features yield a very gradual break in control loads when the retreating-blade limit is eventually reached. Computational and experimental results were generally found to be in good agreement.

Blade tip, finite aspect ratio, and dynamic stall effects on the Darrieus rotor

NASA Astrophysics Data System (ADS)

Paraschivoiu, I.; Desy, P.; Masson, C.

1988-02-01

The objective of the work described in this paper was to apply the Boeing-Vertol dynamic stall model in an asymmetric manner to account for the asymmetry of the flow between the left and right sides of the rotor. This phenomenon has been observed by the flow visualization of a two-straight-bladed Darrieus rotor in the IMST water tunnel. Also introduced into the aerodynamic model are the effects of the blade tip and finite aspect ratio on the aerodynamic performance of the Darrieus wind turbine. These improvements are compatible with the double-multiple-streamtube model and have been included in the CARDAAV computer code for predicting the aerodynamic performance. Very good agreement has been observed between the test data (Sandia 17 m) and theoretical predictions; a significant improvement over the previous dynamic stall model was obtained for the rotor power at low tip speed ratios, while the inclusion of the finite aspect ratio effects enhances the prediction of the rotor power for high tip speed ratios. The tip losses and finite aspect ratio effects were also calculated for a small-scale vertical-axis wind turbine, with a two-straight-bladed (NACA 0015) rotor.

Large-Eddy Simulation of Crashback in a Ducted Propulsor

NASA Astrophysics Data System (ADS)

Jang, Hyunchul; Mahesh, Krishnan

2011-11-01

Crashback is an operating condition to quickly stop a propelled vehicle, where the propeller is rotated in the reverse direction to yield negative thrust. The crashback condition is dominated by the interaction of free stream flow with strong reverse flow. Crashback causes highly unsteady loads and flow separation on blade surface. This study uses Large-Eddy Simulation to predict the highly unsteady flow field in crashback for a ducted propulsor. Thrust mostly arises from the blade surface, but most of side-force is generated from the duct surface. Both mean and RMS of pressure are much higher on inner surface of duct, especially near blade tips. This implies that side-force on the ducted propulsor is caused by the blade-duct interaction. Strong tip leakage flow is observed behind the suction side at the tip gap. The physical source of the tip leakage flow is seen to be the large pressure difference between pressure and suction sides. The conditional average during high amplitude event shows that the tip leakage flow and pressure difference are significantly higher. This work is supported by the United States Office of Naval Research under ONR Grant N00014-05-1-0003.

A Microwave Blade Tip Clearance Sensor for Propulsion Health Monitoring

NASA Technical Reports Server (NTRS)

Woike, Mark R.; Abdul-Aziz, Ali; Bencic, Timothy J.

2010-01-01

Microwave sensor technology is being investigated by the NASA Glenn Research Center as a means of making non-contact structural health measurements in the hot sections of gas turbine engines. This type of sensor technology is beneficial in that it is accurate, it has the ability to operate at extremely high temperatures, and is unaffected by contaminants that are present in turbine engines. It is specifically being targeted for use in the High Pressure Turbine (HPT) and High Pressure Compressor (HPC) sections to monitor the structural health of the rotating components. It is intended to use blade tip clearance to monitor blade growth and wear and blade tip timing to monitor blade vibration and deflection. The use of microwave sensors for this application is an emerging concept. Techniques on their use and calibration needed to be developed. As a means of better understanding the issues associated with the microwave sensors, a series of experiments have been conducted to evaluate their performance for aero engine applications. This paper presents the results of these experiments.

Theoretical prediction of nonlinear propagation effects on noise signatures generated by subsonic or supersonic propeller or rotor-blade tips

NASA Technical Reports Server (NTRS)

Barger, R. L.

1980-01-01

The nonlinear propagation equations for sound generated by a constant speed blade tip are presented. Propagation from a subsonic tip is treated as well as the various cases that can occur at supersonic speeds. Some computed examples indicate that the nonlinear theory correlates with experimental results better than linear theory for large amplitude waves. For swept tips that generate a wave with large amplitude leading expansion, the nonlinear theory predicts a cancellation effect that results in a significant reduction of both amplitude and impulse.

High-Speed, capacitance-based tip clearance sensing

NASA Astrophysics Data System (ADS)

Haase, W. C.; Haase, Z. S.

This paper discusses recent advances in tip clearance measurement systems for turbine engines using capacitive probes. Real time measurements of individual blade pulses are generated using wideband signal processing providing 3 dB bandwidths of typically 5 MHz. Subsequent mixed-signal processing circuitry provide real-time measurements of maximum, minimum, and average clearance with latencies of one blade-to-blade time interval. Both guarded and unguarded probe configurations are possible with the system. Calibration techniques provide high accuracy measurements.

Aerodynamic Characteristics of a Two-blade NACA 10-(3)(062)-045 Propeller and of a Two-blade NACA 10-(3)(08)-045 Propeller

NASA Technical Reports Server (NTRS)

Solomon, William

1953-01-01

Characteristics are given for the two-blade NACA 10-(3)(062)-045 propeller and for the two-blade NACA 10-(3)(08)-045 propeller over a range of advance ratio from 0.5 to 3.8, through a blade-angle range from 20 degrees to 55 degrees measured at the 0.75 radius. Maximum efficiencies of the order of 91.5 to 92 percent were obtained for the propellers. The propeller with the thinner airfoil sections over the outboard portion of the blades, the NACA 10-(3)(062)-045 propeller, had lower losses at high tip speeds, the difference amounting to about 5 percent at a helical tip Mach number of 1.10.

Blade vortex interaction noise reduction techniques for a rotorcraft

NASA Technical Reports Server (NTRS)

Charles, Bruce D. (Inventor); Hassan, Ahmed A. (Inventor); Tadghighi, Hormoz (Inventor); JanakiRam, Ram D. (Inventor); Sankar, Lakshmi N. (Inventor)

1996-01-01

An active control device for reducing blade-vortex interactions (BVI) noise generated by a rotorcraft, such as a helicopter, comprises a trailing edge flap located near the tip of each of the rotorcraft's rotor blades. The flap may be actuated in any conventional way, and is scheduled to be actuated to a deflected position during rotation of the rotor blade through predetermined regions of the rotor azimuth, and is further scheduled to be actuated to a retracted position through the remaining regions of the rotor azimuth. Through the careful azimuth-dependent deployment and retraction of the flap over the rotor disk, blade tip vortices which are the primary source for BVI noise are (a) made weaker and (b) pushed farther away from the rotor disk (that is, larger blade-vortex separation distances are achieved).

Blade vortex interaction noise reduction techniques for a rotorcraft

NASA Technical Reports Server (NTRS)

Charles, Bruce D. (Inventor); JanakiRam, Ram D. (Inventor); Hassan, Ahmed A. (Inventor); Tadghighi, Hormoz (Inventor); Sankar, Lakshmi N. (Inventor)

1998-01-01

An active control device for reducing blade-vortex interactions (BVI) noise generated by a rotorcraft, such as a helicopter, comprises a trailing edge flap located near the tip of each of the rotorcraft's rotor blades. The flap may be actuated in any conventional way, and is scheduled to be actuated to a deflected position during rotation of the rotor blade through predetermined regions of the rotor azimuth, and is further scheduled to be actuated to a retracted position through the remaining regions of the rotor azimuth. Through the careful azimuth-dependent deployment and retraction of the flap over the rotor disk, blade tip vortices which are the primary source for BVI noise are (a) made weaker and (b) pushed farther away from the rotor disk (that is, larger blade-vortex separation distances are achieved).

Large Eddy Simulation of Crashback in Marine Propulsors

NASA Astrophysics Data System (ADS)

Jang, Hyunchul

Crashback is an operating condition to quickly stop a propelled vehicle, where the propeller is rotated in the reverse direction to yield negative thrust. The crashback condition is dominated by the interaction of the free stream flow with the strong reverse flow. This interaction forms a highly unsteady vortex ring, which is a very prominent feature of crashback. Crashback causes highly unsteady loads and flow separation on the blade surface. The unsteady loads can cause propulsor blade damage, and also affect vehicle maneuverability. Crashback is therefore well known as one of the most challenging propeller states to analyze. This dissertation uses Large-Eddy Simulation (LES) to predict the highly unsteady flow field in crashback. A non-dissipative and robust finite volume method developed by Mahesh et al. (2004) for unstructured grids is applied to flow around marine propulsors. The LES equations are written in a rotating frame of reference. The objectives of this dissertation are: (1) to understand the flow physics of crashback in marine propulsors with and without a duct, (2) to develop a finite volume method for highly skewed meshes which usually occur in complex propulsor geometries, and (3) to develop a sliding interface method for simulations of rotor-stator propulsor on parallel platforms. LES is performed for an open propulsor in crashback and validated against experiments performed by Jessup et al. (2004). The LES results show good agreement with experiments. Effective pressures for thrust and side-force are introduced to more clearly understand the physical sources of thrust and side-force. Both thrust and side-force are seen to be mainly generated from the leading edge of the suction side of the propeller. This implies that thrust and side-force have the same source---the highly unsteady leading edge separation. Conditional averaging is performed to obtain quantitative information about the complex flow physics of high- or low-amplitude events. The events for thrust and side force show the same tendency. The conditional averages show that during high amplitude events, the vortex ring core is closer to the propeller blades, the reverse flow induced by the propeller rotation is lower, the forward flow is higher at the root of the blades, and leading and trailing edge flow separations are larger. The instantaneous flow field shows that during low amplitude events, the vortex ring is more axisymmetric and the stronger reverse flow induced by the vortex ring suppresses the forward flow so that flow separation on the blades is smaller. During high amplitude events, the vortex ring is less coherent and the weaker reverse flow cannot overcome the forward flow. The stronger forward flow makes flow separation on the blades larger. The effect of a duct on crashback is studied with LES. Thrust mostly arises from the blade surface, but most of side-force is generated from the duct surface. Both mean and RMS of pressure are much higher on inner surface of duct, especially near blade tips. This implies that side-force on the ducted propulsor is caused by the blade-duct interaction. Strong tip leakage flow is observed behind the suction side at the tip gap. The physical source of the tip leakage flow is seen to be the large pressure difference between pressure and suction sides. The conditional average for high amplitude event shows consistent results; the tip leakage flow and pressure difference are significantly higher when thrust and side-force are higher. A sliding interface method is developed to allow simulations of rotor-stator propulsor in crashback. The method allows relative rotations between different parts of the computational grid. Search algorithm for sliding elements, data structures for message passing, and accurate interpolation scheme at the sliding interface are developed for arbitrary shaped unstructured grids on parallel computing platforms. Preliminary simulations of open propulsor in crashback show reasonable performance.

The results of a wind tunnel investigation of a model rotor with a free tip

NASA Technical Reports Server (NTRS)

Stroub, Robert H.; Young, Larry A.

1985-01-01

The results of a wind-tunnel test of the free tip rotor are presented. The free tip extended over the outer 10% of the rotor blade and included a simple, passive controller mechanism. Wind-tunnel test hardware is described. The free-tip assembly, which includes the controller, functioned flawlessly throughout the test. The tip pitched freely and responded to airflow perturbation in a sharp, quick, and stable manner. Tip pitch-angle responses are presented for an advance ratio range of 0.1 to 0.397 and for a thrust coefficient range of 0.038 to 0.092. The free tip reduced power requirements, loads going into the control system, and some flatwise blade-bending moments. Chordwise loads were not reduced by the free tip.

A method to estimate weight and dimensions of aircraft gas turbine engines. Volume 1: Method of analysis

NASA Technical Reports Server (NTRS)

Pera, R. J.; Onat, E.; Klees, G. W.; Tjonneland, E.

1977-01-01

Weight and envelope dimensions of aircraft gas turbine engines are estimated within plus or minus 5% to 10% using a computer method based on correlations of component weight and design features of 29 data base engines. Rotating components are estimated by a preliminary design procedure where blade geometry, operating conditions, material properties, shaft speed, hub-tip ratio, etc., are the primary independent variables used. The development and justification of the method selected, the various methods of analysis, the use of the program, and a description of the input/output data are discussed.

Aerodynamic tip desensitization in axial flow turbines

NASA Astrophysics Data System (ADS)

Dey, Debashis

The leakage flow near the tip of unshrouded rotor blades in axial turbines imposes significant thermal loads on the blade. It is also responsible for up to a third of aerodynamic losses in a turbine stage. The leakage flow, mainly induced by the pressure differential across the rotor tip section, usually rolls into a stream-wise vertical structure near the suction side part of the blade tip. The current study uses several concepts to reduce the severity of losses introduced by the leakage vortex. Three tip desensitization techniques, both active and passive, are examined. Coolant flow from a tip trench is used to counter the momentum of the leakage jet. Next, a very short winglet obtained by slightly extending the tip platform in the tangential direction is investigated. Lastly, the widely used concept of squealer tip is studied. The current investigation is performed in the Axial Flow Turbine Research Facility (AFTRF) of the Pennsylvania State University. Rotating frame five hole probe measurements as well as stationary frame phase averaged total pressure measurements downstream of a single stage turbine facility were taken. The study enables one to draw conclusions about the nature of the flowfield in the rotor tip region. It also shows that significant efficiency gains could be obtained by using some of these techniques.

An investigation of rotor tip leakage flows in the rear-block of a multistage compressor

NASA Astrophysics Data System (ADS)

Brossman, John Richard

An effective method to improve gas turbine propulsive efficiency is to increase the bypass ratio. With fan diameter reaching a practical limit, increases in bypass ratio can be obtained from reduced core engine size. Decreasing the engine core, results in small, high pressure compressor blading, and large relative tip clearances. At general rule of 1% reduction in compressor efficiency with a 1% increase in tip clearance, a 0.66% change in SFC indicates the entire engine is sensitive to high pressure compressor tip leakage flows. Therefore, further investigations and understanding of the rotor tip leakage flows can help to improve gas turbine engine efficiency. The objectives of this research were to investigate tip leakage flows through computational modeling, examine the baseline experimental steady-stage performance, and acquire unsteady static pressure, over-the rotor to observe the tip leakage flow structure. While tip leakage flows have been investigated in the past, there have been no facilities capable of matching engine representative Reynolds number and Mach number while maintaining blade row interactions, presenting a unique and original flow field to investigate at the Purdue 3-stage axial compressor facility. To aid the design of experimental hardware and determine the influence of clearance geometry on compressor performance, a computational model of the Purdue 3-stage compressor was investigated using a steady RANS CFD analysis. A cropped rotor and casing recess design was investigated to increase the rotor tip clearance. While there were small performance differences between the geometries, the tip leakage flow field was found independent of the design therefore designing future experimental hardware around a casing recess is valid. The largest clearance with flow margin past the design point was 4% tip clearance based on the computational model. The Purdue 3-stage axial compressor facility was rebuilt and setup for high quality, detailed flow measurements during this investigation. A detailed investigation and sensitivity analysis of the inlet flow field found the influence by the inlet total temperature profile was important to performance calculations. This finding was significant and original as previous investigations have been conducted on low-speed machines where there is minimal temperature rise. The steady state performance of the baseline 1.5% tip clearance case was outlined at design speed and three off-design speeds. The leakage flow from the rear seal, the inlet flow field and a thermal boundary condition over the casing was recorded at each operating point. Stage 1 was found to be the limiting stage independent of speed. Few datasets exist on multistage compressor performance with full boundary condition definitions, especially with off-design operating points presenting this as a unique dataset for CFD comparison. The detailed unsteady pressure measurements were conducted over Rotor 1 at design and a near-stall operating condition to characterize the leakage trajectory and position. The leakage flow initial point closer to the leading edge and trajectory angle increased at the higher loading condition. The over-the-rotor static pressure field on Rotor 1 indicated similar trends between the computational model and the leakage trajectory.

Combined Amplitude and Frequency Measurements for Non-Contacting Turbomachinery Blade Vibration

NASA Technical Reports Server (NTRS)

Jagodnik, John J. (Inventor); Platt, Michael J. (Inventor)

2013-01-01

A method and apparatus for measuring the vibration of rotating blades, such as turbines, compressors, fans, or pumps, including sensing the return signal from projected energy and/or field changes from a plurality of sensors mounted on the machine housing. One or more of the sensors has a narrow field of measurement and the data is processed to provide the referenced time of arrival of each blade, and therefore the blade tip deflection due to vibration. One or more of the sensors has a wide field of measurement, providing a time history of the approaching and receding blades, and the data is processed to provide frequency content and relative magnitudes of the active mode(s) of blade vibration. By combining the overall tip deflection magnitude with the relative magnitudes of the active modes, the total vibratory stress state of the blade can be determined.

Comparison of the effects of 23-gauge and 25-gauge microincision vitrectomy blade designs on incision architecture.

PubMed

Inoue, Makoto; Abulon, Dina Joy K; Hirakata, Akito

2014-01-01

To compare the effects of different 23- and 25-gauge microincision vitrectomy trocar cannula entry systems on incision architecture. We tested one ridged microvitreoretinal (MVR), one non-ridged MVR, one pointed beveled, and one round-tipped beveled blade (n=10 per blade design per incision type). Each blade's straight and oblique incision architecture was assessed in a silicone disc simulating the sclera. Wound leakage under pressure and endoscopic observations were conducted on sclerotomy sites of isolated porcine eyes (n=4 per blade design) after simulated vitrectomy. Differences in blade design created distinct incision architecture. Incisions were linear with the ridged MVR blade, flattened "M-shaped" with the non-ridged MVR blade, asymmetrical chevron-shaped with the pointed beveled blade, and curved with the round-tipped beveled blade. With the exception of oblique entry incision thickness, both MVR blade designs created thinner incisions than the beveled blades at entry and exit sites. Only the ridged MVR blade created incisions with no leakage. Vitreous incarceration was observed with all trocar cannula systems. Wound closure in porcine eyes was similar with all blades despite differences in incision architecture. Wound leakage occurred at low to moderate infusion pressures with most blades; no wound leakage was observed with ridged MVR blades.

Preliminary Design Study of a Tail Rotor Blade Jettison Concept

DTIC Science & Technology

1978-07-01

mannar licai^infi tha hohtar or any olhar parion or torporatlon, or convaylni any ti«hti or l>armiHlon, to manuUiMura ut«, or »rti m\\ fvii «i\\(<<>l...Twist (from center of rotation to blade tip) deg - 20.0 Outboard Blade Airfoil Section m SC-1095 Tip Loss Factor - 0.97 First Flatwise Frequency...basic elements. The prototype system designed for fabrication and evaluation testing differed from a fully productionized configuration In several

Evaluation of a Microwave Blade Tip Clearance Sensor for Propulsion Health Monitoring

NASA Technical Reports Server (NTRS)

Woike, Mark R.

2013-01-01

The NASA Glenn Research Center has investigated a microwave blade tip clearance system for the structural health monitoring of gas turbine engines. This presentation describes the sensors and the experiments that have been conducted to evaluate their performance along with future plans for their use on an engine ground test.

A Visualization Study of Secondary Flows in Cascades

NASA Technical Reports Server (NTRS)

Herzig, Howard Z; Hansen, Arthur G; Costello, George R

1954-01-01

Flow-visualization techniques are employed to ascertain the streamline patterns of the nonpotential secondary flows in the boundary layers of cascades, and thereby to provide a basis for more extended analyses in turbomachines. The three-dimensional deflection of the end-wall boundary layer results in the formation of a vortex within each cascade passage. The size and tightness of the vortex generated depend upon the main-flow turning in the cascade passage. Once formed, a vortex resists turning in subsequent blade rows, with consequent unfavorable angles of attack and possible flow disturbances on the pressure surfaces of subsequent blade rows when the vortices impinge on these surfaces. Two major tip-clearance effects are observed, the formation of a tip-clearance vortex and the scraping effect of a blade with relative motion past the wall boundary layer. The flow patterns indicate methods for improving the blade tip-loading characteristics of compressors and of low- and high-speed turbulence.

High-tip-speed, low-loading transonic fan stage. Part 3: Final report

NASA Technical Reports Server (NTRS)

Ware, T. C.; Kobayashi, R. J.; Jackson, R. J.

1974-01-01

Tests were conducted on a high-tip-speed, low-loading transonic fan stage to determine the performance and inlet flow distortion tolerance of the design. The fan was designed for high efficiency at a moderate pressure ratio by designing the hub section to operate at minimum loss when the tip operates with an oblique shock. The design objective was an efficiency of 86 percent at a pressure ratio of 1.5, a specific flow (flow per unit annulus area) of 42 lb/sec-sq. ft (205.1 kgm/sec-m sq), and a tip speed of 1600 ft/sec (488.6 m/sec). During testing, a peak efficiency of 84 percent was achieved at design speed and design specific flow. At the design speed and pressure ratio, the flow was 4 percent greater than design, efficiency was 81 percent, and a stall margin of 24 percent was obtained. The stall line was improved with hub radial distortion but was reduced when the stage was tested with tip radial and circumferential flow distortions. Blade-to-blade values of static pressures were measured over the rotor blade tips.

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 5: Analysis and design of stages D and E

NASA Technical Reports Server (NTRS)

Brent, J. A.; Cheatham, J. G.; Clemmons, D. R.

1972-01-01

A conventional and a tandem bladed stage were designed for a comparative experimental evaluation in a 0.8 hub/tip ratio single-stage compressor. Based on a preliminary design study, a radially constant work input distribution was selected for the rotor designs. Velocity diagrams and blade leading and trailing edge angles selected for the conventional rotor and stator were used in the design of the tandem blading. The effects of axial velocity ratio and secondary flow on turning were included in the selection of blade leading and trailing edge angles. Design values of rotor tip velocity and stage pressure ratio were 757 ft/sec and 1.26, respectively.

16 CFR 1205.3 - Definitions.

Code of Federal Regulations, 2010 CFR

2010-01-01

...: (1) Blade means any rigid or semi-rigid device or means that is intended to cut grass during mowing operations and includes all blades of a multi-bladed mower. (2) Blade tip circle means the path described by the outermost point of the blade as it moves about its axis. (3) Crack means a visible external...

16 CFR 1205.3 - Definitions.

Code of Federal Regulations, 2012 CFR

2012-01-01

...: (1) Blade means any rigid or semi-rigid device or means that is intended to cut grass during mowing operations and includes all blades of a multi-bladed mower. (2) Blade tip circle means the path described by the outermost point of the blade as it moves about its axis. (3) Crack means a visible external...

16 CFR 1205.3 - Definitions.

Code of Federal Regulations, 2011 CFR

2011-01-01

...: (1) Blade means any rigid or semi-rigid device or means that is intended to cut grass during mowing operations and includes all blades of a multi-bladed mower. (2) Blade tip circle means the path described by the outermost point of the blade as it moves about its axis. (3) Crack means a visible external...

16 CFR § 1205.3 - Definitions.

Code of Federal Regulations, 2013 CFR

2013-01-01

...: (1) Blade means any rigid or semi-rigid device or means that is intended to cut grass during mowing operations and includes all blades of a multi-bladed mower. (2) Blade tip circle means the path described by the outermost point of the blade as it moves about its axis. (3) Crack means a visible external...

16 CFR 1205.3 - Definitions.

Code of Federal Regulations, 2014 CFR

2014-01-01

...: (1) Blade means any rigid or semi-rigid device or means that is intended to cut grass during mowing operations and includes all blades of a multi-bladed mower. (2) Blade tip circle means the path described by the outermost point of the blade as it moves about its axis. (3) Crack means a visible external...

Comparison of the effects of 23-gauge and 25-gauge microincision vitrectomy blade designs on incision architecture

PubMed Central

Inoue, Makoto; Abulon, Dina Joy K; Hirakata, Akito

2014-01-01

Purpose To compare the effects of different 23- and 25-gauge microincision vitrectomy trocar cannula entry systems on incision architecture. Methods We tested one ridged microvitreoretinal (MVR), one non-ridged MVR, one pointed beveled, and one round-tipped beveled blade (n=10 per blade design per incision type). Each blade’s straight and oblique incision architecture was assessed in a silicone disc simulating the sclera. Wound leakage under pressure and endoscopic observations were conducted on sclerotomy sites of isolated porcine eyes (n=4 per blade design) after simulated vitrectomy. Results Differences in blade design created distinct incision architecture. Incisions were linear with the ridged MVR blade, flattened “M-shaped” with the non-ridged MVR blade, asymmetrical chevron-shaped with the pointed beveled blade, and curved with the round-tipped beveled blade. With the exception of oblique entry incision thickness, both MVR blade designs created thinner incisions than the beveled blades at entry and exit sites. Only the ridged MVR blade created incisions with no leakage. Vitreous incarceration was observed with all trocar cannula systems. Conclusion Wound closure in porcine eyes was similar with all blades despite differences in incision architecture. Wound leakage occurred at low to moderate infusion pressures with most blades; no wound leakage was observed with ridged MVR blades. PMID:25429201

Test Rig for Evaluating Active Turbine Blade Tip Clearance Control Concepts

NASA Technical Reports Server (NTRS)

Lattime, Scott B.; Steinetz, Bruce M.; Robbie, Malcolm G.

2003-01-01

Improved blade tip sealing in the high pressure compressor and high pressure turbine can provide dramatic improvements in specific fuel consumption, time-on-wing, compressor stall margin and engine efficiency as well as increased payload and mission range capabilities of both military and commercial gas turbine engines. The preliminary design of a mechanically actuated active clearance control (ACC) system for turbine blade tip clearance management is presented along with the design of a bench top test rig in which the system is to be evaluated. The ACC system utilizes mechanically actuated seal carrier segments and clearance measurement feedback to provide fast and precise active clearance control throughout engine operation. The purpose of this active clearance control system is to improve upon current case cooling methods. These systems have relatively slow response and do not use clearance measurement, thereby forcing cold build clearances to set the minimum clearances at extreme operating conditions (e.g., takeoff, re-burst) and not allowing cruise clearances to be minimized due to the possibility of throttle transients (e.g., step change in altitude). The active turbine blade tip clearance control system design presented herein will be evaluated to ensure that proper response and positional accuracy is achievable under simulated high-pressure turbine conditions. The test rig will simulate proper seal carrier pressure and temperature loading as well as the magnitudes and rates of blade tip clearance changes of an actual gas turbine engine. The results of these evaluations will be presented in future works.

Heat Transfer in a Complex Trailing Edge Passage for a High Pressure Turbine Blade. Part 2:; Simulation Results

NASA Technical Reports Server (NTRS)

Rigby, David L.; Bunker, Ronald S.

2002-01-01

A combined experimental and numerical study to investigate the heat transfer distribution in a complex blade trailing edge passage was conducted. The geometry consists of a two pass serpentine passage with taper toward the trailing edge, as well as from hub to tip. The upflow channel has an average aspect ratio of roughly 14:1, while the exit passage aspect ratio is about 5:1. The upflow channel is split in an interrupted way and is smooth on the trailing edge side of the split and turbulated on the other side. A turning vane is placed near the tip of the upflow channel. Reynolds numbers in the range of 31,000 to 61,000, based on inlet conditions, were simulated numerically. The simulation was performed using the Glenn-HT code, a full three-dimensional Navier-Stokes solver using the Wilcox k-omega turbulence model. A structured multi-block grid is used with approximately 4.5 million cells and average y+ values on the order of unity. Pressure and heat transfer distributions are presented with comparison to the experimental data. While there are some regions with discrepancies, in general the agreement is very good for both pressure and heat transfer.

Tangential RIF turbine with particle removing means

DOE Office of Scientific and Technical Information (OSTI.GOV)

Linhardt, H.D.

1980-12-02

A radial inflow turbine is disclosed utilizing a star wheel type of blade and wherein the inlet nozzles are located along the sides or lateral of the blade tips rather than located radially beyond the blade tips. The side approach enables the turbine to include an annular chamber outside of the moving blades and where small particles are centrifuged during operation and collected. The incoming hot gases, that propel the blades, pass through approximately a 90/sup 0/ turn and continue radially inwardly. The large particles hit the turbine blades and are projected outwardly into the annular chamber for collection andmore » subsequent removal while the smaller particles may be centrifuged in the stream of gas and immediately thrown outwardly. The turbine is capable of removing particles to as small as 2-3 microns (..mu..m).« less

Noise from propellers with symmetrical sections at zero blade angle, II

NASA Technical Reports Server (NTRS)

Deming, A F

1938-01-01

In a previous paper (Technical Note No. 605), a theory was developed that required an empirical relation to calculate sound pressures for the higher harmonics. Further investigation indicated that the modified theory agrees with experiment and that the empirical relation was due to an interference phenomenon peculiar to the test arrangement used. Comparison is made between the test results for a two-blade arrangement and the theory. The comparison is made for sound pressures in the plane of the revolving blades for varying values of tip velocity. Comparison is also made at constant tip velocity for all values of azimuth angle B. A further check is made between the theory and the experimental results for the fundamental of a four-blade arrangement with blades of the same dimensions as those used in the two-blade arrangement.

A similitude method and the corresponding blade design of a low-speed large-scale axial compressor rotor

NASA Astrophysics Data System (ADS)

Yu, Chenghai; Ma, Ning; Wang, Kai; Du, Juan; Van den Braembussche, R. A.; Lin, Feng

2014-04-01

A similitude method to model the tip clearance flow in a high-speed compressor with a low-speed model is presented in this paper. The first step of this method is the derivation of similarity criteria for tip clearance flow, on the basis of an inviscid model of tip clearance flow. The aerodynamic parameters needed for the model design are then obtained from a numerical simulation of the target high-speed compressor rotor. According to the aerodynamic and geometric parameters of the target compressor rotor, a large-scale low-speed rotor blade is designed with an inverse blade design program. In order to validate the similitude method, the features of tip clearance flow in the low-speed model compressor are compared with the ones in the high-speed compressor at both design and small flow rate points. It is found that not only the trajectory of the tip leakage vortex but also the interface between the tip leakage flow and the incoming main flow in the high-speed compressor match well with that of its low speed model. These results validate the effectiveness of the similitude method for the tip clearance flow proposed in this paper.

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 1: Analysis and design of stages A, B, and C

NASA Technical Reports Server (NTRS)

Brent, J. A.; Cheatham, J. G.; Nilsen, A. W.

1972-01-01

A conventional rotor and stator, two dual-airfoil tandem rotors, and one dual-airfoil tandem stator were designed. The two tandem rotors were each designed with different percentages of the overall lift produced by the front airfoil. Velocity diagrams and blade leading and trailing edge metal angles selected for the conventional rotor and stator blading were used in the design of the tandem blading. Rotor inlet hub/tip ratio was 0.8. Design values of rotor tip velocity and stage pressure ratio were 757 ft/sec and 1.30, respectively.

Blade Tip Rubbing Stress Prediction

NASA Technical Reports Server (NTRS)

Davis, Gary A.; Clough, Ray C.

1991-01-01

An analytical model was constructed to predict the magnitude of stresses produced by rubbing a turbine blade against its tip seal. This model used a linearized approach to the problem, after a parametric study, found that the nonlinear effects were of insignificant magnitude. The important input parameters to the model were: the arc through which rubbing occurs, the turbine rotor speed, normal force exerted on the blade, and the rubbing coefficient of friction. Since it is not possible to exactly specify some of these parameters, values were entered into the model which bracket likely values. The form of the forcing function was another variable which was impossible to specify precisely, but the assumption of a half-sine wave with a period equal to the duration of the rub was taken as a realistic assumption. The analytical model predicted resonances between harmonics of the forcing function decomposition and known harmonics of the blade. Thus, it seemed probable that blade tip rubbing could be at least a contributor to the blade-cracking phenomenon. A full-scale, full-speed test conducted on the space shuttle main engine high pressure fuel turbopump Whirligig tester was conducted at speeds between 33,000 and 28,000 RPM to confirm analytical predictions.

Shape memory alloy adaptive control of gas turbine engine compressor blade tip clearance

NASA Astrophysics Data System (ADS)

Schetky, Lawrence M.; Steinetz, Bruce M.

1998-06-01

The ambient air ingested through the inlet of a gas turbine is first compressed by an axial compressor followed by further compression in a centrifugal compressor and then fed into the combustion chamber where ignition and expansion take place to produce the engine thrust. The axial compressor typically has five or more stages which consist of revolving blades and stators and the overall performance of the turbine is strongly affected by the compressor efficiency. When the turbine is turned on, to accommodate the rapid initial increase in the compressor blade length due to centrifugal force, the cold turbine has a built in clearance between the turbine blade tip and the casing. As the turbine reached its operating temperature there is a further increase in the blade length due to thermal expansion and, at the same time, the diameter of the casing increases. The net result is that when these various components have reached their equilibrium temperatures, the initial cold build clearance is reduced, but there remains a residual clearance. The magnitude of this clearance has a direct effect on the compressor efficiency and can be stated as: Δη/Δ CLR equals 0.5 where η is efficiency and CLR is the tip clearance. The concept of adaptive tip clearance control is based on the ability of a shape memory alloy ring to shrink to a predetermined diameter when heated to the temperature of a particular stage, and thus reducing the tip clearance. The ring is fabricated from a CuAlNi shape memory alloy and is mounted in the casing so as to be coaxial with the rotating blades of the particular stage. When cold, the ring dimensions are such as to provide the required cold build clearance, but when at operating temperature the reduced diameter creates a very small tip clearance. The clearance provided by this concept is much smaller than the clearance normally obtained for a turbine of the size being studied.

Turbine blade tip clearance measurements using skewed dual optical beams of tip timing

NASA Astrophysics Data System (ADS)

Ye, De-chao; Duan, Fa-jie; Guo, Hao-tian; Li, Yangzong; Wang, Kai

2011-12-01

Optimization and active control of the clearance between turbine blades and case of the engine is identified, especially in aerospace community, as a key technology to increase engine efficiency, reduce fuel consumption and emissions and increase service life .However, the tip clearance varies during different operating conditions. Thus a reliable non-contact and online detection system is essential and ultimately used to close the tip clearance control loop. This paper described a fiber optical clearance measuring system applying skewed dual optical beams to detect the traverse time of passing blades. Two beams were specially designed with an outward angle of 18 degree and the beam spot diameters are less than 100μm within 0-4mm working range to achieve high signal-to-noise and high sensitivity. It could be theoretically analyzed that the measuring accuracy is not compromised by degradation of signal intensity caused by any number of environmental conditions such as light source instability, contamination and blade tip imperfection. Experimental tests were undertaken to achieve a high resolution of 10µm in the rotational speed range 2000-18000RPM and a measurement accuracy of 15μm, indicating that the system is capable of providing accurate and reliable data for active clearance control (ACC).

Ducted fan inlet/exit and rotor tip flow improvements for vertical lift systems

NASA Astrophysics Data System (ADS)

Akturk, Ali

The current research utilized experimental and computational techniques in 5" and 22" diameter ducted fan test systems that have been custom designed and manufactured. Qualitative investigation of flow around the ducted fan was also performed using smoke flow visualizations. Quantitative measurements consisted of 2D and 3D velocity measurements using planar and Stereoscopic Particle Image Velocimetry (PIV and SPIV), high resolution total pressure measurements using Kiel total pressure probes and real time six-component force and torque measurements. The computational techniques used in this thesis included a recently developed radial equilibrium based rotor model(REBRM) and a three dimensional Reynolds-Averaged Navier Stokes (RANS) based CFD model. A radial equilibrium based rotor model (REBRM) developed by the author was effectively integrated into a three-dimensional RANS based computational system. The PIV measurements and computational flow predictions using (REBRM) near the fan inlet plane were in a good agreement at hover and forward flight conditions. The aerodynamic modifications resulting from the fan inlet flow distortions in forward flight regime were clearly captured in 2D PIV results. High resolution total pressure measurements at the downstream of the fan rotor showed that tip leakage, rotor hub separation, and passage flow related total pressure losses were dominant in hover condition. However, the losses were dramatically increased in forward flight because of inlet lip separation and distortion. A novel ducted fan inlet flow conditioning concept named "Double Ducted Fan" (DDF) was developed. The (DDF) concept has a potential to significantly improve the performance and controllability of VTOL UAVs and many other ducted fan based vertical lift systems. The new concept that will significantly reduce the inlet lip separation related performance penalties used a secondary stationary duct system to control "inlet lip separation" occurring especially at elevated forward flight velocities. The (DDF) is self-adjusting in a wide forward flight velocity range. DDFs corrective aerodynamic in influence becomes more pronounced with increasing flight velocity due to its inherent design properties. RANS simulations of the flow around rotor blades and duct geometry in the rotating frame of reference provided a comprehensive description of the tip leakage and passage flow in the flow environment of the two ducted fan research facilities developed throughout this thesis. The aerodynamic measurements and results of the RANS simulation showed good agreement especially near the tip region. A number of novel tip treatments based on custom designed pressure side extensions were introduced. Various tip leakage mitigation schemes were introduced by varying the chordwise location and the width of the extension in the circumferential direction. The current study showed that a proper selection of the pressure side bump location and width were the two critical parameters in influencing the success of the tip leakage mitigation approach. Significant gains in axial mean velocity component were observed when a proper pressure side tip extension was used. It is also observed that an effective tip leakage mitigation scheme significantly reduced the tangential velocity component near the tip of the axial fan blade. Reduced tip clearance related flow interactions were essential in improving the energy efficiency and range of ducted fan based vehicle. Full and inclined pressure side tip squealers were designed. Squealer tips were effective in changing the overall trajectory of the tip vortex to a higher path in radial direction. The interaction of rotor blades and tip vortex was effectively reduced and aerodynamic performance of the rotor blades was improved. The overall aerodynamic gain was a measurable reduction in leakage mass flow rate. This leakage reduction increased thrust significantly. Full and inclined pressure side tip squealers increased thrust obtained in hover condition by 9.1 % and 9.6 % respectively. A reduction or elimination of the momentum deficit in tip vortices is essential to reduce the adverse performance effects originating from the unsteady and highly turbulent tip leakage flows rotating against a stationary casing. The novel tip treatments developed throughout this thesis research are highly effective in reducing the adverse performance effects of ducted fan systems developed for VTOL vehicles. (Abstract shortened by UMI.)

Aerodynamic Inner Workings of Circumferential Grooves in a Transonic Axial Compressor

NASA Technical Reports Server (NTRS)

Hah, Chunill; Mueller, Martin; Schiffer, Heinz-Peter

2007-01-01

The current paper reports on investigations of the fundamental flow mechanisms of circumferential grooves applied to a transonic axial compressor. Experimental results show that the compressor stall margin is significantly improved with the current set of circumferential grooves. The primary focus of the current investigation is to advance understanding of basic flow mechanics behind the observed improvement of stall margin. Experimental data and numerical simulations of a circumferential groove were analyzed in detail to unlock the inner workings of the circumferential grooves in the current transonic compressor rotor. A short length scale stall inception occurs when a large flow blockage is built on the pressure side of the blade near the leading edge and incoming flow spills over to the adjacent blade passage due to this blockage. The current study reveals that a large portion of this blockage is created by the tip clearance flow originating from 20% to 50% chord of the blade from the leading edge. Tip clearance flows originating from the leading edge up to 20% chord form a tip clearance core vortex and this tip clearance core vortex travels radially inward. The tip clearance flows originating from 20% to 50% chord travels over this tip clearance core vortex and reaches to the pressure side. This part of tip clearance flow is of low momentum as it is coming from the casing boundary layer and the blade suction surface boundary layer. The circumferential grooves disturb this part of the tip clearance flow close to the casing. Consequently the buildup of the induced vortex and the blockage near the pressure side of the passage is reduced. This is the main mechanism of the circumferential grooves that delays the formation of blockage near the pressure side of the passage and delays the onset of short length scale stall inception. The primary effect of the circumferential grooves is preventing local blockage near the pressure side of the blade leading edge that directly determines flow spillage around the leading edge. The circumferential grooves do not necessarily reduce the over all blockage built up at the rotor tip section.

Multiple piece turbine rotor blade

DOEpatents

Jones, Russell B; Fedock, John A

2013-05-21

A multiple piece turbine rotor blade with a shell having an airfoil shape and secured between a spar and a platform with the spar including a tip end piece. a snap ring fits around the spar and abuts against the spar tip end piece on a top side and abuts against a shell on the bottom side so that the centrifugal loads from the shell is passed through the snap ring and into the spar and not through a tip cap dovetail slot and projection structure.

Performance with and without inlet radial distortion of a transonic fan stage designed for reduced loading in the tip region

NASA Technical Reports Server (NTRS)

Schmidt, J. F.; Ruggeri, R. S.

1978-01-01

A transonic compressor stage designed for a reduced loading in the tip region of the rotor blades was tested with and without inlet radial distortion. The rotor was 50 cm in diameter and designed for an operating tip speed of 420 m/sec. Although the rotor blade loading in the tip region was reduced to provide additional operating range, analysis of the data indicates that the flow around the damper appears to be critical and limited the stable operating range of this stage. For all levels of tip and hub radial distortion, there was a large reduction in the rotor stall margin.

Integrated Fiber-Optic Light Probe: Measurement of Static Deflections in Rotating Turbomachinery

NASA Technical Reports Server (NTRS)

Kurkov, Anatole P.

1998-01-01

At the NASA Lewis Research Center, in cooperation with Integrated Fiber Optic Systems, Inc., an integrated fiber-optic light probe system was designed, fabricated, and tested for monitoring blade tip deflections, vibrations, and to some extent, changes in the blade tip clearances of a turbomachinery fan or a compressor rotor. The system comprises a set of integrated fiber-optic light probes that are positioned to detect the passing blade tip at the leading and trailing edges. In this configuration, measurements of both nonsynchronous blade vibrations and steady-state blade deflections can be made from the timing information provided by each light probe-consisting of an integrated fiber-optic transmitting channel and numerical aperture receiving fibers, all mounted in the same cylindrical housing. With integrated fiber-optic technology, a spatial resolution of 50 mm is possible while the outer diameter is kept below 2.5 mm. To evaluate these probes, we took measurements in a single-stage compressor facility and an advanced fan rig in Lewis' 9- by 15-Foot Low-Speed Wind Tunnel.

Observations in Flight of the Region of Stalled Flow over the Blades of an Autogiro Rotor

NASA Technical Reports Server (NTRS)

Bailey, F J , Jr; Gustafon, F B

1939-01-01

The flow over the inner halves of the rotor blades on a Kellet YG-1B autogiro was investigated in flight by making camera records of the motion of silk streamers attached to the upper surfaces of the blades. These records were analyzed to determine the boundaries of the region within which the flow over the blade sections was stalled for various tip-speed ratios. For the sake of comparison, corresponding theoretical boundaries were obtained. Both the size of the stalled area and its rate of growth with increasing tip-speed ratio were found to be larger than the theory predicted, although experiment agreed with theory with regard to shape and general location of the stalled area. The stalled region may be an important factor in both the rotor lift-drag ratio and the blade flapping motion at the higher tip-speed ratios. The method of study used in this paper should be useful in further studies of the problem, including the reduction of the size of the region.

A General Multidisciplinary Turbomachinery Design Optimization system Applied to a Transonic Fan

NASA Astrophysics Data System (ADS)

Nemnem, Ahmed Mohamed Farid

The blade geometry design process is integral to the development and advancement of compressors and turbines in gas generators or aeroengines. A new airfoil section design capability has been added to an open source parametric 3D blade design tool. Curvature of the meanline is controlled using B-splines to create the airfoils. The curvature is analytically integrated to derive the angles and the meanline is obtained by integrating the angles. A smooth thickness distribution is then added to the airfoil to guarantee a smooth shape while maintaining a prescribed thickness distribution. A leading edge B-spline definition has also been implemented to achieve customized airfoil leading edges which guarantees smoothness with parametric eccentricity and droop. An automated turbomachinery design and optimization system has been created. An existing splittered transonic fan is used as a test and reference case. This design was more general than a conventional design to have access to the other design methodology. The whole mechanical and aerodynamic design loops are automated for the optimization process. The flow path and the geometrical properties of the rotor are initially created using the axi-symmetric design and analysis code (T-AXI). The main and splitter blades are parametrically designed with the created geometry builder (3DBGB) using the new added features (curvature technique). The solid model creation of the rotor sector with a periodic boundaries combining the main blade and splitter is done using MATLAB code directly connected to SolidWorks including the hub, fillets and tip clearance. A mechanical optimization is performed with DAKOTA (developed by DOE) to reduce the mass of the blades while keeping maximum stress as a constraint with a safety factor. A Genetic algorithm followed by Numerical Gradient optimization strategies are used in the mechanical optimization. The splittered transonic fan blades mass is reduced by 2.6% while constraining the maximum stress below 50% material yield strength using 2D sections thickness and chord multipliers. Once the initial design was mechanically optimized, a CFD optimization was performed to maximize efficiency and/or stall margin. The CFD grid generator (AUTOGRID) reads 3DBGB output and accounts for hub fillets and tip gaps. Single and Multi-objective Genetic Algorithm (SOGA, MOGA) optimization have been used with the CFD analysis system. In SOGA optimization, efficiency was increased by 3.525% from 78.364% to 81.889% while only changing 4 design parameters. For MOGA optimization with higher weighting efficiency than stall margin, the efficiency was increased by 2.651% from 78.364% to 81.015% while the static pressure recovery factor was increased from 0.37407 to 0.4812286 that consequently increases the stall margin. The design process starts with a hot shape design, and then a hot to cold transformation process is explained once the optimization process ends which smoothly subtracts the mechanical deflections from the hot shape. This transformation ensures an accurate tip clearance. The optimization modules can be customized by the user as one full optimization or multiple small ones. This allows the designer not to be eliminated from the design loop which helps in taking the right choice of parameters for the optimization and the final feasible design.

Effect of Blade Cutout on Power Required by Helicopters Operating at High Tip-Speed Ratios

NASA Technical Reports Server (NTRS)

Gessow, Alfred; Gustafson, F. B.

1960-01-01

A numerical study was made of the effects of blade cutout on the power required by a sample helicopter rotor traveling at tip-speed ratios of 0.3, 0.4, and 0.5. The amount of cutout varied from 0 to 0.5 of the rotor radius and the calculations were carried out for a thrust coefficient-solidity ratio of 0.04. In these calculations the blade within the cutout radius was assumed to have zero chord. The effect of such cutout on profile-drag power ranged from almost no effect at a tip-speed ratio of 0.3 to as much as a 60 percent reduction at a tip-speed ratio of 0.5. Optimum cutout was about 0.3 of the rotor radius. Part of the large power reduction at a tip-speed ratio of 0.5 resulted from a reduction in tip-region stall, brought about by cutout. For tip-speed ratios greater than 0.3, cutout also effected a significant increase in the ability of the rotor to overcome helicopter parasite drag. It is thus seen that the adverse trends (at high tip-speed ratios) indicated by the uniform-chord theoretical charts are caused in large measure by the center portion of the rotor. The extent to which a modified-design rotor can actually be made more efficient at high speeds than a uniform-chord rotor will depend in practice on the degree of success in minimizing the blade plan form near the center and on special modifications in center-section profiles. A few suggestions and estimates in regard to such modifications are included herein.

On the transonic aerodynamics of a compressor blade row

NASA Technical Reports Server (NTRS)

Erickson, J. C., Jr.; Lordi, J. A.; Rae, W. J.

1971-01-01

Linearized analyses have been carried out for the induced velocity and pressure fields within a compressor blade row operating in an infinite annulus at transonic Mach numbers of the flow relative to the blades. In addition, the relationship between the induced velocity and the shape of the mean blade surface has been determined. A computational scheme has been developed for evaluating the blade mean surface ordinates and surface pressure distributions. The separation of the effects of a specified blade thickness distribution from the effects of a specified distribution of the blade lift has been established. In this way, blade mean surface shapes that are necessary for the blades to be locally nonlifting have been computed and are presented for two examples of blades with biconvex parabolic arc sections of radially tapering thickness. Blade shapes that are required to achieve a zero thickness, uniform chordwise loading, constant work spanwise loading are also presented for two examples. In addition, corresponding surface pressure distributions are given. The flow relative to the blade tips has a high subsonic Mach number in the examples that have been computed. The results suggest that at near-sonic relative tip speeds the effective blade shape is dominated by the thickness distribution, with the lift distribution playing only a minor role.

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 3: Data and performance for stage C

NASA Technical Reports Server (NTRS)

Brent, J. A.; Clemmons, D.

1972-01-01

Stage C, comprised of tandem-airfoil rotor C and tandem-airfoil stator B, was designed and tested to establish performance data for comparison with the performance of conventional single-airfoil blading. Velocity diagrams and blade leading and trailing edge metal angles selected for the conventional rotor and stator blading were used in the design of the tandem blading. The rotor had an inlet hub/tip ratio of 0.8 and a design tip velocity of 757 ft/sec. At design equivalent rotor speed, rotor C achieved a maximum adiabatic efficiency of 91.8% at a pressure ratio of 1.31. The stage maximum adiabatic efficiency was 86.5% at a pressure ratio of 1.31.

Optimization of the blade trailing edge geometric parameters for a small scale ORC turbine

NASA Astrophysics Data System (ADS)

Zhang, L.; Zhuge, W. L.; Peng, J.; Liu, S. J.; Zhang, Y. J.

2013-12-01

In general, the method proposed by Whitfield and Baines is adopted for the turbine preliminary design. In this design procedure for the turbine blade trailing edge geometry, two assumptions (ideal gas and zero discharge swirl) and two experience values (WR and γ) are used to get the three blade trailing edge geometric parameters: relative exit flow angle β6, the exit tip radius R6t and hub radius R6h for the purpose of maximizing the rotor total-to-static isentropic efficiency. The method above is established based on the experience and results of testing using air as working fluid, so it does not provide a mathematical optimal solution to instruct the optimization of geometry parameters and consider the real gas effects of the organic, working fluid which must be taken into consideration for the ORC turbine design procedure. In this paper, a new preliminary design and optimization method is established for the purpose of reducing the exit kinetic energy loss to improve the turbine efficiency ηts, and the blade trailing edge geometric parameters for a small scale ORC turbine with working fluid R123 are optimized based on this method. The mathematical optimal solution to minimize the exit kinetic energy is deduced, which can be used to design and optimize the exit shroud/hub radius and exit blade angle. And then, the influence of blade trailing edge geometric parameters on turbine efficiency ηts are analysed and the optimal working ranges of these parameters for the equations are recommended in consideration of working fluid R123. This method is used to modify an existing ORC turbine exit kinetic energy loss from 11.7% to 7%, which indicates the effectiveness of the method. However, the internal passage loss increases from 7.9% to 9.4%, so the only way to consider the influence of geometric parameters on internal passage loss is to give the empirical ranges of these parameters, such as the recommended ranges that the value of γ is at 0.3 to 0.4, and the value of τ is at 0.5 to 0.6.

Turbomachine blade assembly

DOE Office of Scientific and Technical Information (OSTI.GOV)

Garcia Crespo, Andres Jose

Embodiments of the present disclosure include a system comprising a turbomachine blade assembly having a blade portion, a shank portion, and a mounting portion, wherein the blade portion, the shank portion, and the mounting portion comprise a first plurality of plies extending from a tip of the airfoil to a base of the dovetail.

A comparison with theory of peak to peak sound level for a model helicopter rotor generating blade slap at low tip speeds

NASA Technical Reports Server (NTRS)

Fontana, R. R.; Hubbard, J. E., Jr.

1983-01-01

Mini-tuft and smoke flow visualization techniques have been developed for the investigation of model helicopter rotor blade vortex interaction noise at low tip speeds. These techniques allow the parameters required for calculation of the blade vortex interaction noise using the Widnall/Wolf model to be determined. The measured acoustics are compared with the predicted acoustics for each test condition. Under the conditions tested it is determined that the dominating acoustic pulse results from the interaction of the blade with a vortex 1-1/4 revolutions old at an interaction angle of less than 8 deg. The Widnall/Wolf model predicts the peak sound pressure level within 3 dB for blade vortex separation distances greater than 1 semichord, but it generally over predicts the peak S.P.L. by over 10 dB for blade vortex separation distances of less than 1/4 semichord.

Structural optimization procedure of a composite wind turbine blade for reducing both material cost and blade weight

NASA Astrophysics Data System (ADS)

Hu, Weifei; Park, Dohyun; Choi, DongHoon

2013-12-01

A composite blade structure for a 2 MW horizontal axis wind turbine is optimally designed. Design requirements are simultaneously minimizing material cost and blade weight while satisfying the constraints on stress ratio, tip deflection, fatigue life and laminate layup requirements. The stress ratio and tip deflection under extreme gust loads and the fatigue life under a stochastic normal wind load are evaluated. A blade element wind load model is proposed to explain the wind pressure difference due to blade height change during rotor rotation. For fatigue life evaluation, the stress result of an implicit nonlinear dynamic analysis under a time-varying fluctuating wind is converted to the histograms of mean and amplitude of maximum stress ratio using the rainflow counting algorithm Miner's rule is employed to predict the fatigue life. After integrating and automating the whole analysis procedure an evolutionary algorithm is used to solve the discrete optimization problem.

Comparison of tip apex distance and cut-out complications between helical blades and lag screws in intertrochanteric fractures among the elderly: a meta-analysis.

PubMed

Li, Shuang; Chang, Shi-Min; Niu, Wen-Xin; Ma, Hui

2015-11-01

To investigate whether helical blade implant systems have advantages in terms of tip apex distance (TAD) and cut-out rate in comparison to conventional lag screws for intertrochanteric fractures in a geriatric population. Methods: Relevant articles were sourced from the MEDLINE, Embase, Ovid and Cochrane Library databases from inception through March 2015. All randomized controlled trials (RCTs) comparing outcomes between helical blade and lag screw implant systems were selected. Mean TAD values and reported cut-out complications were noted. Each author independently assessed the relevance of the enrolled studies and the quality of the extracted data. Data were analyzed using R software. Ten studies including 1831 patients were eligible for this review, seven of which were included in a combined analysis of dichotomous outcomes and five in a combined analysis of continuous outcomes. The results revealed that, compared with lag screw implantations, the use of helical blades led to a lower rate of cut-out complications (95 % CI: 0.28–0.96, P = 0.036). Patients who experienced cut-out complications had a significantly greater tip apex distance (95 % CI: 0.68–1.34, P < 0.001). However, the actual tip apex distances were similar between the screw group and blade group (95 % CI: −0.44–0.79, P = 0.58). No difference in TAD values was found between blades and screws. In addition, the cut-out risk in the blade-design group was lower than that of the screw group. Therefore, TAD is not an accurate predictor of cut-out risk.

Multiple pure tone noise generated by fans at supersonic tip speeds

NASA Technical Reports Server (NTRS)

Sofrin, T. G.; Pickett, G. F.

1974-01-01

The existence of clusters of pure tones at integral multiples of shaft speed has been noted for supersonic-tip-speed operation of fans and compressors. A continuing program to explore this phenomenon, often called combination-tone noise, has been in effect for several years. This paper reviews the research program, which involves a wide range of engines, compressor rigs, and special apparatus. Elements of the aerodynamics of the blade-associated shock waves are outlined and causes of blade-to-blade shock inequalities, responsible for the multiple tones, are described.

Experimental investigation of the flow field and power consumption characteristics of regular and fractal blade impellers in a dynamic mixer

NASA Astrophysics Data System (ADS)

Steiros, K.; Bruce, P. J. K.; Buxton, O. R. H.; Vassilicos, J. C.

2015-11-01

Experiments have been performed in an octagonal un-baffled water tank, stirred by three radial turbines with different geometry impellers: (1) regular rectangular blades; (2) single-iteration fractal blades; (3) two-iteration fractal blades. Shaft torque was monitored and the power number calculated for each case. Both impellers with fractal geometry blades exhibited a decrease of turbine power number compared to the regular one (15% decrease for single-iteration and 19% for two iterations). Phase locked PIV in the discharge region of the blades revealed that the vortices emanating from the regular blades are more coherent, have higher kinetic energy, and advect faster towards the tank's walls where they are dissipated, compared to their fractal counterparts. This suggests a strong link between vortex production and behaviour and the energy input for the different impellers. Planar PIV measurements in the bulk of the tank showed an increase of turbulence intensity of over 20% for the fractal geometry blades, suggesting higher mixing efficiency. Experiments with pressure measurements on the different geometry blade surfaces are ongoing to investigate the distribution of forces, and calculate hydrodynamic centres of pressure. The authors would like to acknowledge the financial support given by European Union FP7 Marie Curie MULTISOLVE project (Grant Agreement No. 317269).

Fiber-optic laser Doppler turbine tip clearance probe

NASA Astrophysics Data System (ADS)

Büttner, Lars; Pfister, Thorsten; Czarske, Jürgen

2006-05-01

A laser Doppler based method for in situ single blade tip clearance measurements of turbomachines with high precision is presented for what we believe is the first time. The sensor is based on two superposed fanlike interference fringe systems generated by two laser wavelengths from a fiber-coupled, passive, and therefore compact measurement head employing diffractive optics. Tip clearance measurements at a transonic centrifugal compressor performed during operation at 50,000 rpm (833 Hz, 586 m/s tip speed) are reported. At these speeds the measured uncertainty of the tip position was less than 20 μm, a factor of 2 more accurate than that of capacitive probes. The sensor offers great potential for in situ and online high-precision tip clearance measurements of metallic and nonmetallic turbine blades.

Fiber-optic laser Doppler turbine tip clearance probe.

PubMed

Büttner, Lars; Pfister, Thorsten; Czarske, Jürgen

2006-05-01

A laser Doppler based method for in situ single blade tip clearance measurements of turbomachines with high precision is presented for what we believe is the first time. The sensor is based on two superposed fanlike interference fringe systems generated by two laser wavelengths from a fiber-coupled, passive, and therefore compact measurement head employing diffractive optics. Tip clearance measurements at a transonic centrifugal compressor performed during operation at 50,000 rpm (833 Hz, 586 m/s tip speed) are reported. At these speeds the measured uncertainty of the tip position was less than 20 microm, a factor of 2 more accurate than that of capacitive probes. The sensor offers great potential for in situ and online high-precision tip clearance measurements of metallic and nonmetallic turbine blades.

High-Tip-Speed, Low-Loading Transonic Fan Stage. Part 1: Aerodynamic and Mechanical Design

NASA Technical Reports Server (NTRS)

Wright, L. C.; Vitale, N. G.; Ware, T. C.; Erwin, J. R.

1973-01-01

A high-tip-speed, low-loading transonic fan stage was designed to deliver an overall pressure ratio of 1.5 with an adiabatic efficiency of 86 percent. The design flow per unit annulus area is 42.0 pounds per square foot. The fan features a hub/tip ratio of 0.46, a tip diameter of 28.74 in. and operates at a design tip speed of 1600 fps. For these design conditions, the rotor blade tip region operates with supersonic inlet and supersonic discharge relative velocities. A sophisticated quasi-three-dimensional characteristic section design procedure was used for the all-supersonic sections and the inlet of the midspan transonic sections. For regions where the relative outlet velocities are supersonic, the blade operates with weak oblique shocks only.

Noise comparison of two 1.2-pressure-ratio fans with 15 and 42 rotor blades

NASA Technical Reports Server (NTRS)

Woodward, R. P.; Glaser, F. W.; Wazyniak, J. A.

1973-01-01

Two 1.829-m-(6-ft-) diameter fans suitable for a quiet engine for future short-takeoff-and-landing (STOL) aircraft were compared. Both fans were designed for a 1.2 pressure ratio with similar weight flows, thrusts, and tip speeds. The first fan, designated QF-9, had 15 rotor blades and 11 stator blades. The rotor was highly loaded and the tip solidity was less than 1. The QF-9 rotor blades had an adjustable pitch feature which can be used for thrust reversal. The second fan, designated QF-6, operated at a moderate loading with a rotor tip solidity greater than 1. Fan QF-6 had 42 rotor blades and 50 stator blades. The low number of rotor blades for QF-9 reduced the frequency of the blade-passage tone below the range of maximum annoyance. In addition to this difference, the QF-9 fan had a somewhat smaller rotor-stator separation than the QF-6 fan. In terms of sound pressure level and sound power level, QF-9 was the noisier fan, with the power level results for QF-9 being about 1 db above those for QF-6 at equivalent operating points as determined by similar stage pressure ratios. At the same equivalent operating points, the maximum perceived noise along a 152.5-m (500-ft) sideline for QF-9 was about 2.5 PNdb below that for QF-6, which indicated that QF-9 was less objectionable to human hearing.

Improving the performances of H-Darrieus cross-flow turbines through proper detached end plate designs

NASA Astrophysics Data System (ADS)

Villeneuve, Thierry; Boudreau, Matthieu; Dumas, Guy; CFD Laboratory LMFN Team

2017-11-01

Previous studies on H-Darrieus cross-flow turbines have highlighted the fact that their performances are highly sensitive to the detrimental effects associated with the blades tips. Wingtip devices could be designed in order to attenuate these effects, but the benefits of such devices are always impaired by their added viscous drag since they are moving with the turbine's blades. In this context, the development of fixed and detached end plates, i.e., which are not in contact with the turbine's blades, could reduce the tip losses without the undesirable added drag of typical wingtip devices moving with the blades. The case of a single stationary blade with detached end plates has first been investigated with RANS simulations in order to understand the mechanisms responsible for the increase of the blade's lift. An analysis of the vorticity lines' dynamics provides crucial insights into the effects of the gap width between the blade and the detached end plate on the blade's loading and on the intensity of the tip vortices. Based on these observations, various configurations of detached end plates are tested on cross-flow turbines via RANS and DDES simulations. Preliminary results show that appropriate detached end plates can significantly increase the turbines' efficiency. The authors gratefully acknowledge the Natural Sciences and Engineering Research Council of Canada (NSERC) for their financial support as well as Compute Canada and Calcul Québec for their supercomputer allocations.

Analysis of the behavior of a wiper blade around the reversal in consideration of dynamic and static friction

NASA Astrophysics Data System (ADS)

Unno, M.; Shibata, A.; Yabuno, H.; Yanagisawa, D.; Nakano, T.

2017-04-01

Reducing noise generated by automobile windshield wipers during reversals is a desirable feature. For this purpose, details of the behavior of the wiper blade need to be ascertained. In this study, we present theoretical and experimental clarification of this behavior during reversals. Using simulation algorithms to consider exactly the effects of dynamic and static friction, we determined theoretical predictions for the vibrational response caused by friction and the response frequency and compared these results with experimental ones obtained from a mock-up incorporating an actual wiper blade. We introduce an analytical link model with two degrees of freedom and consider two types of states at the blade tip. In the stick and the slip states, static friction and dynamic friction, respectively, act on the blade tip. In the theoretical approach, the static friction is expressed by a set-valued function. The transition between the two states is repeated and an evaluation of an exact transition time leads to an accurate prediction of the behavior of the wiper system. In the analysis, the slack variable method is used to find the exact transition time. Assuming low blade speeds during reversal, a parameter study indicates that the blade tip transitions between slip and stick states and the frequency of the vibration caused by this transitions is close to the natural frequency of the neck of the wiper blade. The theoretical predictions are in good agreement with experimental observations.

Techniques for blade tip clearance measurements with capacitive probes

NASA Astrophysics Data System (ADS)

Steiner, Alexander

2000-07-01

This article presents a proven but advantageous concept for blade tip clearance evaluation in turbomachinery. The system is based on heavy duty probes and a high frequency (HF) and amplifying electronic unit followed by a signal processing unit. Measurements are taken under high temperature and other severe conditions such as ionization. Every single blade can be observed. The signals are digitally filtered and linearized in real time. The electronic set-up is highly integrated. Miniaturized versions of the electronic units exist. The small and robust units can be used in turbo engines in flight. With several probes at different angles in one radial plane further information is available. Shaft eccentricity or blade oscillations can be calculated.

Correlation of SA349/2 helicopter flight-test data with a comprehensive rotorcraft model

NASA Technical Reports Server (NTRS)

Yamauchi, Gloria K.; Heffernan, Ruth M.; Gaubert, Michel

1986-01-01

A comprehensive rotorcraft analysis model was used to predict blade aerodynamic and structural loads for comparison with flight test data. The data were obtained from an SA349/2 helicopter with an advanced geometry rotor. Sensitivity of the correlation to wake geometry, blade dynamics, and blade aerodynamic effects was investigated. Blade chordwise pressure coefficients were predicted for the blade transonic regimes using the model coupled with two finite-difference codes.

Noninterference Systems Developed for Measuring and Monitoring Rotor Blade Vibrations

NASA Technical Reports Server (NTRS)

Kurkov, Anatole P.

2003-01-01

In the noninterference measurement of blade vibrations, a laser light beam is transmitted to the rotor blade tips through a single optical fiber, and the reflected light from the blade tips is collected by a receiving fiber-optic bundle and conducted to a photodetector. Transmitting and receiving fibers are integrated in an optical probe that is enclosed in a metal tube which also houses a miniature lens that focuses light on the blade tips. Vibratory blade amplitudes can be deduced from the measurement of the instantaneous time of arrival of the blades and the knowledge of the rotor speed. The in-house noninterference blade-vibration measurement system was developed in response to requirements to monitor blade vibrations in several tests where conventional strain gauges could not be installed or where there was a need to back up strain gauges should critical gauges fail during the test. These types of measurements are also performed in the aircraft engine industry using proprietary in-house technology. Two methods of measurement were developed for vibrations that are synchronous with a rotor shaft. One method requires only one sensor; however, it is necessary to continuously record the data while the rotor is being swept through the resonance. In the other method, typically four sensors are employed and the vibratory amplitude is deduced from the data by performing a least square fit to a harmonic function. This method does not require continuous recording of data through the resonance and, therefore, is better suited for monitoring. The single-probe method was tested in the Carl facility at the Wright- Patterson Air Force Base, and the multiple-probe method was tested in NASA Glenn Research Center's Spin Rig facility, which uses permanent magnets to excite synchronous vibrations. Representative results from this test are illustrated in the bar chart. Nonsynchronous vibrations were measured online during testing of the Quiet High Speed Fan in Glenn s 9- by 15-Foot Low-Speed Wind Tunnel. Three sensors were employed, enabling a reconstruction of the vibratory patterns at the leading and trailing edges at the tip span, as well as a determination of vibratory amplitudes for every blade.

Experimental Investigation of an Air-Cooled Turbine Operating in a Turbojet Engine at Turbine Inlet Temperatures up to 2500 F

NASA Technical Reports Server (NTRS)

Cochran, Reeves P.; Dengler, Robert P.

1961-01-01

An experimental investigation was made of an air-cooled turbine at average turbine inlet temperatures up to 2500 F. A modified production-model 12-stage axial-flow-compressor turbojet engine operating in a static sea-level stand was used as the test vehicle. The modifications to the engine consisted of the substitution of special combustor and turbine assemblies and double-walled exhaust ducting for the standard parts of the engine. All of these special parts were air-cooled to withstand the high operating temperatures of the investigation. The air-cooled turbine stator and rotor blades were of the corrugated-insert type. Leading-edge tip caps were installed on the rotor blades to improve leading-edge cooling by diverting the discharge of coolant to regions of lower gas pressure toward the trailing edge of the blade tip. Caps varying in length from 0.15- to 0.55-chord length were used in an attempt to determine the optimum cap length for this blade. The engine was operated over a range of average turbine inlet temperatures from about 1600 to about 2500 F, and a range of average coolant-flow ratios of 0.012 to 0.065. Temperatures of the air-cooled turbine rotor blades were measured at all test conditions by the use of thermocouples and temperature-indicating paints. The results of the investigation indicated that this type of blade is feasible for operation in turbojet engines at the average turbine inlet temperatures and stress levels tested(maximums of 2500 F and 24,000 psi, respectively). An average one-third-span blade temperature of 1300 F could be maintained on 0.35-chord tip cap blades with an average coolant-flow ratio of about 0.022 when the average turbine inlet temperature was 2500 F and cooling-air temperature was about 260 F. All of the leading-edge tip cap lengths improved the cooling of the leading-edge region of the blades, particularly at low average coolant-flow ratios. At high gas temperatures, such parts as the turbine stator and the combustor liners are likely to be as critical as the turbine rotor blades.

Single stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 4: Data and performance for stage B

NASA Technical Reports Server (NTRS)

Brent, J. A.; Cheatham, J. G.

1973-01-01

Stage B, composed of tandem-airfoil rotor B and stator B, was tested with uniform inlet flow and with hub radial, tip radial and 90 degree one-per-revolution circumferential distortion of the inlet flow as part of an overall program to evaluate the effectiveness of tandem airfoils for increasing the design point loading capability and stable operating range of rotor and stator blading. The results of this series of tests provide overall performance and blade element data for evaluating: (1) the potential of tandem blading for extending the loading limit and stable operating range of a stage representative of a middle stage of an advanced high pressure compressor, (2) the effect of loading split between the two airfoils in tandem on the performance of tandem blading, and (3) the effects of inlet flow distortion on the stage performance. The rotor had an inlet hub/tip ratio of 0.8 and a design tip velocity of 757 ft/sec. With uniform inlet flow, rotor B achieved a maximum adiabatic efficiency of 88.4% at design equivalent rotor speed and a pressure ratio of 1.31. The stage maximum adiabatic efficiency at design equivalent rotor speed with uniform inlet flow was 82.5% at a pressure ratio of 1.28. Tip radial and circumferential distortion of the inlet flow caused substantial reductions in surge margin.

Optical and thermal performance of bladed receivers

NASA Astrophysics Data System (ADS)

Pye, John; Coventry, Joe; Ho, Clifford; Yellowhair, Julius; Nock, Ian; Wang, Ye; Abbasi, Ehsan; Christian, Joshua; Ortega, Jesus; Hughes, Graham

2017-06-01

Bladed receivers use conventional receiver tube-banks rearranged into bladed/finned structures, and offer better light trapping, reduced radiative and convective losses, and reduced tube mass, based on the presented optical and thermal analysis. Optimising for optical performance, deep blades emerge. Considering thermal losses leads to shallower blades. Horizontal blades perform better, in both windy and no-wind conditions, than vertical blades, at the scales considered so far. Air curtains offer options to further reduce convective losses; high flux on blade-tips is still a concern.

Time-of-Flight Tip-Clearance Measurements

NASA Technical Reports Server (NTRS)

Dhadwal, H. S.; Kurkov, A. P.; Janetzke, D. C.

1999-01-01

In this paper a time-of-flight probe system incorporating the two integrated fiber optic probes which are tilted equally relative to the probe holder centerline, is applied for the first time to measure the tip clearance of an advanced fan prototype. Tip clearance is largely independent of the signal amplitude and it relies on timing measurement. This work exposes optical effects associated with the fan blade stagger angle that were absent during the original spin-rig experiment on the zero stagger rotor. Individual blade tip clearances were measured with accuracy of +/- 127-mm (+/- 0.005-in). Probe features are discussed and improvements to the design are suggested.

CERT: Center of Excellence in Rotorcraft Technology

NASA Technical Reports Server (NTRS)

2002-01-01

The research objectives of this effort are to understand the physical processes that influence the formation of the tip vortex of a rotor in advancing flight, and to develop active and passive means of weakening the tip vortex during conditions when strong blade-vortex-interaction effects are expected. A combined experimental, analytical, and computational effort is being employed. Specifically, the following efforts are being pursued: 1. Analytical evaluation and design of combined elastic tailoring and active material actuators applicable to rotor blade tips. 2. Numerical simulations of active and passive tip devices. 3. LDV Measurement of the near and far wake behind rotors in forward flight.

Reynolds-averaged Navier-Stokes computation on tip clearance flow in a compressor cascade using an unstructured grid

NASA Astrophysics Data System (ADS)

Shin, Sangmook

2001-07-01

A three-dimensional unstructured incompressible RANS code has been developed using artificial compressibility and Spalart-Allmaras eddy viscosity model. A node-based finite volume method is used in which all flow variables are defined at the vertices of tetrahedrons in an unstructured grid. The inviscid fluxes are computed by using the Roe's flux difference splitting method, and higher order accuracy is attained by data reconstruction based on Taylor series expansion. Gauss theorem is used to formulate necessary gradients. For time integration, an implicit scheme based on linearized Euler backward method is used. A tetrahedral unstructured grid generation code has been also developed and applied to the tip clearance flow in a highly staggered cascade. Surface grids are first generated in the flow passage and blade tip by using several triangulation methods including Delaunay triangulation, advancing front method and advancing layer method. Then the whole computational domain including tip gap region is filled with prisms using the surface grids. The code has been validated by comparisons with available computational and experimental results for several test cases: inviscid flow around NACA section, laminar and turbulent flow over a flat plate, turbulent flow through double-circular arc cascade and laminar flow through a square duct with 90° bend. Finally the code is applied to a linear cascade that has GE rotor B section with tip clearance and a high stagger angle of 56.9°. The overall structure of the tip clearance flow is well predicted. Loss of loading due to tip leakage flow and reloading due to tip leakage vortex are presented. On the end wall, separation line of the tip leakage vortex and reattachment line of passage vortex are identified. Prediction of such an interaction presents a challenge to RANS computations. The effects of blade span on the flow structure have been also investigated. Two cascades with blades of aspect ratios of 0.5 and 1.0 are considered. By comparing pressure distributions on the blade, it is shown that the aspect ratio has strong effects on loading distribution on the blade although the tip gap height is very small (0.016 chord). Grid convergence study has been carried out with three different grids for pressure distributions and limiting streamlines on the end wall. (Abstract shortened by UMI.)

A Model Rotor in Axial Flight

NASA Technical Reports Server (NTRS)

McAlister, K. W.; Huang, S. S.; Abrego, A. I.

2001-01-01

A model rotor was mounted horizontally in the settling chamber of a wind tunnel to obtain performance and wake structure data under low climb conditions. The immediate wake of the rotor was carefully surveyed using 3-component particle image velocimetry to define the velocity and vortical content of the flow, and used in a subsequent study to validate a theory for the separate determination of induced and profile drag. Measurements were obtained for two collective pitch angles intended to render a predominately induced drag state and another with a marked increase in profile drag. A majority of the azimuthally directed vorticity in the wake was found to be concentrated in the tip vortices. However, adjacent layers of inboard vorticity with opposite sense were clearly present. At low collective, the close proximity of the tip vortex from the previous blade caused the wake from the most recent blade passage to be distorted. The deficit velocity component that was directed along the azimuth of the rotor blade was never more that 15 percent of the rotor tip speed, and except for the region of the tip vortex, appeared to have totally disappeared form the wake left by the previous blade.

Investigation of Rotor-Airframe Interaction Noise Associated with Small-Scale Rotary-Wing Unmanned Aircraft Systems

NASA Technical Reports Server (NTRS)

Zawodny, Nikolas S.; Boyd, D. Douglas, Jr.

2017-01-01

In this study, hover acoustic measurements are taken on isolated rotor-airframe configurations representative of smallscale, rotary-wing unmanned aircraft systems (UAS). Each rotor-airframe configuration consists of two fixed-pitch blades powered by a brushless motor, with a simplified airframe geometry intended to represent a generic multicopter arm. In addition to acoustic measurements, CFD-based aeroacoustic predictions are implemented on a subset of the experimentally tested rotor-airframe configurations in an effort to better understand the noise content of the rotor-airframe systems. Favorable agreements are obtained between acoustic measurements and predictions, based on both time- and frequency-domain post-processing techniques. Results indicate that close proximity of airframe surfaces result in the generation of considerable tonal acoustic content in the form of harmonics of the rotor blade passage frequency (BPF). Analysis of the acoustic prediction data shows that the presence of the airframe surfaces can generate noise levels either comparable to or greater than the rotor blade surfaces under certain rotor tip clearance conditions. Analysis of the on-surface Ffowcs Williams and Hawkings (FW-H) source terms provide insight as to the predicted physical noise-generating mechanisms on the rotor and airframe surfaces.

Measurement and prediction of model-rotor flow fields

NASA Technical Reports Server (NTRS)

Owen, F. K.; Tauber, M. E.

1985-01-01

This paper shows that a laser velocimeter can be used to measure accurately the three-component velocities induced by a model rotor at transonic tip speeds. The measurements, which were made at Mach numbers from 0.85 to 0.95 and at zero advance ratio, yielded high-resolution, orthogonal velocity values. The measured velocities were used to check the ability of the ROT22 full-potential rotor code to predict accurately the transonic flow field in the crucial region around and beyond the tip of a high-speed rotor blade. The good agreement between the calculated and measured velocities established the code's ability to predict the off-blade flow field at transonic tip speeds. This supplements previous comparisons in which surface pressures were shown to be well predicted on two different tips at advance ratios to 0.45, especially at the critical 90 deg azimuthal blade position. These results demonstrate that the ROT22 code can be used with confidence to predict the important tip-region flow field, including the occurrence, strength, and location of shock waves causing high drag and noise.

Three dimensional mean velocity and turbulence characteristics in the annulus wall region of an axial flow compressor rotor passage

NASA Technical Reports Server (NTRS)

Davino, R.; Lakshminarayana, B.

1982-01-01

The experiment was performed using the rotating hot-wire technique within the rotor blade passage and the stationary hot-wire technique for the exitflow of the rotor blade passage. The measurements reveal the effect of rotation and subsequent flow interactions upon the rotor blade flowfield and wake development in the annulus-wall region. The flow near the rotor blade tips is found to be highly complex due to the interaction of the annulus-wall boundary layer, the blade boundary layers, the tip leakage flow, and the secondary flow. Within the blade passage, this interaction results in an appreciable radial inward flow as well as a defect in the mainstream velocity near the mid-passage. Turbulence levels within this region are very high. This indicates a considerable extent of flow mixing due to the viscous flow interactions. The size and strength of this loss core is found to grow with axial distance from the blade trailing edge. The nature of the rotor blade exit-flow was dominated by the wake development.

Investigation of helicopter rotor blade/wake interactive impulsive noise

NASA Technical Reports Server (NTRS)

Miley, S. J.; Hall, G. F.; Vonlavante, E.

1987-01-01

An analysis of the Tip Aerodynamic/Aeroacoustic Test (TAAT) data was performed to identify possible aerodynamic sources of blade/vortex interaction (BVI) impulsive noise. The identification is based on correlation of measured blade pressure time histories with predicted blade/vortex intersections for the flight condition(s) where impulsive noise was detected. Due to the location of the recording microphones, only noise signatures associated with the advancing blade were available, and the analysis was accordingly restricted to the first and second azimuthal quadrants. The results show that the blade tip region is operating transonically in the azimuthal range where previous BVI experiments indicated the impulsive noise to be. No individual blade/vortex encounter is identifiable in the pressure data; however, there is indication of multiple intersections in the roll-up region which could be the origin of the noise. Discrete blade/vortex encounters are indicated in the second quadrant; however, if impulsive noise were produced here, the directivity pattern would be such that it was not recorded by the microphones. It is demonstrated that the TAAT data base is a valuable resource in the investigation of rotor aerodynamic/aeroacoustic behavior.

Detailed noise measurements on the SR-7A propeller: Tone behavior with helical tip Mach number

NASA Technical Reports Server (NTRS)

Dittmar, James H.; Hall, David G.

1991-01-01

Detailed noise measurements were taken on the SR-7A propeller to investigate the behavior of the noise with helical tip Mach number and then to level off as Mach number was increased further. This behavior was further investigated by obtaining detailed pressure-time histories of data. The pressure-time histories indicate that a portion of the primary pressure pulse is progressively cancelled by a secondary pulse which results in the noise leveling off as the helical tip Mach number is increased. This second pulse appears to originate on the same blade as the primary pulse and is in some way connected to the blade itself. This leaves open the possibility of redesigning the blade to improve the cancellation; thereby, the propeller noise is reduced.

Detailed noise measurements on the SR-7A propeller: Tone behavior with helical tip Mach number

NASA Astrophysics Data System (ADS)

Dittmar, James H.; Hall, David G.

1991-12-01

Detailed noise measurements were taken on the SR-7A propeller to investigate the behavior of the noise with helical tip Mach number and then to level off as Mach number was increased further. This behavior was further investigated by obtaining detailed pressure-time histories of data. The pressure-time histories indicate that a portion of the primary pressure pulse is progressively cancelled by a secondary pulse which results in the noise leveling off as the helical tip Mach number is increased. This second pulse appears to originate on the same blade as the primary pulse and is in some way connected to the blade itself. This leaves open the possibility of redesigning the blade to improve the cancellation; thereby, the propeller noise is reduced.

Rub tolerance evaluation of two sintered NiCrAl gas path seal materials. [wear tests of gas turbine engine seals

NASA Technical Reports Server (NTRS)

Bill, R. C.

1978-01-01

Two strength level variations of sintered NiCrAl (about 40 percent dense), candidate high pressure turbine seal materials, were subject to rub tolerance testing against AM 355 steel blade tips. The high strength material (17 N/sq mm tensile strength) showed frictional and radial loads that were 20 to 50 percent higher than those measured for the low strength material (15.5 N/ sq mm tensile strength). Measured wear to the AM 355 blade tips was not significantly different for the two sintered NiCrAl seal materials. Wear of the sintered NiCrAl was characterized by material removal to a depth greater than the depth to which blade tips were driven into the seal, indicating self-erosion effects.

Blade design and analysis using a modified Euler solver

NASA Technical Reports Server (NTRS)

Leonard, O.; Vandenbraembussche, R. A.

1991-01-01

An iterative method for blade design based on Euler solver and described in an earlier paper is used to design compressor and turbine blades providing shock free transonic flows. The method shows a rapid convergence, and indicates how much the flow is sensitive to small modifications of the blade geometry, that the classical iterative use of analysis methods might not be able to define. The relationship between the required Mach number distribution and the resulting geometry is discussed. Examples show how geometrical constraints imposed upon the blade shape can be respected by using free geometrical parameters or by relaxing the required Mach number distribution. The same code is used both for the design of the required geometry and for the off-design calculations. Examples illustrate the difficulty of designing blade shapes with optimal performance also outside of the design point.

Cut-cell method based large-eddy simulation of tip-leakage flow

NASA Astrophysics Data System (ADS)

Pogorelov, Alexej; Meinke, Matthias; Schröder, Wolfgang

2015-07-01

The turbulent low Mach number flow through an axial fan at a Reynolds number of 9.36 × 105 based on the outer casing diameter is investigated by large-eddy simulation. A finite-volume flow solver in an unstructured hierarchical Cartesian setup for the compressible Navier-Stokes equations is used. To account for sharp edges, a fully conservative cut-cell approach is applied. A newly developed rotational periodic boundary condition for Cartesian meshes is introduced such that the simulations are performed just for a 72° segment, i.e., the flow field over one out of five axial blades is resolved. The focus of this numerical analysis is on the development of the vortical flow structures in the tip-gap region. A detailed grid convergence study is performed on four computational grids with 50 × 106, 250 × 106, 1 × 109, and 1.6 × 109 cells. Results of the instantaneous and the mean fan flow field are thoroughly analyzed based on the solution with 1 × 109 cells. High levels of turbulent kinetic energy and pressure fluctuations are generated by a tip-gap vortex upstream of the blade, the separating vortices inside the tip gap, and a counter-rotating vortex on the outer casing wall. An intermittent interaction of the turbulent wake, generated by the tip-gap vortex, with the downstream blade, leads to a cyclic transition with high pressure fluctuations on the suction side of the blade and a decay of the tip-gap vortex. The disturbance of the tip-gap vortex results in an unsteady behavior of the turbulent wake causing the intermittent interaction. For this interaction and the cyclic transition, two dominant frequencies are identified which perfectly match with the characteristic frequencies in the experimental sound power level and therefore explain their physical origin.

Study of the capacitance technique for measuring high-temperature blade tip clearance on ceramic rotors

NASA Technical Reports Server (NTRS)

Barranger, John P.

1993-01-01

Higher operating temperatures required for increased engine efficiency can be achieved by using ceramic materials for engine components. Ceramic turbine rotors are subject to the same limitations with regard to gas path efficiency as their superalloy predecessors. In this study, a modified frequency-modulation system is proposed for the measurement of blade tip clearance on ceramic rotors. It is expected to operate up to 1370 C (2500 F), the working temperature of present engines with ceramic turbine rotors. The design of the system addresses two special problems associated with nonmetallic blades: the capacitance is less than that of a metal blade and the effects of temperature may introduce uncertainty with regard to the blade tip material composition. To increase capacitance and stabilize the measurement, a small portion of the rotor is modified by the application of 5-micron-thick platinum films. The platinum surfaces on the probe electrodes and rotor that are exposed to the high-velocity gas stream are coated with an additional 10-micron-thick protective ceramic topcoat. A finite-element method is applied to calculate the capacitance as a function of clearance.

Validation of the actuator disc and actuator line techniques for yawed rotor flows using the New MEXICO experimental data

NASA Astrophysics Data System (ADS)

Breton, S.-P.; Shen, W. Z.; Ivanell, S.

2017-05-01

Experimental data acquired in the New MEXICO experiment on a yawed 4.5m diameter rotor model turbine are used here to validate the actuator line (AL) and actuator disc (AD) models implemented in the Large Eddy Simulation code EllipSys3D in terms of loading and velocity field. Even without modelling the geometry of the hub and nacelle, the AL and AD models produce similar results that are generally in good agreement with the experimental data under the various configurations considered. As expected, the AL model does better at capturing the induction effects from the individual blade tip vortices, while the AD model can reproduce the averaged features of the flow. The importance of using high quality airfoil data (including 3D corrections) as well as a fine grid resolution is highlighted by the results obtained. Overall, it is found that both models can satisfactorily predict the 3D velocity field and blade loading of the New MEXICO rotor under yawed inflow.

Capacitance-type blade-tip clearance measurement system using a dual amplifier with ramp/dc inputs and integration

NASA Technical Reports Server (NTRS)

Sarma, Garimella R.; Barranger, John P.

1992-01-01

The analysis and prototype results of a dual-amplifier circuit for measuring blade-tip clearance in turbine engines are presented. The capacitance between the blade tip and mounted capacitance electrode within a guard ring of a probe forms one of the feedback elements of an operational amplifier (op amp). The differential equation governing the circuit taking into consideration the nonideal features of the op amp was formulated and solved for two types of inputs (ramp and dc) that are of interest for the application. Under certain time-dependent constraints, it is shown that (1) with a ramp input the circuit has an output voltage proportional to the static tip clearance capacitance, and (2) with a dc input, the output is proportional to the derivative of the clearance capacitance, and subsequent integration recovers the dynamic capacitance. The technique accommodates long cable lengths and environmentally induced changes in cable and probe parameters. System implementation for both static and dynamic measurements having the same high sensitivity is also presented.

Capacitance-type blade-tip clearance measurement system using a dual amplifier with ramp/dc inputs and integration

NASA Astrophysics Data System (ADS)

Sarma, Garimella R.; Barranger, John P.

1992-10-01

The analysis and prototype results of a dual-amplifier circuit for measuring blade-tip clearance in turbine engines are presented. The capacitance between the blade tip and mounted capacitance electrode within a guard ring of a probe forms one of the feedback elements of an operational amplifier (op amp). The differential equation governing the circuit taking into consideration the nonideal features of the op amp was formulated and solved for two types of inputs (ramp and dc) that are of interest for the application. Under certain time-dependent constraints, it is shown that (1) with a ramp input the circuit has an output voltage proportional to the static tip clearance capacitance, and (2) with a dc input, the output is proportional to the derivative of the clearance capacitance, and subsequent integration recovers the dynamic capacitance. The technique accommodates long cable lengths and environmentally induced changes in cable and probe parameters. System implementation for both static and dynamic measurements having the same high sensitivity is also presented.

Electrical capacitance clearanceometer

NASA Technical Reports Server (NTRS)

Hester, Norbert J. (Inventor); Hornbeck, Charles E. (Inventor); Young, Joseph C. (Inventor)

1992-01-01

A hot gas turbine engine capacitive probe clearanceometer is employed to measure the clearance gap or distance between blade tips on a rotor wheel and its confining casing under operating conditions. A braze sealed tip of the probe carries a capacitor electrode which is electrically connected to an electrical inductor within the probe which is inserted into a turbine casing to position its electrode at the inner surface of the casing. Electrical power is supplied through a voltage controlled variable frequency oscillator having a tuned circuit in which the probe is a component. The oscillator signal is modulated by a change in electrical capacitance between the probe electrode and a passing blade tip surface while an automatic feedback correction circuit corrects oscillator signal drift. A change in distance between a blade tip and the probe electrode is a change in capacitance therebetween which frequency modulates the oscillator signal. The modulated oscillator signal which is then processed through a phase detector and related circuitry to provide an electrical signal is proportional to the clearance gap.

Local Mass and Heat Transfer on a Turbine Blade Tip

DOE PAGES

Jin, P.; Goldstein, R. J.

2003-01-01

Locmore » al mass and heat transfer measurements on a simulated high-pressure turbine blade-tip surface are conducted in a linear cascade with a nonmoving tip endwall, using a naphthalene sublimation technique. The effects of tip clearance (0.86–6.90% of chord) are investigated at various exit Reynolds numbers (4–7 × 10 5 ) and turbulence intensities (0.2 and 12.0%). The mass transfer on the tip surface is significant along its pressure edge at the smallest tip clearance. At the two largest tip clearances, the separation bubble on the tip surface can cover the whole width of the tip on the second half of the tip surface. The average mass-transfer rate is highest at a tip clearance of 1.72% of chord. The average mass-transfer rate on the tip surface is four and six times as high as on the suction and the pressure surface, respectively. A high mainstream turbulence level of 12.0% reduces average mass-transfer rates on the tip surface, while the higher mainstream Reynolds number generates higher local and average mass-transfer rates on the tip surface.« less

The Three Dimensional Flow Field at the Exit of an Axial-Flow Turbine Rotor

NASA Technical Reports Server (NTRS)

Lakshminarayana, B.; Ristic, D.; Chu, S.

1998-01-01

A systematic and comprehensive investigation was performed to provide detailed data on the three dimensional viscous flow phenomena downstream of a modem turbine rotor and to understand the flow physics such as origin, nature, development of wakes, secondary flow, and leakage flow. The experiment was carried out in the Axial Flow Turbine Research Facility (AFTRF) at Penn State, with velocity measurements taken with a 3-D LDV System. Two radial traverses at 1% and 10% of chord downstream of the rotor have been performed to identify the three-dimensional flow features at the exit of the rotor blade row. Sufficient spatial resolution was maintained to resolve blade wake, secondary flow, and tip leakage flow. The wake deficit is found to be substantial, especially at 1% of chord downstream of the rotor. At this location, negative axial velocity occurs near the tip, suggesting flow separation in the tip clearance region. Turbulence intensities peak in the wake region, and cross- correlations are mainly associated with the velocity gradient of the wake deficit. The radial velocities, both in the wake and in the endwall region, are found to be substantial. Two counter-rotating secondary flows are identified in the blade passage, with one occupying the half span close to the casino and the other occupying the half span close to the hub. The tip leakage flow is well restricted to 10% immersion from the blade tip. There are strong vorticity distributions associated with these secondary flows and tip leakage flow. The passage averaged data are in good agreement with design values.

The singing vortex.

PubMed

Arndt, R; Pennings, P; Bosschers, J; van Terwisga, T

2015-10-06

Marine propellers display several forms of cavitation. Of these, propeller-tip vortex cavitation is one of the important factors in propeller design. The dynamic behaviour of the tip vortex is responsible for hull vibration and noise. Thus, cavitation in the vortices trailing from tips of propeller blades has been studied extensively. Under certain circumstances cavitating vortices have been observed to have wave-like disturbances on the surfaces of vapour cores. Intense sound at discrete frequencies can result from a coupling between tip vortex disturbances and oscillating sheet cavitation on the surfaces of the propeller blades. This research article focuses on the dynamics of vortex cavitation and more in particular on the energy and frequency content of the radiated pressures.

Design and performance of an 0.8 hub-tip ratio axial flow pump rotor with a blade tip diffusion factor of 0.55

NASA Technical Reports Server (NTRS)

Urasek, D. C.

1972-01-01

A 22.9-centimeter diameter axial flow rotor with a 0.8 hub-tip radius ratio, a design flow coefficient of 0.466, and a blade tip design diffusion factor of 0.55 was tested in cold water under both cavitating and noncavitating conditions. Radial surveys of the flow conditions at the rotor inlet and outlet were made. At design flow, the rotor produced an overall headrise coefficient of 0.360 with an overall efficiency of 95.0 percent. The efficiency remained greater than 88 percent over the entire flow coefficient range which varied from 0.350 to 0.615.

The singing vortex

PubMed Central

Arndt, R.; Pennings, P.; Bosschers, J.; van Terwisga, T.

2015-01-01

Marine propellers display several forms of cavitation. Of these, propeller-tip vortex cavitation is one of the important factors in propeller design. The dynamic behaviour of the tip vortex is responsible for hull vibration and noise. Thus, cavitation in the vortices trailing from tips of propeller blades has been studied extensively. Under certain circumstances cavitating vortices have been observed to have wave-like disturbances on the surfaces of vapour cores. Intense sound at discrete frequencies can result from a coupling between tip vortex disturbances and oscillating sheet cavitation on the surfaces of the propeller blades. This research article focuses on the dynamics of vortex cavitation and more in particular on the energy and frequency content of the radiated pressures. PMID:26442147

Control of Leakage Flow by Triple Squealer Configuration in Axial Flow Turbine

NASA Astrophysics Data System (ADS)

El-Ghandour, Mohamed; Ibrahim, Mohammed K.; Mori, Koichi; Nakamura, Yoshiaki

A new turbine blade tip shape called triple squealer is proposed. This shape is based on the conventional double squealer, and the cavity on the tip surface is divided into two parts by using a third squealer along the blade camber line. The effect of the ratio of groove depth to span (GDS ratio) was investigated. The flat-tip case (baseline case) and double-squealer case were calculated for comparison. In-house, unstructured, 3D, Navier-Stokes, finite volume, multiblock code with DES (Detached Eddy Simulation) as turbulence model was used to calculate the flow field around the tip. The computational results show that the reduction in the mass flow rate of the leakage flow for the triple squealer is 15.69% compared to the flat-tip case.

Turbine Engine Disk Rotor Health Monitoring Assessment Using Spin Tests Data

NASA Technical Reports Server (NTRS)

Abdul-Aziz, Ali; Woike, Mark; Baalini, George; Bodis, James R.

2012-01-01

Detecting rotating engine component malfunctions and structural anomalies is increasingly becoming a crucial key feature that will help boost safety and lower maintenance cost. However, achievement of such technology, which can be referred to as a health monitoring remains somewhat challenging to implement. This is mostly due to presence of scattered loading conditions, crack sizes, component geometry and material properties that hinders the simplicity of imposing such application. Different approaches are being considered to assist in developing other means of health monitoring or nondestructive techniques to detect hidden flaws and mini cracks before any catastrophic events occur. These methods extend further to assess material discontinuities and other defects that have matured to the level where a failure is very likely. This paper is focused on presenting data obtained from spin test experiments of a turbine engine like rotor disk and their correlation to the development of a structural health monitoring and fault detection system. The data collected includes blade tip clearance, blade tip timing measurements and shaft displacements. The experimental results are collected at rotational speeds up to 10,000 Rpm and tests are conducted at the NASA Glenn Research Center s Rotordynamics Laboratory via a high precision spin system. Additionally, this study offers a closer glance at a selective online evaluation of a rotating disk using advanced capacitive, microwave and eddy current sensor technology.

Turbine engine disk rotor health monitoring assessment using spin tests data

NASA Astrophysics Data System (ADS)

Abdul-Aziz, Ali; Woike, Mark; Baaklini, George; Bodis, James R.

2012-04-01

Detecting rotating engine component malfunctions and structural anomalies is increasingly becoming a crucial key feature that will help boost safety and lower maintenance cost. However, achievement of such technology, which can be referred to as a health monitoring remains somewhat challenging to implement. This is mostly due to presence of scattered loading conditions, crack sizes, component geometry and material properties that hinders the simplicity of imposing such application. Different approaches are being considered to assist in developing other means of health monitoring or nondestructive techniques to detect hidden flaws and mini cracks before any catastrophic events occur. These methods extend further to assess material discontinuities and other defects that have matured to the level where a failure is very likely. This paper is focused on presenting data obtained from spin test experiments of a turbine engine like rotor disk and their correlation to the development of a structural health monitoring and fault detection system. The data collected includes blade tip clearance, blade tip timing measurements and shaft displacements. The experimental results are collected at rotational speeds up to 10,000 Rpm and tests are conducted at the NASA Glenn Research Center's Rotordynamics Laboratory via a high precision spin system. Additionally, this study offers a closer glance at a selective online evaluation of a rotating disk using advanced capacitive, microwave and eddy current sensor technology.

Does the Miller blade truly provide a better laryngoscopic view and intubating conditions than the Macintosh blade in small children?

PubMed

Varghese, Elsa; Kundu, Ratul

2014-08-01

Both Miller and Macintosh blades are widely used for laryngoscopy in small children, though the Miller blade is more commonly recommended in pediatric anesthetic literature. The aim of this study was to compare laryngoscopic views and ease and success of intubation with Macintosh and Miller blades in small children under general anesthesia. One hundred and twenty children aged 1-24 months were randomized for laryngoscopy to be performed in a crossover manner with either the Miller or the Macintosh blade first, following induction of anesthesia and neuromuscular blockade. The tips of both the blades were placed at the vallecula. Intubation was performed following the second laryngoscopy. The glottic views with and without external laryngeal maneuver (ELM) and ease of intubation were observed. Similar glottic views with both blades were observed in 52/120 (43%) children, a better view observed with the Miller blade in 35/120 (29%) children, and with the Macintosh blade in 33/120 (28%). Laryngoscopy was easy in 65/120 (54%) children with both the blades. Restricted laryngoscopy was noted in 55 children: in 27 children with both the blades, 15 with Miller, and 13 with Macintosh blade. Laryngoscopic view improved following ELM with both the blades. In children aged 1-24 months, the Miller and the Macintosh blades provide similar laryngoscopic views and intubating conditions. When a restricted view is obtained, a change of blade may provide a better view. Placing the tip of the Miller blade in the vallecula provides satisfactory intubating conditions in this age group. © 2014 John Wiley & Sons Ltd.

Adaptive Development

NASA Technical Reports Server (NTRS)

2005-01-01

The goal of this research is to develop and demonstrate innovative adaptive seal technologies that can lead to dramatic improvements in engine performance, life, range, and emissions, and enhance operability for next generation gas turbine engines. This work is concentrated on the development of self-adaptive clearance control systems for gas turbine engines. Researchers have targeted the high-pressure turbine (HPT) blade tip seal location for following reasons: Current active clearance control (ACC) systems (e.g., thermal case-cooling schemes) cannot respond to blade tip clearance changes due to mechanical, thermal, and aerodynamic loads. As such they are prone to wear due to the required tight running clearances during operation. Blade tip seal wear (increased clearances) reduces engine efficiency, performance, and service life. Adaptive sealing technology research has inherent impact on all envisioned 21st century propulsion systems (e.g. distributed vectored, hybrid and electric drive propulsion concepts).

Some observations of the effects of radial distortions on performance of a transonic rotating blade row

NASA Technical Reports Server (NTRS)

Sandercock, D. M.; Sanger, N. L.

1974-01-01

A single rotating blade row was tested with two magnitudes of tip radial distortion and two magnitudes of hub radial distortion imposed on the inlet flow. The rotor was about 50 centimeters (20 in.) in diameter and had a design operating tip speed of approximately 420 meters per second (1380 ft/sec). Overall performance at 60, 80, and 100 percent of equivalent design speed generally showed a decrease (compared to undistorted flow) in rotor stall margin with tip radial distortion but no change, or a slight increase, in rotor stall margin with hub radial distortion. At design speed there was a decrease in rotor overall total pressure ratio and choke flow with all inlet flow distortions. Radial distributions of blade element parameters are presented for selected operating conditions at design speed.

Numerical study on air-structure coupling dynamic characteristics of the axial fan blade

NASA Astrophysics Data System (ADS)

Chen, Q. G.; Xie, B.; Li, F.; Gu, W. G.

2013-12-01

In order to understand the dynamic characteristics of the axial-flow fan blade due to the effect of rotating stress and the action of unsteady aerodynamic forces caused by the airflow, a numerical simulation method for air-structure coupling in an axial-flow fan with fixed rear guide blades was performed. The dynamic characteristics of an axial-flow fan rotating blade were studied by using the two-way air-structure coupling method. Based on the standard k-ε turbulence model, and using weak coupling method, the preceding six orders modal parameters of the rotating blade were obtained, and the distributions of stress and strain on the rotating blade were presented. The results show that the modal frequency from the first to the sixth order is 3Hz higher than the modal frequency without considering air-structure coupling interaction; the maximum stress and the maximum strain are all occurred in the vicinity of root area of the blade no matter the air-structure coupling is considered or not, thus, the blade root is the dangerous location subjected to fatigue break; the position of maximum deformation is at the blade tip, so the vibration of the blade tip is significant. This study can provide theoretical references for the further study on the strength analysis and mechanical optimal design.

Acoustic results of supersonic tip speed fan blade modification

NASA Technical Reports Server (NTRS)

Jutras, R. R.; Kazin, S. B.

1974-01-01

A supersonic tip speed single stage fan was modified with the intent of reducing multiple pure tone (MPT) or buzz saw noise. There were three modifications to the blades from the original design. The modifications to the blade resulted in an increase in cascade throat area causing the shock to start at a lower corrected fan speed. The acoustic results without acoustically absorbing liners showed substantial reduction in multiple pure tone levels. However, an increase in the blade passing frequency noise at takeoff fan speed accompanied the MPT reduction. The net result however, was a reduction in the maximum 1000-foot (304.8 m) altitude level flyover PNL. For the case with acoustic treatment in the inlet outer wall, the takeoff noise increased relative to an acoustically treated baseline. This was largely due to the increased blade passing frequency noise which was not effectively reduced by the liner.

Numerical study of the trailing vortex of a wing with wing-tip blowing

NASA Technical Reports Server (NTRS)

Lim, Hock-Bin

1994-01-01

Trailing vortices generated by lifting surfaces such as helicopter rotor blades, ship propellers, fixed wings, and canard control surfaces are known to be the source of noise, vibration, cavitation, degradation of performance, and other hazardous problems. Controlling these vortices is, therefore, of practical interest. The formation and behavior of the trailing vortices are studied in the present research. In addition, wing-tip blowing concepts employing axial blowing and spanwise blowing are studied to determine their effectiveness in controlling these vortices and their effects on the performance of the wing. The 3D, unsteady, thin-layer compressible Navier-Stokes equations are solved using a time-accurate, implicit, finite difference scheme that employs LU-ADI factorization. The wing-tip blowing is simulated using the actuator plane concept, thereby, not requiring resolution of the jet slot geometry. Furthermore, the solution blanking feature of the chimera scheme is used to simplify the parametric study procedure for the wing-tip blowing. Computed results are shown to compare favorably with experimental measurements. It is found that axial wing-tip blowing, although delaying the rolling-up of the trailing vortices and the near-field behavior of the flowfield, does not dissipate the circulation strength of the trailing vortex farther downstream. Spanwise wing-tip blowing has the effect of displacing the trailing vortices outboard and upward. The increased 'wing-span' due to the spanwise wing-tip blowing has the effect of lift augmentation on the wing and the strengthening of the trailing vortices. Secondary trailing vortices are created at high spanwise wing-tip blowing intensities.

Numerical Analysis of the Acoustic Field of Tip-Clearance Flow

NASA Astrophysics Data System (ADS)

Alavi Moghadam, S. M.; M. Meinke Team; W. Schröder Team

2015-11-01

Numerical simulations of the acoustic field generated by a shrouded axial fan are studied by a hybrid fluid-dynamics-acoustics method. In a first step, large-eddy simulations are performed to investigate the dynamics of tip clearance flow for various tip gap sizes and to determine the acoustic sources. The simulations are performed for a single blade out of five blades with periodic boundary conditions in the circumferential direction on a multi-block structured mesh with 1.4 ×108 grid points. The turbulent flow is simulated at a Reynolds number of 9.36 ×105 at undisturbed inflow condition and the results are compared with experimental data. The diameter and strength of the tip vortex increase with the tip gap size, while simultaneously the efficiency of the fan decreases. In a second step, the acoustic field on the near field is determined by solving the acoustic perturbation equations (APE) on a mesh for a single blade consisting of approx. 9.8 ×108 grid points. The overall agreement of the pressure spectrum and its directivity with measurements confirm the correct identification of the sound sources and accurate prediction of the acoustic duct propagation. The results show that the longer the tip gap size the higher the broadband noise level. Senior Scientist, Institute of Aerodynamics, RWTH Aachen University.

Measurements of the tip leakage vortex structures and turbulence in the meridional plane of an axial water-jet pump

NASA Astrophysics Data System (ADS)

Wu, Huixuan; Miorini, Rinaldo L.; Katz, Joseph

2011-04-01

Particle image velocimetry (PIV) measurements at varying resolutions focus on the flow structures in the tip region of a water-jet pump rotor, including the tip-clearance flow and the rollup process of a tip leakage vortex (TLV). Unobstructed views of these regions are facilitated by matching the optical refractive index of the transparent pump with that of the fluid. High-magnification data reveal the flow non-uniformities and associated turbulence within the tip gap. Instantaneous data and statistics of spatial distributions and strength of vortices in the rotor passage reveal that the leakage flow emerges as a wall jet with a shear layer containing a train of vortex filaments extending from the tip of the blade. These vortices are entrained into the TLV, but do not have time to merge. TLV breakdown in the aft part of the blade passage further fragments these structures, increasing their number and reducing their size. Analogy is made between the circumferential development of the TLV in the blade passage and that of the starting jet vortex ring rollup. Subject to several assumptions, these flows display similar trends, including conditions for TLV separation from the shear layer feeding vorticity into it.

Analysis of the sweeped actuator line method

DOE PAGES

Nathan, Jörn; Masson, Christian; Dufresne, Louis; ...

2015-10-16

The actuator line method made it possible to describe the near wake of a wind turbine more accurately than with the actuator disk method. Whereas the actuator line generates the helicoidal vortex system shed from the tip blades, the actuator disk method sheds a vortex sheet from the edge of the rotor plane. But with the actuator line come also temporal and spatial constraints, such as the need for a much smaller time step than with actuator disk. While the latter one only has to obey the Courant-Friedrichs-Lewy condition, the former one is also restricted by the grid resolution andmore » the rotor tip-speed. Additionally the spatial resolution has to be finer for the actuator line than with the actuator disk, for well resolving the tip vortices. Therefore this work is dedicated to examining a method in between of actuator line and actuator disk, which is able to model the transient behavior, such as the rotating blades, but which also relaxes the temporal constraint. Therefore a larger time-step is used and the blade forces are swept over a certain area. As a result, the main focus of this article is on the aspect of the blade tip vortex generation in comparison with the standard actuator line and actuator disk.« less

Analysis of the sweeped actuator line method

DOE Office of Scientific and Technical Information (OSTI.GOV)

Nathan, Jörn; Masson, Christian; Dufresne, Louis

The actuator line method made it possible to describe the near wake of a wind turbine more accurately than with the actuator disk method. Whereas the actuator line generates the helicoidal vortex system shed from the tip blades, the actuator disk method sheds a vortex sheet from the edge of the rotor plane. But with the actuator line come also temporal and spatial constraints, such as the need for a much smaller time step than with actuator disk. While the latter one only has to obey the Courant-Friedrichs-Lewy condition, the former one is also restricted by the grid resolution andmore » the rotor tip-speed. Additionally the spatial resolution has to be finer for the actuator line than with the actuator disk, for well resolving the tip vortices. Therefore this work is dedicated to examining a method in between of actuator line and actuator disk, which is able to model the transient behavior, such as the rotating blades, but which also relaxes the temporal constraint. Therefore a larger time-step is used and the blade forces are swept over a certain area. As a result, the main focus of this article is on the aspect of the blade tip vortex generation in comparison with the standard actuator line and actuator disk.« less

Impact of tip-gap size and periodicity on turbulent transition

NASA Astrophysics Data System (ADS)

Pogorelov, Alexej; Meinke, Matthias; Schroeder, Wolfgang

2015-11-01

Large-Eddy Simulations of the flow field in an axial fan are performed at a Reynolds number of 936.000 based on the diameter and the rotational speed of the casing wall. A finite-volume flow solver based on a conservative Cartesian cut-cell method is used to solve the unsteady compressible Navier-Stokes equations. Computations are performed at a flow rate coefficient of 0.165 and a tip-gap size of s/D =0.01, for a 72 degrees fan section resolving only one out of five blades and a full fan resolving all five blades to investigate the impact of the periodic boundary condition. Furthermore, a grid convergence study is performed using four computational grids. Results of the flow field are analyzed for the computational grid with 1 billion cells. An interaction of the turbulent wake, generated by the tip-gap vortex, with the downstream blade, is observed, which leads to a cyclic transition with high pressure fluctuations on the suction side of the blade. Two dominant frequencies are identified which perfectly match with the characteristic frequencies in the experimental sound power level such that their physical origin is explained. A variation of the tip-gap size alters the transition on the suction side, i.e., no cyclic transition is observed.

Measurement of Rotating Blade Tip Clearance with Fibre-Optic Probe

NASA Astrophysics Data System (ADS)

Cao, S. Z.; Duan, F. J.; Zhang, Y. G.

2006-10-01

This paper described a tip clearance measuring system with fibre-optic probe. The system is based on a novel tip clearance sensor of optical fibre-bundle mounted on the casing, rotating speed synchronization sensor mounted on the rotating shaft, the tip clearance preamplification processing circuit followed by high speed data-acquisition unit. A novel tip clearance sensor of trifurcated optical fibre bundle was proposed and demonstrated. It is independent of material of measured surface but capacitive probe demands target conductive. Measurements can be taken under severe conditions such as ionization. Sensor circuitry and data acquisition circuit were successfully designed. With the help of Rotation synchronized sensor, all the blades can be detected in real-time. Because of fibre-optic sensor, the measuring system has commendably frequency response, which can work well in high rotating speed from 0-15000rpm.The measurement range of tip clearance is 0-3mm with 25um precision.

Numerical investigation of tip clearance cavitation in Kaplan runners

NASA Astrophysics Data System (ADS)

Nikiforova, K.; Semenov, G.; Kuznetsov, I.; Spiridonov, E.

2016-11-01

There is a gap between the Kaplan runner blade and the shroud that makes for a special kind of cavitation: cavitation in the tip leakage flow. Two types of cavitation caused by the presence of clearance gap are known: tip vortex cavitation that appears at the core of the rolled up vortex on the blade suction side and tip clearance cavitation that appears precisely in the gap between the blade tip edge and the shroud. In the context of this work numerical investigation of the model Kaplan runner has been performed taking into account variable tip clearance for several cavitation regimes. The focus is put on investigation of structure and origination of mechanism of cavitation in the tip leakage flow. Calculations have been performed with the help of 3-D unsteady numerical model for two-phase medium. Modeling of turbulent flow in this work has been carried out using full equations of Navier-Stokes averaged by Reynolds with correction for streamline curvature and system rotation. For description of this medium (liquid-vapor) simplification of Euler approach is used; it is based on the model of interpenetrating continuums, within the bounds of this two- phase medium considered as a quasi-homogeneous mixture with the common velocity field and continuous distribution of density for both phases. As a result, engineering techniques for calculation of cavitation conditioned by existence of tip clearance in model turbine runner have been developed. The detailed visualization of the flow was carried out and vortex structure on the suction side of the blade was reproduced. The range of frequency with maximum value of pulsation was assigned and maximum energy frequency was defined; it is based on spectral analysis of the obtained data. Comparison between numerical computation results and experimental data has been also performed. The location of cavitation zone has a good agreement with experiment for all analyzed regimes.

Experimental study of rotating wind turbine breakdown characteristics in large scale air gaps

NASA Astrophysics Data System (ADS)

Wang, Yu; Qu, Lu; Si, Tianjun; Ni, Yang; Xu, Jianwei; Wen, Xishan

2017-06-01

When a wind turbine is struck by lightning, its blades are usually rotating. The effect of blade rotation on a turbine’s ability to trigger a lightning strike is unclear. Therefore, an arching electrode was used in a wind turbine lightning discharge test to investigate the difference in lightning triggering ability when blades are rotating and stationary. A negative polarity switching waveform of 250/2500 μs was applied to the arching electrode and the up-and-down method was used to calculate the 50% discharge voltage. Lightning discharge tests of a 1:30 scale wind turbine model with 2, 4, and 6 m air gaps were performed and the discharge process was observed. The experimental results demonstrated that when a 2 m air gap was used, the breakdown voltage increased as the blade speed was increased, but when the gap length was 4 m or longer, the trend was reversed and the breakdown voltage decreased. The analysis revealed that the rotation of the blades changes the charge distribution in the blade-tip region, promotes upward leader development on the blade tip, and decreases the breakdown voltage. Thus, the blade rotation of a wind turbine increases its ability to trigger lightning strikes.

Three-dimensional analysis of the Pratt and Whitney alternate design SSME fuel turbine

NASA Technical Reports Server (NTRS)

Kirtley, K. R.; Beach, T. A.; Adamczyk, J. J.

1991-01-01

The three dimensional viscous time-mean flow in the Pratt and Whitney alternate design space shuttle main engine fuel turbine is simulated using the average passage Navier-Stokes equations. The migration of secondary flows generated by upstream blade rows and their effect on the performance of downstream blade rows is studied. The present simulation confirms that the flow in this two stage turbine is highly three dimensional and dominated by the tip leakage flow. The tip leakage vortex generated by the first blade persists through the second blade and adversely affects its performance. The greatest mixing of the inlet total temperature distortion occurs in the second vane and is due to the large leakage vortex generated by the upstream rotor. It is assumed that the predominant spanwise mixing mechanism in this low aspect ratio turbine is the radial transport due to the deterministically unsteady vortical flow generated by upstream blade rows. A by-product of the analysis is accurate pressure and heat loads for all blade rows under the influence of neighboring blade rows. These aero loads are useful for advanced structural analysis of the vanes and blades.

A laser-optical sensor system for blade vibration detection of high-speed compressors

NASA Astrophysics Data System (ADS)

Neumann, Mathias; Dreier, Florian; Günther, Philipp; Wilke, Ulrich; Fischer, Andreas; Büttner, Lars; Holzinger, Felix; Schiffer, Heinz-Peter; Czarske, Jürgen

2015-12-01

Improved efficiency as well as increased lifetime of turbines and compressors are important goals in turbomachinery development. A significant enhancement to accomplish these aims can be seen in online monitoring of the operating parameters of the machines. During the operation of compressors it is of high interest to predict critical events like flutter or stall which can be achieved by observing blade deformations and vibrations. We have developed a laser Doppler distance sensor (LDDS), which is capable of simultaneously measuring the radial blade expansions, the circumferential blade deflections as well as the circumferential velocities of the rotor blade tips. As a result, an increase of blade vibrations is measured before stall at characteristic frequencies. While the detected vibration frequencies and the vibration increase are in agreement with the measurement results of a commercial capacitive blade tip timing system, the measured values of the vibration amplitudes differ by a factor of three. This difference can be mainly attributed to the different measurement locations and to the different measurement approaches. Since the LDDS is applicable to metal as well as ceramic, carbon-fiber and glass-fiber reinforced composite blades, a universally applicable sensor system for stall prediction and status monitoring is presented.

Wind-Tunnel Investigation of the Effect of Angle of Attack and Flapping-Hinge Offset on Periodic Bending Moments and Flapping of a Small Rotor

NASA Technical Reports Server (NTRS)

McCarty, John Locke; Brooks, George W.; Maglieri, Domenic J.

1959-01-01

A two-blade rotor having a diameter of 4 feet and a solidity of 0.037 was tested in the Langley 300-MPH 7- by 10-foot tunnel to obtain information on the effect of certain rotor variables on the blade periodic bending moments and flapping angles during the various stages of transformation between the helicopter and autogiro configuration. Variables studied included collective pitch angle, flapping-hinge offset, rotor angle of attack, and tip-speed ratio. The results show that the blade periodic bending moments generally increase with tip-speed ratio up into the transition region, diminish over a certain range of tip-speed ratio, and increase again at higher tip-speed ratios. Above the transition region, the bending moments increase with collective pitch angle and rotor angle of attack. The absence of a flapping hinge results in a significant amplification of the periodic bending moments, the magnitudes of which increase with tip-speed ratio. When the flapping hinge is used, an increase in flapping-hinge offset results in reduced period bending moments. The aforementioned trends exhibited by the bending moments for changes in the variables are essentially duplicated by the periodic flapping motions. The existence of substantial amounts of blade stall increased both the periodic bending moments and the flapping angles. Harmonic analysis of the bending moments shows significant contributions of the higher harmonics, particularly in the transition region.

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

NASA Technical Reports Server (NTRS)

Podboy, Gary G.; Krupar, Martin J.; Sutliff, Daniel L.; Horvath, Csaba

2007-01-01

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.

Rotorcraft In-Plane Noise Reduction Using Active/Passive Approaches with Induced Vibration Tracking

NASA Astrophysics Data System (ADS)

Chia, Miang Hwee

A comprehensive study of the use of active and passive approaches for in-plane noise reduction, including the vibrations induced during noise reduction, was conducted on a hingeless rotor configuration resembling the MBB BO-105 rotor. First, a parametric study was performed to examine the effects of rotor blade stiffness on the vibration and noise reduction performance of a 20%c plain trailing edge flap and a 1.5%c sliding microflap. This was accomplished using a comprehensive code AVINOR (for Active VIbration and NOise Reduction). A two-dimensional unsteady reduced order aerodynamic model (ROM), using the Rational Function Approximation approach and CFD-based oscillatory aerodynamic load data, was used in the comprehensive code. The study identified a hingeless blade configuration with torsional frequency of 3.17/rev as an optimum configuration for studying vibration and noise reduction using on-blade control devices such as flaps or microflaps. Subsequently, a new suite of computational tools capable of predicting in-plane low frequency sound pressure level (LFSPL) rotorcraft noise and its control was developed, replacing the acoustic module WOPWOP in AVINOR with a new acoustic module HELINOIR (for HELIcopter NOIse Reduction), which overcomes certain limitations associated with WOPWOP. The new suite, consisting of the AVINOR/HELINOIR combination, was used to study active flaps, as well as microflaps operating in closed-loop mode for in-plane noise reduction. An alternative passive in-plane noise reduction approach using modification to the blade tip in the 10%R outboard region was also studied. The new suite consisting of the AVINOR/HELINOIR combination based on a compact aeroacoustic model was validated by comparing with wind tunnel test results, and subsequently verified by comparing with computational results. For active control, the in-plane noise reduction obtained with a single 20%c plain trailing edge flap during level flight at a moderate advance ratio was examined. Different configurations of far-field and near-field feedback microphone locations were examined to develop a fundamental understanding of the feedback microphone locations on the noise reduction process A near-field microphone located on the tip of a nose boom was found to produce a LFSPL reduction of up to 6dB. However, this noise reduction was accompanied by an out-of-plane noise increase of 18dB and 60% increase in vertical hub shear. For passive control, three tip geometries having sweep, dihedral, and anhedral, were considered. The tip dihedral reduced LFSPL by up to 2dB without a vibratory load penalty. However, this was accompanied by an increase in the mid frequency sound pressure levels (MFSPL). The tip sweep and tip anhedral produced an increase in in-plane LFSPL below the horizon. A comparison of the active and passive approaches indicated that active approaches implemented by a plain flap with a feedback microphone located on the nose boom is superior to the passive control approaches. However, there is a general trade-off between LFSPL reduction, MFSPL generation and vibratory hub loads induced by noise control.

Blade Vibration Measurement System

NASA Technical Reports Server (NTRS)

Platt, Michael J.

2014-01-01

The Phase I project successfully demonstrated that an advanced noncontacting stress measurement system (NSMS) could improve classification of blade vibration response in terms of mistuning and closely spaced modes. The Phase II work confirmed the microwave sensor design process, modified the sensor so it is compatible as an upgrade to existing NSMS, and improved and finalized the NSMS software. The result will be stand-alone radar/tip timing radar signal conditioning for current conventional NSMS users (as an upgrade) and new users. The hybrid system will use frequency data and relative mode vibration levels from the radar sensor to provide substantially superior capabilities over current blade-vibration measurement technology. This frequency data, coupled with a reduced number of tip timing probes, will result in a system capable of detecting complex blade vibrations that would confound traditional NSMS systems. The hardware and software package was validated on a compressor rig at Mechanical Solutions, Inc. (MSI). Finally, the hybrid radar/tip timing NSMS software package and associated sensor hardware will be installed for use in the NASA Glenn spin pit test facility.

Analytical design of an advanced radial turbine. [automobile engines

NASA Technical Reports Server (NTRS)

Large, G. D.; Finger, D. G.; Linder, C. G.

1981-01-01

The aerodynamic and mechanical potential of a single stage ceramic radial inflow turbine was evaluated for a high temperature single stage automotive engine. The aerodynamic analysis utilizes a turbine system optimization technique to evaluate both radial and nonradial rotor blading. Selected turbine rotor configurations were evaluated mechanically with three dimensional finite element techniques. Results indicate that exceptionally high rotor tip speeds (2300 ft/sec) and performance potential are feasible with radial bladed rotors if the projected ceramic material properties are realized. Nonradial rotors reduced tip speed requirements (at constant turbine efficiency) but resulted in a lower cumulative probability of success due to higher blade and disk stresses.

Rotorcraft Blade-Vortex Interaction Controller

NASA Technical Reports Server (NTRS)

Schmitz, Fredric H. (Inventor)

1995-01-01

Blade-vortex interaction noises, sometimes referred to as 'blade slap', are avoided by increasing the absolute value of inflow to the rotor system of a rotorcraft. This is accomplished by creating a drag force which causes the angle of the tip-path plane of the rotor system to become more negative or more positive.

Preventing pressure ulcers

MedlinePlus

... Hips Spine Tailbone area Elbows Shoulders and shoulder blades Back of the head Ears Call your health ... your tailbone area Under your shoulders and shoulder blades Under your elbows Other tips are: DO NOT ...

Initial Aerodynamic and Acoustic Study of an Active Twist Rotor Using a Loosely Coupled CFD/CSD Method

NASA Technical Reports Server (NTRS)

Boyd, David D. Jr.

2009-01-01

Preliminary aerodynamic and performance predictions for an active twist rotor for a HART-II type of configuration are performed using a computational fluid dynamics (CFD) code, OVERFLOW2, and a computational structural dynamics (CSD) code, CAMRAD -II. These codes are loosely coupled to compute a consistent set of aerodynamics and elastic blade motions. Resultant aerodynamic and blade motion data are then used in the Ffowcs-Williams Hawkins solver, PSU-WOPWOP, to compute noise on an observer plane under the rotor. Active twist of the rotor blade is achieved in CAMRAD-II by application of a periodic torsional moment couple (of equal and opposite sign) at the blade root and tip at a specified frequency and amplitude. To provide confidence in these particular active twist predictions for which no measured data is available, the rotor system geometry and computational set up examined here are identical to that used in a previous successful Higher Harmonic Control (HHC) computational study. For a single frequency equal to three times the blade passage frequency (3P), active twist is applied across a range of control phase angles at two different amplitudes. Predicted results indicate that there are control phase angles where the maximum mid-frequency noise level and the 4P non -rotating hub vibrations can be reduced, potentially, both at the same time. However, these calculated reductions are predicted to come with a performance penalty in the form of a reduction in rotor lift-to-drag ratio due to an increase in rotor profile power.

Altitude-Wind-Tunnel Investigation of Performance of Several Propellers on YP-47M Airplane at High Blade Loadings. 6; Hamilton Standard 6507A-2 Four- and Three-Blade Propellers

NASA Technical Reports Server (NTRS)

Saari, Martin J.; Sorin, Solomon M.

1946-01-01

An altitude-wind-tunnel investigation has been made to determine the performance of Hamilton Standard 6507A-2 four-blade and three-blade propellers on a YP-47M airplane at high blade loadings and high engine powers. Characteristics of the four-blase propeller were obtained for a range of power coefficients from 0.10 to 1.00 at free-stream Mach numbers of 0.20, 0.30, 0.40. Characteristics of the three-blade propeller were obtained for a range of power coefficients from 0.30 to 1.00 at a free-stream Mach number of 0.40. Results of the force measurements indicate primarily the trend of propeller efficiency for changes in power coefficient or advance-diameter ratio because no corrections for the effects of tunnel-wall constriction on the installation were applied. Slipstream surveys are presented to illustrate blade thrust load distribution for certain operating conditions. Within the range of advance-diameter ratios investigated at each free-stream Mach number, the efficiency of the four-blade propeller decreased as the power coefficient was increased from 0.10 to 1.00. For the three-blade propeller, nearly constant maximum efficiencies were obtained for power coefficients from 0.32 to 0.63 at advance-diameter ratios between 1.90 and 3.00. In general, for conditions below the stall and critical tip Mach number, the maximum thrust load shifted from the inboard sections toward the tip sections as the power coefficient was increased or as the advance-diameter ratio was decreased. For conditions beyond the stall or critical tip Mach number, losses in thrust occurred on the outboard blade sections owing to flow break-down; the thrust load increased slightly on the inboard sections.

Analysis of helicopter blade vortex structure by laser velocimetry

NASA Astrophysics Data System (ADS)

Boutier, A.; Lefèvre, J.; Micheli, F.

1996-05-01

In descent flight, helicopter external noise is mainly generated by the Blade Vortex Interaction (BVI). To under-stand the dynamics of this phenomenon, the vortex must be characterized before its interaction with the blade, which means that its viscous core radius, its strength and its distance to the blade have to be determined by non-intrusive measurement techniques. As part of the HART program (Higher Harmonic Control Aeroacoustic Rotor Test, jointly conducted by US Army, NASA, DLR, DNW and ONERA), a series of tests have been made in the German Dutch Wind Tunnel (DNW) on a helicopter rotor with 2 m long blades, rotating at 1040 rpm; several flight configurations, with an advance ratio of 0.15 and a shaft angle of 5.3°, have been studied with different higher harmonic blade pitch angles superposed on the conventional one (corresponding to the baseline case). The flow on the retreating side has been analyzed with an especially designed 3D laser velocimeter, and, simultaneously, the blade tip attitude has been determined in order to get the blade-vortex miss distance, which is a crucial parameter in the noise reduction. A 3D laser velocimeter, in backscatter mode with a working distance of 5 m, was installed on a platform 9 m high, and flow seeding with submicron incense smoke was achieved in the settling chamber using a remotely controlled displacement device. Acquisition of instantaneous velocity vectors by an IFA 750 yielded mean velocity and turbulence maps across the vortex as well as the vortex position, intensity and viscous radius. The blade tip attitude (altitude, jitter, angle of incidence) was recorded by the TART method (Target Attitude in Real Time) which makes use of a CCD camera on which is formed the image of two retroreflecting targets attached to the blade tip and lighted by a flash lamp. In addition to the mean values of the aforementioned quantities, spectra of their fluctuations have been established up to 8 Hz.

The Effect of Unsteady Wakes on Turbine Tip Gap Leakage

DTIC Science & Technology

2013-05-10

low speed wind tunnel. The turbine blade shape for the experiment was the GE E 3 high pressure turbine stage 1 blade (Halila et al. 1982). The E 3...DATES COVERED (From - To) 4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER The Effect of Unsteady Wakes on Turbine Tip Gap Leakage...Gas turbines are found in military and civilian aircraft, ships, and power plants. Because of this widespread use, relatively small improvements

Rotor blades for turbine engines

DOEpatents

Piersall, Matthew R; Potter, Brian D

2013-02-12

A tip shroud that includes a plurality of damping fins, each damping fin including a substantially non-radially-aligned surface that is configured to make contact with a tip shroud of a neighboring rotor blade. At least one damping fin may include a leading edge damping fin and at least one damping fin may include a trailing edge damping fin. The leading edge damping fin may be configured to correspond to the trailing edge damping fin.

Propeller rotation noise due to torque and thrust

NASA Technical Reports Server (NTRS)

Deming, Arthur F

1940-01-01

Sound pressure of the first four harmonics of rotation from a full-scale two-blade propeller were measured and are compared with values calculated from theory. The comparison is made (1) for the space distribution with constant tip speed and (2) for fixed space angles with variable tip speed. A relation for rotation noise from an element of radius developed by Gutin is given showing the effect of number of blades on the rotation noise.

Numerical investigation & comparison of a tandem-bladed turbocharger centrifugal compressor stage with conventional design

NASA Astrophysics Data System (ADS)

Danish, Syed Noman; Qureshi, Shafiq Rehman; EL-Leathy, Abdelrahman; Khan, Salah Ud-Din; Umer, Usama; Ma, Chaochen

2014-12-01

Extensive numerical investigations of the performance and flow structure in an unshrouded tandem-bladed centrifugal compressor are presented in comparison to a conventional compressor. Stage characteristics are explored for various tip clearance levels, axial spacings and circumferential clockings. Conventional impeller was modified to tandem-bladed design with no modifications in backsweep angle, meridional gas passage and camber distributions in order to have a true comparison with conventional design. Performance degradation is observed for both the conventional and tandem designs with increase in tip clearance. Linear-equation models for correlating stage characteristics with tip clearance are proposed. Comparing two designs, it is clearly evident that the conventional design shows better performance at moderate flow rates. However; near choke flow, tandem design gives better results primarily because of the increase in throat area. Surge point flow rate also seems to drop for tandem compressor resulting in increased range of operation.

JT8D high pressure compressor performance improvement

NASA Technical Reports Server (NTRS)

Gaffin, W. O.

1981-01-01

An improved performance high pressure compressor with potential application to all models of the JT8D engine was designed. The concept consisted of a trenched abradable rubstrip which mates with the blade tips for each of the even rotor stages. This feature allows tip clearances to be set so blade tips run at or near the optimum radius relative to the flowpath wall, without the danger of damaging the blades during transients and maneuvers. The improved compressor demonstrated thrust specific fuel consumption and exhaust gas temperature improvements of 1.0 percent and at least 10 C over the takeoff and climb power range at sea level static conditions, compared to a bill-of-material high pressure compressor. Surge margin also improved 4 percentage points over the high power operating range. A thrust specific fuel consumption improvement of 0.7 percent at typical cruise conditions was calculated based on the sea level test results.

Rotator cuff - self-care

MedlinePlus

... good posture in general to keep your shoulder blade and joint in their right positions. Other tips ... joint Good range of motion of your shoulder blade and upper spine No pain during certain physical ...

Preliminary Aerodynamic Investigation of Fan Rotor Blade Morphing

NASA Technical Reports Server (NTRS)

Tweedt, Daniel L.

2012-01-01

Various new technologies currently under development may enable controlled blade shape variability, or so-called blade morphing, to be practically employed in aircraft engine fans and compressors in the foreseeable future. The current study is a relatively brief, preliminary computational fluid dynamics investigation aimed at partially demonstrating and quantifying the aerodynamic potential of fan rotor blade morphing. The investigation is intended to provide information useful for near-term planning, as well as aerodynamic solution data sets that can be subsequently analyzed using advanced acoustic diagnostic tools, for the purpose of making fan noise comparisons. Two existing fan system models serve as baselines for the investigation: the Advanced Ducted Propulsor fan with a design tip speed of 806 ft/sec and a pressure ratio of 1.294, and the Source Diagnostic Test fan with a design tip speed of 1215 ft/sec and a pressure ratio of 1.470. Both are 22-in. sub-scale, low-noise research fan/nacelle models that have undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. The study, restricted to fan rotor blade morphing only, involves a fairly simple blade morphing technique. Specifically, spanwise-linear variations in rotor blade-section setting angle are applied to alter the blade shape; that is, the blade is linearly retwisted from hub to tip. Aerodynamic performance comparisons are made between morphed-blade and corresponding baseline configurations on the basis of equal fan system thrust, where rotor rotational speed for the morphed-blade fan is varied to change the thrust level for that configuration. The results of the investigation confirm that rotor blade morphing could be a useful technology, with the potential to enable significant improvements in fan aerodynamic performance. Even though the study is very limited in scope and confined to simple geometric perturbations of two existing fan systems, the aerodynamic effectiveness of blade morphing is demonstrated by the configurations analyzed. In particular, for the Advanced Ducted Propulsor fan it is demonstrated that the performance levels of the original variable-pitch baseline design can be achieved using blade morphing instead of variable pitch, and for the Source Diagnostic Test fan the performance at important off-design operating points is substantially increased with blade morphing.

Computational study of the effects of shroud geometric variation on turbine performance in a 1.5-stage high-loaded turbine

NASA Astrophysics Data System (ADS)

Jia, Wei; Liu, Huoxing

2013-10-01

Generally speaking, main flow path of gas turbine is assumed to be perfect for standard 3D computation. But in real engine, the turbine annulus geometry is not completely smooth for the presence of the shroud and associated cavity near the end wall. Besides, shroud leakage flow is one of the dominant sources of secondary flow in turbomachinery, which not only causes a deterioration of useful work but also a penalty on turbine efficiency. It has been found that neglect shroud leakage flow makes the computed velocity profiles and loss distribution significantly different to those measured. Even so, the influence of shroud leakage flow is seldom taken into consideration during the routine of turbine design due to insufficient understanding of its impact on end wall flows and turbine performance. In order to evaluate the impact of tip shroud geometry on turbine performance, a 3D computational investigation for 1.5-stage turbine with shrouded blades was performed in this paper. The following geometry parameters were varied respectively: Inlet cavity length and exit cavity length

Experimental performance and analysis of 15.04-centimeter-tip-diameter, radial-inflow turbine with work factor of 1.126 and thick blading

NASA Technical Reports Server (NTRS)

Mclallin, K. L.; Haas, J. E.

1980-01-01

The aerodynamic design, the performance, and an internal loss breakdown were examined for a 15.04 cm tip diameter, radial-inflow turbine. The design application was to drive a two stage, 10 to 1 pressure ratio compressor with a mass flow of 0.952 kg/sec and a rotative speed of 70,000 rmp. The turbine inlet temperature was 1478 K, and the turbine was designed with blades thick enough for internal cooling passages. The rotor tip diameter was limited to 86 percent of optimum in order to obtain a reduced tip speed design. The turbine was fabricated with solid, uncooled blading and tested in air at nominal inlet pressure and temperature of 1.379 x 10000 N/sq m and 322.2 K, respectively. Results indicated the turbine total efficiency to be 5.3 points less than design. Analysis of these results has indicated the deficit in performance to be due to stator secondary flow losses, vaneless space surface friction losses, and trailing edge wake mixing losses.

Analytic investigation of helicopter rotor blade appended aeroelastic devices

NASA Technical Reports Server (NTRS)

Bielawa, Richard L.

1984-01-01

Analytic evaluations of four different passive aeroelastic devices appended to helicopter rotor blades are presented. The devices consist of a passive tuned tab, a control coupled tab, an all-flying tip and a harmonic dilational airfoil tip. Each device was conceived for improving either aerodynamic performance or reducing vibratory control loads or hub shears. The evaluation was performed using a comprehensive rotor aeroelastic analysis (the G400PA code with appropriate modifications), together with data for a realistic helicopter rotor blade (the UH-60A Blackhawk), in high speed flight (90 m/s, 175 kts). The results of this study show that significant performance (L/(D sub e)) gains can be achieved with the all-flying free tip. Results from the harmonic dilational airfoil tip show the potential for moderate improvements in L/(D sub e). Finally, the results for the passive tuned tab and the control coupled tab, as configured for this study, show these devices to be impractical. Sections are included which describe the operation of each device, the required G400PA modifications, and the detailed results obtained for each device.

Effect of chord-to-diameter ratio on vertical-axis wind turbine wake development

NASA Astrophysics Data System (ADS)

Parker, Colin M.; Araya, Daniel B.; Leftwich, Megan C.

2017-12-01

The wake structure of a vertical-axis wind turbine (VAWT) is strongly dependent on the tip-speed ratio, λ, or the tangential speed of the turbine blade relative to the incoming wind speed. The geometry of a turbine can influence λ, but the precise relationship among VAWT geometric parameters and VAWT wake characteristics remains unknown. To investigate this relationship, we present the results of an experiment to characterize the wakes of three VAWTs that are geometrically similar except for the ratio of the turbine diameter ( D), to blade chord ( c), which was chosen to be D/c = 3, 6, and 9. For a fixed freestream Reynolds number based on the blade chord of Re_c = 1.6× 10^3, both two-component particle image velocimetry (PIV) and single-component hot-wire anemometer measurements are taken at the horizontal mid-plane in the wake of each turbine. PIV measurements are ensemble averaged in time and phase averaged with each rotation of the turbine. Hot-wire measurement points are selected to coincide with the edge of the shear layer of each turbine wake, as deduced from the PIV data, which allows for an analysis of the frequency content of the wake due to vortex shedding by the turbine.

Application of chromatic confocal displacement sensor in measurement of tip clearance

NASA Astrophysics Data System (ADS)

Bi, Chao; Li, Di; Fang, Jianguo; Zhang, Bin

2016-10-01

In the field of aeronautics, the tip clearance of rotor exerts a crucial influence on the performance of the aero engine. As defined as the radial distance between the top of the blade and the inner wall of the casing, the tip clearance of too large or small size will adversely affect the normal running of the engine. In order to realize accurate measurement of the tip clearance in a simple way, a non-contact measuring method by the chromatic confocal displacement sensor is proposed in the paper. The sensor possesses the advantages such as small volume, good signal-to-noise ratio, high accuracy and response frequency etc., which make it be widely used in engineering and industry. For testing the performance and potential application of the sensor, a simulation testing platform is established. In the platform, a simulation blisk is installed on the air bearing spindle and a chromatic confocal displacement sensor is fixed on the platform to measure the displacement variation of the blade tip, which can be used to characterize the variation of the tip clearance. In the simulation experiments, both of single and continuous measurement of the tip clearance of the 36 blades on the blisk is executed. As the results of experiments show, the chromatic confocal displacement sensor can meet the requirements of measuring task, in which both of high measuring efficiency and accuracy could be achieved. Therefore, the measuring method proposed in the paper can be utilized in the actual assembling sites of the aero engine.

High-loading low-speed fan study. 4: Data and performance with redesign stator and including a rotor tip casing treatment

NASA Technical Reports Server (NTRS)

Harley, K. G.; Odegard, P. A.; Burdsall, E. A.

1972-01-01

A single stage fan with a rotor tip speed of 1000 ft/sec(304.8 m/sec) and a hub-to-tip ratio of 0.392 was retested with a redesigned stator. Tests were conducted with uniform inlet, tip-radial, hub-radial, and circumferential inlet distortions. With uniform inlet flow, stall margin was improved 12 percentage points above that with the original stator. The fan demonstrated an efficiency of 0.883 and a stall margin of 15 percent at a pressure ratio of 1.488 and a specific flow of 41.17 lb/sec/sq ft. Tests were also made with a redesigned casing treatment consisting of skewed slots over the rotor blade tips. This casing treatment gave a 7 percentage point improvement in stall margin when tested with tip radial distortion (when the rotor tip initiated stall). Noise measurements at the fan inlet and exit indicate no effect from closing the stator 10 degrees, nor were there measurable effects from adding skewed slots over the blade tips.

In-place recalibration technique applied to a capacitance-type system for measuring rotor blade tip clearance

NASA Technical Reports Server (NTRS)

Barranger, J. P.

1978-01-01

The rotor blade tip clearance measurement system consists of a capacitance sensing probe with self contained tuning elements, a connecting coaxial cable, and remotely located electronics. Tests show that the accuracy of the system suffers from a strong dependence on probe tip temperature and humidity. A novel inplace recalibration technique was presented which partly overcomes this problem through a simple modification of the electronics that permits a scale factor correction. This technique, when applied to a commercial system significantly reduced errors under varying conditions of humidity and temperature. Equations were also found that characterize the important cable and probe design quantities.

Turbine airfoil fabricated from tapered extrusions

DOEpatents

Marra, John J

2013-07-16

An airfoil (30) and fabrication process for turbine blades with cooling channels (26). Tapered tubes (32A-32D) are bonded together in a parallel sequence, forming a leading edge (21), a trailing edge (22), and pressure and suction side walls (23, 24) connected by internal ribs (25). The tapered tubes may be extruded without camber to simplify the extrusion process, then bonded along matching surfaces (34), forming a non-cambered airfoil (28), which may be cambered in a hot forming process and cut (48) to length. The tubes may have tapered walls that are thinner at the blade tip (T1) than at the base (T2), reducing mass. A cap (50) may be attached to the blade tip. A mounting lug (58) may be forged (60) on the airfoil base and then machined, completing the blade for mounting in a turbine rotor disk.

Cost/benefit studies of advanced materials technologies for future aircraft turbine engines: Materials for advanced turbine engines

NASA Technical Reports Server (NTRS)

Stearns, M.; Wilbers, L.

1982-01-01

Cost benefit studies were conducted on six advanced materials and processes technologies applicable to commercial engines planned for production in the 1985 to 1990 time frame. These technologies consisted of thermal barrier coatings for combustor and high pressure turbine airfoils, directionally solidified eutectic high pressure turbine blades, (both cast and fabricated), and mixers, tail cones, and piping made of titanium-aluminum alloys. A fabricated titanium fan blisk, an advanced turbine disk alloy with improved low cycle fatigue life, and a long-life high pressure turbine blade abrasive tip and ceramic shroud system were also analyzed. Technologies showing considerable promise as to benefits, low development costs, and high probability of success were thermal barrier coating, directionally solidified eutectic turbine blades, and abrasive-tip blades/ceramic-shroud turbine systems.

Study of noise sources in a subsonic fan using measured blade pressures and acoustic theory

NASA Technical Reports Server (NTRS)

Hanson, D. B.

1975-01-01

Sources of noise in a 1.4 m (4.6 ft) diameter subsonic tip speed propulsive fan running statically outdoors are studied using a combination of techniques. Signals measured with pressure transducers on a rotor blade are plotted in a format showing the space-time history of inlet distortion. Study of these plots visually and with statistical correlation analysis confirms that the inlet flow contains long, thin eddies of turbulence. Turbulence generated in the boundary layer of the shroud upstream of the rotor tips was not found to be an important noise source. Fan noise is diagnosed by computing narrowband spectra of rotor and stator sound power and comparing these with measured sound power spectra. Rotor noise is computed from spectra of the measured blade pressures and stator noise is computed using the author's stator noise theory. It is concluded that the rotor and stator sources contribute about equally at frequencies in the vicinity of the first three harmonics of blade passing frequency. At higher frequencies, the stator contribution diminishes rapidly and the rotor/inlet turbulence mechanism dominates. Two parametric studies are performed by using the rotor noise calculation procedure which was correlated with test. In the first study, the effects on noise spectrum and directivity are calculated for changes in turbulence properties, rotational Mach number, number of blades, and stagger angle. In the second study the influences of design tip speed and blade number on noise are evaluated.

Testing of a Microwave Blade Tip Clearance Sensor at the NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Woike, Mark R.; Roeder, James W.; Hughes, Christopher E.; Bencic, Timothy J.

2009-01-01

The development of new active tip clearance control and structural health monitoring schemes in turbine engines and other types of rotating machinery requires sensors that are highly accurate and can operate in a high-temperature environment. The use of a microwave sensor to acquire blade tip clearance and tip timing measurements is being explored at the NASA Glenn Research Center. The microwave blade tip clearance sensor works on principles that are very similar to a short-range radar system. The sensor sends a continuous microwave signal towards a target and measures the reflected signal. The phase difference of the reflected signal is directly proportional to the distance between the sensor and the target being measured. This type of sensor is beneficial in that it has the ability to operate at extremely high temperatures and is unaffected by contaminants that may be present in turbine engines. The use of microwave sensors for this application is a new concept. Techniques on calibrating the sensors along with installation effects are not well quantified as they are for other sensor technologies. Developing calibration techniques and evaluating installation effects are essential in using these sensors to make tip clearance and tip timing measurements. As a means of better understanding these issues, the microwave sensors were used on a benchtop calibration rig, a large axial vane fan, and a turbofan. Background on the microwave tip clearance sensor, an overview of their calibration, and the results from their use on the axial vane fan and the turbofan will be presented in this paper.

Testing of a Microwave Blade Tip Clearance Sensor at the NASA Glenn Research Center

NASA Technical Reports Server (NTRS)

Woike, Mark R.; Roeder, James W.; Hughes, Christopher E.; Bencic, Timothy J.

2009-01-01

The development of new active tip clearance control and structural health monitoring schemes in turbine engines and other types of rotating machinery requires sensors that are highly accurate and can operate in a high temperature environment. The use of a microwave sensor to acquire blade tip clearance and tip timing measurements is being explored at the NASA Glenn Research Center. The microwave blade tip clearance sensor works on principles that are very similar to a short range radar system. The sensor sends a continuous microwave signal towards a target and measures the reflected signal. The phase difference of the reflected signal is directly proportional to the distance between the sensor and the target being measured. This type of sensor is beneficial in that it has the ability to operate at extremely high temperatures and is unaffected by contaminants that may be present in turbine engines. The use of microwave sensors for this application is a new concept. Techniques on calibrating the sensors along with installation effects are not well quantified as they are for other sensor technologies. Developing calibration techniques and evaluating installation effects are essential in using these sensors to make tip clearance and tip timing measurements. As a means of better understanding these issues, the microwave sensors were used on a bench top calibration rig, a large axial vane fan, and a turbofan. Background on the microwave tip clearance sensor, an overview of their calibration, and the results from their use on the axial vane fan and the turbofan will be presented in this paper.

Numerical simulation on a straight-bladed vertical axis wind turbine with auxiliary blade

NASA Astrophysics Data System (ADS)

Li, Y.; Zheng, Y. F.; Feng, F.; He, Q. B.; Wang, N. X.

2016-08-01

To improve the starting performance of the straight-bladed vertical axis wind turbine (SB-VAWT) at low wind speed, and the output characteristics at high wind speed, a flexible, scalable auxiliary vane mechanism was designed and installed into the rotor of SB-VAWT in this study. This new vertical axis wind turbine is a kind of lift-to-drag combination wind turbine. The flexible blade expanded, and the driving force of the wind turbines comes mainly from drag at low rotational speed. On the other hand, the flexible blade is retracted at higher speed, and the driving force is primarily from a lift. To research the effects of the flexible, scalable auxiliary module on the performance of SB-VAWT and to find its best parameters, the computational fluid dynamics (CFD) numerical calculation was carried out. The calculation result shows that the flexible, scalable blades can automatic expand and retract with the rotational speed. The moment coefficient at low tip speed ratio increased substantially. Meanwhile, the moment coefficient has also been improved at high tip speed ratios in certain ranges.

Annulus wall boundary layer development in a compressor stage, including the effects of tip clearance

NASA Technical Reports Server (NTRS)

Lakshminarayana, B.; Murthy, K. N. S.; Pouagare, M.; Govindan, T. R.

1983-01-01

The end-wall boundary layer development in a compressor stage, including the inlet guide vane (IGV) passage and the rotor passage, was measured. The measurement upstream of the rotor and inside the IGV passage were carried out with a five-hole probe. The data (blade-to-blade) inside the IGV passage were carried out with a five-hole probe. The data (blade-to-blade) inside the rotor passage were measured using a three-sensor rotating hot-wire below the tip clearance region and "V' configuration probe inside the clearance region. The rotor exit measurements (blade-to-blade) were acquired with a laser Doppler velocimeter. The velocity profiles and the integral properties are presented and interpreted. The boundary layer is comparatively well behaved up to the leading edge of the rotor, beyond which complex interactions result in very unconventional profiles. The momentum thicknesses decrease in the leakage flow region of the rotor. The momentum thicknesses and the limiting streamline angles predicted from a momentum integral technique agree well with the data up to the leading edge of the rotor.

Cavitation, Flow Structure and Turbulence in the Tip Region of a Rotor Blade

NASA Technical Reports Server (NTRS)

Wu, H.; Miorini, R.; Soranna, F.; Katz, J.; Michael, T.; Jessup, S.

2010-01-01

Objectives: Measure the flow structure and turbulence within a Naval, axial waterjet pump. Create a database for benchmarking and validation of parallel computational efforts. Address flow and turbulence modeling issues that are unique to this complex environment. Measure and model flow phenomena affecting cavitation within the pump and its effect on pump performance. This presentation focuses on cavitation phenomena and associated flow structure in the tip region of a rotor blade.

Effect of the Macintosh curved blade size on direct laryngoscopic view in edentulous patients.

PubMed

Kim, Hyerim; Chang, Jee-Eun; Han, Sung-Hee; Lee, Jung-Man; Yoon, Soohyuk; Hwang, Jin-Young

2018-01-01

In the present study, we compared the laryngoscopic view depending on the size of the Macintosh curved blade in edentulous patients. Thirty-five edentulous adult patients scheduled for elective surgery were included in the study. After induction of anesthesia, two direct laryngoscopies were performed alternately using a standard-sized Macintosh curved blade (No. 4 for men and No. 3 for women) and smaller-sized Macintosh curved blade (No. 3 for men and No. 2 for women). During direct laryngoscopy with each blade, two digital photographs of the lateral view were taken when the blade tip was placed in the valleculae; the laryngoscope was lifted to achieve the best laryngeal view. Then, the best laryngeal views were assessed using the percentage of glottic opening (POGO) score. On the photographs of the lateral view of direct laryngoscopy, the angles between the line extending along the laryngoscopic handle and the horizontal line were measured. The POGO score was improved with the smaller-sized blade compared with the standard-sized blade (87.3% [11.8%] vs. 71.3% [20.0%], P<0.001, respectively). The angles between the laryngoscopic handle and the horizontal line were greater with the smaller-sized blade compared to the standard-sized blade when the blade tip was placed on the valleculae and when the laryngoscope was lifted to achieve the best laryngeal view (both P<0.001). Compared to a standard-sized Macintosh blade, a smaller-sized Macintosh curved blade improved the laryngeal exposure in edentulous patients. Copyright © 2017 Elsevier Inc. All rights reserved.

Probabilistic Structural Analysis of SSME Turbopump Blades: Probabilistic Geometry Effects

NASA Technical Reports Server (NTRS)

Nagpal, V. K.

1985-01-01

A probabilistic study was initiated to evaluate the precisions of the geometric and material properties tolerances on the structural response of turbopump blades. To complete this study, a number of important probabilistic variables were identified which are conceived to affect the structural response of the blade. In addition, a methodology was developed to statistically quantify the influence of these probabilistic variables in an optimized way. The identified variables include random geometric and material properties perturbations, different loadings and a probabilistic combination of these loadings. Influences of these probabilistic variables are planned to be quantified by evaluating the blade structural response. Studies of the geometric perturbations were conducted for a flat plate geometry as well as for a space shuttle main engine blade geometry using a special purpose code which uses the finite element approach. Analyses indicate that the variances of the perturbations about given mean values have significant influence on the response.

Noise Exposure in TKA Surgery; Oscillating Tip Saw Systems vs Oscillating Blade Saw Systems.

PubMed

Peters, Michiel P; Feczko, Peter Z; Tsang, Karel; van Rietbergen, Bert; Arts, Jacobus J; Emans, Peter J

2016-12-01

Historically it has been suggested that noise-induced hearing loss (NIHL) affects approximately 50% of the orthopedic surgery personnel. This noise may be partially caused by the use of powered saw systems that are used to make the bone cuts. The first goal was to quantify and compare the noise emission of these different saw systems during total knee arthroplasty (TKA) surgery. A second goal was to estimate the occupational NIHL risk for the orthopedic surgery personnel in TKA surgery by quantifying the total daily noise emission spectrum during TKA surgery and to compare this to the Dutch Occupational Health Organization guidelines. A conventional sagittal oscillating blade system with a full oscillating blade and 2 newer oscillating tip saw systems (handpiece and blade) were compared. Noise level measurements during TKA surgery were performed during cutting and hammering, additionally surgery noise profiles were made. The noise level was significantly lower for the oscillating tip saw systems compared to the conventional saw system, but all were in a range that can cause NIHL. The conventional system handpiece produced a considerable higher noise level compared to oscillating tip handpiece. NIHL is an underestimated problem in the orthopedic surgery. Solutions for decreasing the risk of hearing loss should be considered. The use of oscillating tip saw systems have a reduced noise emission in comparison with the conventional saw system. The use of these newer systems might be a first step in decreasing hearing loss among the orthopedic surgery personnel. Copyright © 2016 Elsevier Inc. All rights reserved.

Research on the nonintrusive measurement of the turbine blade vibration

NASA Astrophysics Data System (ADS)

Zhang, Shi hai; Li, Lu-ping; Rao, Hong-de

2008-11-01

It's one of the important ways to monitor the change of dynamic characteristic of turbine blades for ensuring safety operation of turbine unit. Traditional measurement systems for monitoring blade vibration generally use strain gauges attached to the surface of turbine blades, each strain gauge gives out an analogue signal related to blade deformation, it's maximal defect is only a few blades could be monitored which are attached by strain gauge. But the noncontact vibration measurement will be discussed would solve this problem. This paper deals with noncontact vibration measurement on the rotor blades of turbine through experiments. In this paper, the noncontact vibration measurement - Tip Timing Measurement will be presented, and will be improved. The statistics and DFT will be used in the improved measurement. The main advantage of the improved measurement is that only two sensors over the top of blades and one synchronous sensor of the rotor are used to get the exact vibration characteristics of the each blade in a row. In our experiment, we adopt NI Company's DAQ equipment: SCXI1001 and PCI 6221, three optical sensors, base on the graphics program soft LabVIEW to develop the turbine blade monitor system. At the different rotational speed of the rotor (1000r/m and 1200r/m) we do several experiments on the bench of the Turbine characteristic. Its results indicated that the vibration of turbine blade could be real-time monitored and accurately measured by the improved Tip Timing Measurement.

Rotorcraft Brownout Advanced Understanding, Control, and Mitigation

DTIC Science & Technology

2014-10-31

rotor disk loading , blade loading , number and placement of rotors, number of blades, blade twist, blade tip shape, fuselage shape, as well as...Mechanical Engineering • Ramani Duraiswami, Ph.D., Associate Professor, Department of Computer Science & Insti- tute for Advanced Computer Studies • Nail ...23, 2013. 71. Mulinti, R., Corfman, K., and Kiger, K. T., “Particle-Turbulence Interaction of Suspended Load by Forced Jet Impinging on a Mobile

Vibrations of a Marine Propeller Operating in a Nonuniform Inflow.

DTIC Science & Technology

1980-04-01

Expanded Blade Midsurface ......... ........................ ... 73 16 - Calculated Normalized Propeller RMS Vibration Velocity as a Function of...averaged over the blade midsurface ), rather thaft the maximum velocities near the blade tip. Then, for the two test propellers, the rms nonuniform inflow...time- averaged midsurface of the blade, then the instantaneous position S of the vibrating midsurface is _S (ric)+ qct S(r,c,t) = (rc) + q(t) i(rc

Sparse Representation Based Frequency Detection and Uncertainty Reduction in Blade Tip Timing Measurement for Multi-Mode Blade Vibration Monitoring

PubMed Central

Pan, Minghao; Yang, Yongmin; Guan, Fengjiao; Hu, Haifeng; Xu, Hailong

2017-01-01

The accurate monitoring of blade vibration under operating conditions is essential in turbo-machinery testing. Blade tip timing (BTT) is a promising non-contact technique for the measurement of blade vibrations. However, the BTT sampling data are inherently under-sampled and contaminated with several measurement uncertainties. How to recover frequency spectra of blade vibrations though processing these under-sampled biased signals is a bottleneck problem. A novel method of BTT signal processing for alleviating measurement uncertainties in recovery of multi-mode blade vibration frequency spectrum is proposed in this paper. The method can be divided into four phases. First, a single measurement vector model is built by exploiting that the blade vibration signals are sparse in frequency spectra. Secondly, the uniqueness of the nonnegative sparse solution is studied to achieve the vibration frequency spectrum. Thirdly, typical sources of BTT measurement uncertainties are quantitatively analyzed. Finally, an improved vibration frequency spectra recovery method is proposed to get a guaranteed level of sparse solution when measurement results are biased. Simulations and experiments are performed to prove the feasibility of the proposed method. The most outstanding advantage is that this method can prevent the recovered multi-mode vibration spectra from being affected by BTT measurement uncertainties without increasing the probe number. PMID:28758952

Performance of a 1.15-pressure-ratio axial-flow fan stage with a blade tip solidity of 0.5

NASA Technical Reports Server (NTRS)

Osborn, W. M.; Steinke, R. J.

1974-01-01

The overall and blade-element performance of a low-solidity, low-pressure-ratio, low-tip-speed fan stage is presented over the stable operating range at rotative speeds from 90 to 120 percent of design speed. At design speed a stage peak efficiency of 0.836 was obtained at a weight flow of 30.27 kilograms per second and a pressure ratio of 1.111. The pressure ratio was less than design pressure ratio, and the design energy input into the rotor was not achieved. A mismatch of the rotor and stator blade elements resulted due to the lower than design pressure ratio of the rotor.

Tip Vortex and Wake Characteristics of a Counterrotating Open Rotor

NASA Technical Reports Server (NTRS)

VanZante, Dale E.; Wernet, Mark P.

2012-01-01

One of the primary noise sources for Open Rotor systems is the interaction of the forward rotor tip vortex and blade wake with the aft rotor. NASA has collaborated with General Electric on the testing of a new generation of low noise, counterrotating Open Rotor systems. Three-dimensional particle image velocimetry measurements were acquired in the intra-rotor gap of the Historical Baseline blade set. The velocity measurements are of sufficient resolution to characterize the tip vortex size and trajectory as well as the rotor wake decay and turbulence character. The tip clearance vortex trajectory is compared to results from previously developed models. Forward rotor wake velocity profiles are shown. Results are presented in a form as to assist numerical modeling of Open Rotor system aerodynamics and acoustics.

Analysis of middle bearing failure in rotor jet engine using tip-timing and tip-clearance techniques

NASA Astrophysics Data System (ADS)

Rzadkowski, R.; Rokicki, E.; Piechowski, L.; Szczepanik, R.

2016-08-01

The reported problem is the failure of the middle bearing in an aircraft rotor engine. Tip-timing and tip-clearance and variance analyses are carried out on a compressor rotor blade in the seventh stage above the middle bearing. The experimental analyses concern both an aircraft engine with a middle bearing in good working order and an engine with a damaged middle bearing. A numerical analysis of seventh stage blade free vibration is conducted to explain the experimental results. This appears to be an effective method of predicting middle bearing failure. The results show that variance first increases in the initial stages of bearing failure, but then starts to decrease and stabilize, and then again decrease shortly before complete bearing failure.

On vortex-airfoil interaction noise including span-end effects, with application to open-rotor aeroacoustics

NASA Astrophysics Data System (ADS)

Roger, Michel; Schram, Christophe; Moreau, Stéphane

2014-01-01

A linear analytical model is developed for the chopping of a cylindrical vortex by a flat-plate airfoil, with or without a span-end effect. The major interest is the contribution of the tip-vortex produced by an upstream rotating blade in the rotor-rotor interaction noise mechanism of counter-rotating open rotors. Therefore the interaction is primarily addressed in an annular strip of limited spanwise extent bounding the impinged blade segment, and the unwrapped strip is described in Cartesian coordinates. The study also addresses the interaction of a propeller wake with a downstream wing or empennage. Cylindrical vortices are considered, for which the velocity field is expanded in two-dimensional gusts in the reference frame of the airfoil. For each gust the response of the airfoil is derived, first ignoring the effect of the span end, assimilating the airfoil to a rigid flat plate, with or without sweep. The corresponding unsteady lift acts as a distribution of acoustic dipoles, and the radiated sound is obtained from a radiation integral over the actual extent of the airfoil. In the case of tip-vortex interaction noise in CRORs the acoustic signature is determined for vortex trajectories passing beyond, exactly at and below the tip radius of the impinged blade segment, in a reference frame attached to the segment. In a second step the same problem is readdressed accounting for the effect of span end on the aerodynamic response of a blade tip. This is achieved through a composite two-directional Schwarzschild's technique. The modifications of the distributed unsteady lift and of the radiated sound are discussed. The chained source and radiation models provide physical insight into the mechanism of vortex chopping by a blade tip in free field. They allow assessing the acoustic benefit of clipping the rear rotor in a counter-rotating open-rotor architecture.

Striation patterns in serrated blade stabs to cartilage.

PubMed

Pounder, Derrick J; Reeder, Francesca D

2011-05-20

Stab wounds were made in porcine cartilage with 13 serrated knives, amongst which 4 were drop-point and 9 straight-spine; 9 coarsely serrated, 3 finely serrated and 1 with mixed pattern serrations. The walls of the stab tracks were cast with dental impression material, and the casts photographed together with the knife blades for comparison. All 13 serrated blades produced an "irregularly regular" pattern of striations on cartilage in all stabbings. Unusual and distinctive blade serration patterns produced equally distinctive wound striation patterns. A reference collection of striation patterns and corresponding blades might prove useful for striation pattern analysis. Drop-point blades produced similar striations to straight-spine blades except that the striations were not parallel but rather fan-shaped, converging towards the wound exit. The fan-shaped striation pattern characteristic of drop-point blades is explained by the initial lateral movement of the blade through the cartilage imposed by the presence of the drop point shape. It appears that the greater the overall angle of the drop point, the shorter the blade length over which the drop point occurs, and the closer the first serration is to the knife tip, the more obvious is the fan-shaped pattern. We anticipate that micro-irregularities producing individualising characteristics in non-serrated drop point blades, provided they were located at the tip opposite the drop point, should also show a fan-shaped pattern indicative of a drop point blade. The examination of the walls of stab wounds to cartilage represents an under-utilised source of forensic information to assist in knife identification. Copyright © 2010 Elsevier Ireland Ltd. All rights reserved.

Fully Suspended, Five-Axis, Three-Magnetic-Bearing Dynamic Spin Rig With Forced Excitation

NASA Technical Reports Server (NTRS)

Morrison, Carlos R.; Provenza, Andrew; Kurkov, Anatole; Montague, Gerald; Duffy, Kirsten; Mehmed, Oral; Johnson, Dexter; Jansen, Ralph

2004-01-01

The Five-Axis, Three-Magnetic-Bearing Dynamic Spin Rig, a significant advancement in the Dynamic Spin Rig (DSR), is used to perform vibration tests of turbomachinery blades and components under rotating and nonrotating conditions in a vacuum. The rig has as its critical components three magnetic bearings: two heteropolar radial active magnetic bearings and a magnetic thrust bearing. The bearing configuration allows full vertical rotor magnetic suspension along with a feed-forward control feature, which will enable the excitation of various natural blade modes in bladed disk test articles. The theoretical, mechanical, electrical, and electronic aspects of the rig are discussed. Also presented are the forced-excitation results of a fully levitated, rotating and nonrotating, unbladed rotor and a fully levitated, rotating and nonrotating, bladed rotor in which a pair of blades was arranged 180 degrees apart from each other. These tests include the bounce mode excitation of the rotor in which the rotor was excited at the blade natural frequency of 144 Hz. The rotor natural mode frequency of 355 Hz was discerned from the plot of acceleration versus frequency. For nonrotating blades, a blade-tip excitation amplitude of approximately 100 g/A was achieved at the first-bending critical (approximately 144 Hz) and at the first-torsional and second-bending blade modes. A blade-tip displacement of 70 mils was achieved at the first-bending critical by exciting the blades at a forced-excitation phase angle of 908 relative to the vertical plane containing the blades while simultaneously rotating the shaft at 3000 rpm.

Compressor blade clearance measurement using capacitance and phase lock techniques

NASA Astrophysics Data System (ADS)

Demers, Rosario N.

1986-11-01

The clearance measurement system has several unique features which mimimize problems plaguing earlier systems. These include tuning stability and sensitivity drift. Both these problems are intensified by the environmental factors present in compressors i.e., wide temperature fluctuations, vibrations, and conductive contamination of probe tips. The circuitry in this new system provides phase lock feedback to control tuning and shut calibration to measure sensitivity. The use of high frequency excitation lowers the probe tip impedance, thus miminizing the effects of contamination. A prototype has been built and tested. The ability to calibrate has been demonstrated. An eight channel system is now being constructed for use in the Compressor Research Facility at Wright-Patterson AFB. The efficiency of a turbine engine is to a large extent dependent upon the mechanical tolerances maintained between its moving parts. On critical tolerance is the blade span. Although this tolerance may not appear severe, the impact on compressor efficiency is dramatic. The penalty in percent efficiency has been shown to be three times the percent clearance to blade span ratio. In addition, each percent loss in compressor efficiency represents one half percent loss in specific fuel consumption. Factors which affect blade tip clearance are identified.

Effects of Alternate Leading Edge Cutback on the Space Shuttle Main Engine Low Pressure Fuel Pump

NASA Technical Reports Server (NTRS)

Mulder, Andrew; Skelley, Stephen

2016-01-01

A higher order cavitation oscillation observed in the SSME low pressure fuel pump has been eliminated in water flow testing of a modified subscale replica of the inducer. The low pressure pump was modified by removing the outboard sections of two opposing blades of the four-bladed inducer, blending the "cutback" regions into the blades at the leading edge and tip, and removing material on the suction sides to decrease the exposed leading edge thickness. The leading edge tips of the cutback blades were moved approximately 25 degrees from their previous locations, thereby increasing one blade to blade spacing, decreasing the second, while simultaneously moving the cutback tips downstream. The test was conducted in MSFC's inducer test loop at scaled operating conditions in degassed and filtered water. In addition to eliminating HOC across the entire scaled operating regime, rotating cavitation was suppressed while the range of both alternate blade and asymmetric cavitation were increased. These latter phenomena, and more significantly, the shifts between these cavitation modes also resulted in significant changes to the head coefficient at low cavitation numbers. Reverse flow was detected at a slightly larger flow coefficient with the cutback inducer and suction capability was reduced by approximately 1 velocity head at and above approximately 90% of the reference flow coefficient. These performance changes along with more intense reverse flow are consistent with poor flow area management and increased incidence in the cutback region. Although the test demonstrated that the inducer modification was successful at eliminating the higher order cavitation across the entire scaled operating regime, different, previously unobserved, cavitation oscillations were introduced and significant performance penalties were imposed.

Redesigned rotor for a highly loaded, 1800 ft/sec tip speed compressor fan stage 1: Aerodynamic and mechanical design

NASA Technical Reports Server (NTRS)

Halle, J. E.; Ruschak, J. T.

1975-01-01

A highly loaded, high tip-speed fan rotor was designed with multiple-circular-arc airfoil sections as a replacement for a marginally successful rotor which had precompression airfoil sections. The substitution of airfoil sections was the only aerodynamic change. Structural design of the redesigned rotor blade was guided by successful experience with the original blade. Calculated stress levels and stability parameters for the redesigned rotor are within limits demonstrated in tests of the original rotor.

Observations of tip vortex cavitation inception from a model marine propeller

NASA Astrophysics Data System (ADS)

Lodha, R. K.; Arakeri, V. H.

1984-01-01

Cavitation inception characteristics of a model marine propeller having three blades, developed area ratio of 0.34 and at three different pitch to diameter ratios of 0.62, 0.83 and 1.0 are reported. The dominant type of cavitation observed at inception was the tip vortex type. The measured magnitude of inception index is found to agree well with a proposed correlation due to Strasberg. Performance calculations of the propeller based on combined vortex and blade element theory are also presented.

Heat Transfer on a Film-Cooled Rotating Blade

NASA Technical Reports Server (NTRS)

Garg, Vijay K.

1999-01-01

A multi-block, three-dimensional Navier-Stokes code has been used to compute heat transfer coefficient on the blade, hub and shroud for a rotating high-pressure turbine blade with 172 film-cooling holes in eight rows. Film cooling effectiveness is also computed on the adiabatic blade. Wilcox's k-omega model is used for modeling the turbulence. Of the eight rows of holes, three are staggered on the shower-head with compound-angled holes. With so many holes on the blade it was somewhat of a challenge to get a good quality grid on and around the blade and in the tip clearance region. The final multi-block grid consists of 4784 elementary blocks which were merged into 276 super blocks. The viscous grid has over 2.2 million cells. Each hole exit, in its true oval shape, has 80 cells within it so that coolant velocity, temperature, k and omega distributions can be specified at these hole exits. It is found that for the given parameters, heat transfer coefficient on the cooled, isothermal blade is highest in the leading edge region and in the tip region. Also, the effectiveness over the cooled, adiabatic blade is the lowest in these regions. Results for an uncooled blade are also shown, providing a direct comparison with those for the cooled blade. Also, the heat transfer coefficient is much higher on the shroud as compared to that on the hub for both the cooled and the uncooled cases.

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 7: Data and performance for stage E

NASA Technical Reports Server (NTRS)

Cheatham, J. G.

1974-01-01

An axial flow compressor stage, having tandem airfoil blading, was designed for zero rotor prewhirl, constant rotor work across the span, and axial discharge flow. The stage was designed to produce a pressure ratio of 1.265 at a rotor tip velocity of 757 ft/sec. The rotor has an inlet hub/tip ratio of 0.8. The design procedure accounted for the rotor inlet boundary layer and included the effects of axial velocity ratio and secondary flow on blade row performance. The objectives of this experimental program were (1) to obtain performance with uniform and distorted inlet flow for comparison with the performance of a stage consisting of single-airfoil blading designed for the same vector diagrams and (2) to evaluate the effectiveness of accounting for the inlet boundary layer, axial velocity ratio, and secondary flows in the stage design.

Analysis of the aerodynamic performance of the multi-rotor concept

NASA Astrophysics Data System (ADS)

Chasapogiannis, Petros; Prospathopoulos, John M.; Voutsinas, Spyros G.; Chaviaropoulos, Takis K.

2014-06-01

The concept of a large (~20MW) multi-rotor wind turbine intended for offshore installations is analysed with respect to its aerodynamic performance. The effect of closely clustering rotors on a single actuator disk is estimated using two different modelling approaches: a CFD solver in which the rotors are simulated as distinct actuator disks and a vortex based solver in which the blade geometry is exactly considered. In the present work, a system of 7 rotors is simulated with a centre to centre spacing of 1.05D. At nominal conditions (tip speed ratio=9) both models predict an increase in power of ~3% alongside with an increase in thrust of ~1.5%. The analysis of the flow field indicates that in the 7 rotor system the individual wakes merge into one wake at ~2D and that flow recovery starts at approximately the same downstream distance as in the single rotor case. As regards the dynamic implications of the close spacing of the rotors it was found that there is an increase in the loading amplitude ranging from 0.30-2.13% at blade level in rated conditions.

Strength and dynamic characteristics analyses of wound composite axial impeller

NASA Astrophysics Data System (ADS)

Wang, Jifeng; Olortegui-Yume, Jorge; Müller, Norbert

2012-03-01

A low cost, light weight, high performance composite material turbomachinery impeller with a uniquely designed blade patterns is analyzed. Such impellers can economically enable refrigeration plants to use water as a refrigerant (R718). A strength and dynamic characteristics analyses procedure is developed to assess the maximum stresses and natural frequencies of these wound composite axial impellers under operating loading conditions. Numerical simulation using FEM for two-dimensional and three-dimensional impellers was investigated. A commercially available software ANSYS is used for the finite element calculations. Analysis is done for different blade geometries and then suggestions are made for optimum design parameters. In order to avoid operating at resonance, which can make impellers suffer a significant reduction in the design life, the designer must calculate the natural frequency and modal shape of the impeller to analyze the dynamic characteristics. The results show that using composite Kevlar fiber/epoxy matrix enables the impeller to run at high tip speed and withstand the stresses, no critical speed will be matched during start-up and shut-down, and that mass imbalances of the impeller shall not pose a critical problem.

Computational Aerodynamic Simulations of a 1484 ft/sec Tip Speed Quiet High-Speed Fan System Model for Acoustic Methods Assessment and Development

NASA Technical Reports Server (NTRS)

Tweedt, Daniel L.

2014-01-01

Computational Aerodynamic simulations of a 1484 ft/sec tip speed quiet high-speed fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, low-noise research fan/nacelle model that has undergone experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating points simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, which includes a core duct and a bypass duct that merge upstream of the fan system nozzle. As a result, only fan rotational speed and the system bypass ratio, set by means of a translating nozzle plug, were adjusted in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. Computed blade row flow fields at all fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow fields reveals no excessive or critical boundary layer separations or related secondary-flow problems, with the exception of the hub boundary layer at the core duct entrance. At that location a significant flow separation is present. The region of local flow recirculation extends through a mixing plane, however, which for the particular mixing-plane model used is now known to exaggerate the recirculation. In any case, the flow separation has relatively little impact on the computed rotor and FEGV flow fields.

Computational Aerodynamic Simulations of an 840 ft/sec Tip Speed Advanced Ducted Propulsor Fan System Model for Acoustic Methods Assessment and Development

NASA Technical Reports Server (NTRS)

Tweedt, Daniel L.

2014-01-01

Computational Aerodynamic simulations of an 840 ft/sec tip speed, Advanced Ducted Propulsor fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, lownoise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15- foot Low Speed Wind Tunnel at the NASA Glenn Research Center, resulting in quality, detailed aerodynamic and acoustic measurement data. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating conditions simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, excluding a long core duct section downstream of the core inlet guide vane. As a result, only fan rotational speed and system bypass ratio, set by specifying static pressure downstream of the core inlet guide vane row, were adjusted in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. The computed blade row flow fields for all five fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow fields reveals no excessive boundary layer separations or related secondary-flow problems. A few spanwise comparisons between computational and measurement data in the bypass duct show that they are in good agreement, thus providing a partial validation of the computational results.

Tip-path-plane angle effects on rotor blade-vortex interaction noise levels and directivity

NASA Technical Reports Server (NTRS)

Burley, Casey L.; Martin, Ruth M.

1988-01-01

Acoustic data of a scale model BO-105 main rotor acquired in a large aeroacoustic wind tunnel are presented to investigate the parametric effects of rotor operating conditions on blade-vortex interaction (BVI) impulsive noise. Contours of a BVI noise metric are employed to quantify the effects of rotor advance ratio and tip-path-plane angle on BVI noise directivity and amplitude. Acoustic time history data are presented to illustrate the variations in impulsive characteristics. The directionality, noise levels and impulsive content of both advancing and retreating side BVI are shown to vary significantly with tip-path-plane angle and advance ratio over the range of low and moderate flight speeds considered.

Near-blade flow structure modification

NASA Astrophysics Data System (ADS)

Kura, T.; Fornalik-Wajs, E.

2016-10-01

In this paper, the importance of near-blade flow structure influence on the performance of a centrifugal compressor was discussed. The negative effects of eddies and secondary flows appearance were described, together with the proposal of their reduction. Three-dimensional analyses were performed for the rotors. Focus was placed on the blade's 3D curvature impact on the efficiency of compression, and the influence of blade-shroud tip existence. A few design proposals were investigated - their performance maps were the basis of further analysis. Proposed modification of blade shape changed the near-blade flow structure and improved the compressor performance.

Slot configuration for axial-flow turbomachinery blades

NASA Technical Reports Server (NTRS)

Taylor, W. E.

1972-01-01

Machining of slot in turbine blades of axial flow turbines to provide flow path between pressure and suction surfaces is discussed. Slot configuration and improvements in blade performance are described. Diagram of blade slot to show geometry of modification is included.

Blade tip clearance measurement of the turbine engines based on a multi-mode fiber coupled laser ranging system

DOE Office of Scientific and Technical Information (OSTI.GOV)

Guo, Haotian; Duan, Fajie; Wu, Guoxiu

2014-11-15

The blade tip clearance is a parameter of great importance to guarantee the efficiency and safety of the turbine engines. In this article, a laser ranging system designed for blade tip clearance measurement is presented. Multi-mode fiber is utilized for optical transmission to guarantee that enough optical power is received by the sensor probe. The model of the tiny sensor probe is presented. The error brought by the optical path difference of different modes of the fiber is estimated and the length of the fiber is limited to reduce this error. The measurement range in which the optical power receivedmore » by the probe remains essentially unchanged is analyzed. Calibration experiments and dynamic experiments are conducted. The results of the calibration experiments indicate that the resolution of the system is about 0.02 mm and the range of the system is about 9 mm.« less

Distributed collaborative probabilistic design for turbine blade-tip radial running clearance using support vector machine of regression

NASA Astrophysics Data System (ADS)

Fei, Cheng-Wei; Bai, Guang-Chen

2014-12-01

To improve the computational precision and efficiency of probabilistic design for mechanical dynamic assembly like the blade-tip radial running clearance (BTRRC) of gas turbine, a distribution collaborative probabilistic design method-based support vector machine of regression (SR)(called as DCSRM) is proposed by integrating distribution collaborative response surface method and support vector machine regression model. The mathematical model of DCSRM is established and the probabilistic design idea of DCSRM is introduced. The dynamic assembly probabilistic design of aeroengine high-pressure turbine (HPT) BTRRC is accomplished to verify the proposed DCSRM. The analysis results reveal that the optimal static blade-tip clearance of HPT is gained for designing BTRRC, and improving the performance and reliability of aeroengine. The comparison of methods shows that the DCSRM has high computational accuracy and high computational efficiency in BTRRC probabilistic analysis. The present research offers an effective way for the reliability design of mechanical dynamic assembly and enriches mechanical reliability theory and method.

Radiated noise of ducted fans

NASA Astrophysics Data System (ADS)

Eversman, Walter

The differences in the radiated acoustic fields of ducted and unducted propellers of the same thrust operating under similar conditions are investigated. An FEM model is created for the generation, propagation, and radiation of steady, rotor alone noise and exit guide vane interaction noise of a ducted fan. For a specified number of blades, angular mode harmonic, and rotor angular velocity, the acoustic field is described in a cylindrical coordinate system reduced to only the axial and radial directions. It is found that, contrary to the usual understanding of the Tyler and Sofrin (1962) result, supersonic tip speed rotor noise can be cut off if the tip Mach number is only slightly in excess of unity and if the number of blades is relatively small. If there are many blades, the fundamental angular mode number is large, and the Tyler and Sofrin result for thin annuli becomes more relevant. Shrouding of subsonic tip speed propellers is a very effective means of controlling rotor alone noise.

Development and Evaluation of Processes for Deposition of Ni/Co-Cr-Aly (MCrAly) Coatings for Gas Turbine Components.

DTIC Science & Technology

1979-09-01

turbine engines demand increasingly higher operating tem- peratures in blades and vanes for greater thrust and efficiency. The turbine components...limitations; namely, expense and the inability to uniformly coat complex geometries and clustered turbine blade and vane airfoils . Thus, another means of...cost and the ability to uniformly coat turbine components of complex geometries and clustered turbine blade and vane airfoils .

Material Constitutive Models for Creep and Rupture of SiC/SiC Ceramic-Matrix Composites (CMCs) Under Multiaxial Loading

NASA Astrophysics Data System (ADS)

Grujicic, Mica; Galgalikar, R.; Snipes, J. S.; Ramaswami, S.

2016-05-01

Material constitutive models for creep deformation and creep rupture of the SiC/SiC ceramic-matrix composites (CMCs) under general three-dimensional stress states have been developed and parameterized using one set of available experimental data for the effect of stress magnitude and temperature on the time-dependent creep deformation and rupture. To validate the models developed, another set of available experimental data was utilized for each model. The models were subsequently implemented in a user-material subroutine and coupled with a commercial finite element package in order to enable computational analysis of the performance and durability of CMC components used in high-temperature high-stress applications, such as those encountered in gas-turbine engines. In the last portion of the work, the problem of creep-controlled contact of a gas-turbine engine blade with the shroud is investigated computationally. It is assumed that the blade is made of the SiC/SiC CMC, and that the creep behavior of this material can be accounted for using the material constitutive models developed in the present work. The results clearly show that the blade-tip/shroud clearance decreases and ultimately becomes zero (the condition which must be avoided) as a function of time. In addition, the analysis revealed that if the blade is trimmed at its tip to enable additional creep deformation before blade-tip/shroud contact, creep-rupture conditions can develop in the region of the blade adjacent to its attachment to the high-rotational-speed hub.

Aerodynamic interactions between a 1/6 scale helicopter rotor and a body of revolution

NASA Technical Reports Server (NTRS)

Betzina, M. D.; Shinoda, P.

1982-01-01

A wind-tunnel investigation was conducted in which independent, steady state aerodynamic forces and moments were measured on a 2.24-m-diam, two bladed helicopter rotor and a body of revolution. The objective was to determine the interaction of the body on the rotor performance and the effect of the rotor on the body aerodynamics for variations in velocity, thrust, tip-path-plane angle of attack, body angle of attack, rotor/body position, and body nose geometry. Results show that a body of revolution near the rotor can produce significant favorable or unfavorable effects on rotor performance, depending on the operating condition. Body longitudinal aerodynamic characteristics are significantly modified by the presence of an operating rotor and hub.

A Numerical Analysis of Heat Transfer and Effectiveness on Film Cooled Turbine Blade Tip Models

NASA Technical Reports Server (NTRS)

Ameri, A. A.; Rigby, D. L.

1999-01-01

A computational study has been performed to predict the distribution of convective heat transfer coefficient on a simulated blade tip with cooling holes. The purpose of the examination was to assess the ability of a three-dimensional Reynolds-averaged Navier-Stokes solver to predict the rate of tip heat transfer and the distribution of cooling effectiveness. To this end, the simulation of tip clearance flow with blowing of Kim and Metzger was used. The agreement of the computed effectiveness with the data was quite good. The agreement with the heat transfer coefficient was not as good but improved away from the cooling holes. Numerical flow visualization showed that the uniformity of wetting of the surface by the film cooling jet is helped by the reverse flow due to edge separation of the main flow.

Improved Temperature Dynamic Model of Turbine Subcomponents for Facilitation of Generalized Tip Clearance Control

NASA Technical Reports Server (NTRS)

Kypuros, Javier A.; Colson, Rodrigo; Munoz, Afredo

2004-01-01

This paper describes efforts conducted to improve dynamic temperature estimations of a turbine tip clearance system to facilitate design of a generalized tip clearance controller. This work builds upon research previously conducted and presented in and focuses primarily on improving dynamic temperature estimations of the primary components affecting tip clearance (i.e. the rotor, blades, and casing/shroud). The temperature profiles estimated by the previous model iteration, specifically for the rotor and blades, were found to be inaccurate and, more importantly, insufficient to facilitate controller design. Some assumptions made to facilitate the previous results were not valid, and thus improvements are presented here to better match the physical reality. As will be shown, the improved temperature sub- models, match a commercially validated model and are sufficiently simplified to aid in controller design.

Mind the gap - tip leakage vortex in axial turbines

NASA Astrophysics Data System (ADS)

Dreyer, M.; Decaix, J.; Münch-Alligné, C.; Farhat, M.

2014-03-01

The tendency of designing large Kaplan turbines with a continuous increase of output power is bringing to the front the cavitation erosion issue. Due to the flow in the gap between the runner and the discharge ring, axial turbine blades may develop the so called tip leakage vortex (TLV) cavitation with negative consequences. Such vortices may interact strongly with the wake of guide vanes leading to their multiple collapses and rebounds. If the vortex trajectory remains close to the blade tip, these collapses may lead to severe erosion. One is still unable today to predict its occurrence and development in axial turbines with acceptable accuracy. Numerical flow simulations as well as the actual scale-up rules from small to large scales are unreliable. The present work addresses this problematic in a simplified case study representing TLV cavitation to better understand its sensitivity to the gap width. A Naca0009 hydrofoil is used as a generic blade in the test section of EPFL cavitation tunnel. A sliding mounting support allowing an adjustable gap between the blade tip and wall was manufactured. The vortex trajectory is visualized with a high speed camera and appropriate lighting. The three dimensional velocity field induced by the TLV is investigated using stereo particle image velocimetry. We have taken into account the vortex wandering in the image processing to obtain accurate measurements of the vortex properties. The measurements were performed in three planes located downstream of the hydrofoil for different values of the flow velocity, the incidence angle and the gap width. The results clearly reveal a strong influence of the gap width on both trajectory and intensity of the tip leakage vortex.

Performance of a 1380-foot-per-second-tip-speed axial-flow compressor rotor with a blade tip solidity of 1.3

NASA Technical Reports Server (NTRS)

Hager, R. D.; Janetzke, D. C.; Reid, L.

1972-01-01

Aerodynamic design parameters are presented along the overall and blade element performance, of an axial flow compressor rotor designed to study the effects of blade solidity on efficiency and stall margin. At design speed the peak efficiency was 0.844 and occurred at an equivalent weight flow of 63.5 lb/sec with a total pressure ratio of 1.801. Design efficiency, pressure ratio, and weight flow 0.814, 1.65, and 65.3(41.1 lb/sec/sq ft of annulus area), respectively. Stall margin for design speed was 6.4 percent based on the weight flow and pressure ratio values at peak efficiency and just prior to stall.

Two stage low noise advanced technology fan. 1: Aerodynamic, structural, and acoustic design

NASA Technical Reports Server (NTRS)

Messenger, H. E.; Ruschak, J. T.; Sofrin, T. G.

1974-01-01

A two-stage fan was designed to reduce noise 20 db below current requirements. The first-stage rotor has a design tip speed of 365.8 m/sec and a hub/tip ratio of 0.4. The fan was designed to deliver a pressure ratio of 1.9 with an adiabatic efficiency of 85.3 percent at a specific inlet corrected flow of 209.2kg/sec/sq m. Noise reduction devices include acoustically treated casing walls, a flowpath exit acoustic splitter, a translating centerbody sonic inlet device, widely spaced blade rows, and the proper ratio of blades and vanes. Multiple-circular-arc rotor airfoils, resettable stators, split outer casings, and capability to go to close blade-row spacing are also included.

Design of Single Stage Axial Turbine with Constant Nozzle Angle Blading for Small Turbojet

NASA Astrophysics Data System (ADS)

Putra Adnan, F.; Hartono, Firman

2018-04-01

In this paper, an aerodynamic design of a single stage gas generator axial turbine for small turbojet engine is explained. As per design requirement, the turbine should be able to deliver power output of 155 kW at 0.8139 kg/s gas mass flow, inlet total temperature of 1200 K and inlet total pressure of 335330 Pa. The design phase consist of several steps, i.e.: determination of velocity triangles in 2D plane, 2D blading design and 3D flow analysis at design point using Computational Fluid Dynamics method. In the determination of velocity triangles, two conditions are applied: zero inlet swirl (i.e. the gas flow enter the turbine at axial direction) and constant nozzle angle design (i.e. the inlet and outlet angle of the nozzle blade are constant from root to tip). The 2D approach in cascade plane is used to specify airfoil type at root, mean and tip of the blade based on inlet and outlet flow conditions. The 3D approach is done by simulating the turbine in full configuration to evaluate the overall performance of the turbine. The observed parameters including axial gap, stagger angle, and tip clearance affect its output power. Based on analysis results, axial gap and stagger angle are positively correlated with output power up to a certain point at which the power decreases. Tip clearance, however, gives inversely correlation with output power.

JT8D-15/17 High Pressure Turbine Root Discharged Blade Performance Improvement. [engine design

NASA Technical Reports Server (NTRS)

Janus, A. S.

1981-01-01

The JT8D high pressure turbine blade and seal were modified, using a more efficient blade cooling system, improved airfoil aerodynamics, more effective control of secondary flows, and improved blade tip sealing. Engine testing was conducted to determine the effect of these improvements on performance. The modified turbine package demonstrated significant thrust specific fuel consumption and exhaust gas temperature improvements in sea level and altitude engine tests. Inspection of the improved blade and seal hardware after testing revealed no unusual wear or degradation.

Applying Dynamic Wake Models to Induced Power Calculations for an Optimum Rotor

DTIC Science & Technology

2009-08-01

versions being special cases of the general one. Although the rotor blade may be moving at transonic speeds near the tip, the rotor wake is...The effect of a finite number of blades incurs an additional loss in wake energy due to the individual vortex sheets from each blade . In 1929... blades . Up to this point, previous developments have been able to achieve the full description of the wake in all ranges of flight regime

Development of a Novel Erosion Resistant Coating System for Use on Rotorcraft Blades

DTIC Science & Technology

2012-05-01

Technologies Research Center (UTRC) and Sikorsky utilizes a two part metal/ cermet coating system on the leading edge of the blades to provide unmatched...ARL, United Technologies Research Center (UTRC) and Sikorsky utilizes a two part metal/ cermet coating system on the leading edge of the blades to...Rotor Blade Tip Fairing A study by Ely et.al. evaluated dozens of coating technologies and down-selected a two-part metal/ceramic coating system on

Investigation of Unsteady Flow Behavior in Transonic Compressor Rotors with LES and PIV Measurements

NASA Technical Reports Server (NTRS)

Hah, Chunill; Voges, Melanie; Mueller, Martin; Schiffer, Heinz-Peter

2009-01-01

In the present study, unsteady flow behavior in a modern transonic axial compressor rotor is studied in detail with large eddy simulation (LES) and particle image velocimetry (PIV). The main purpose of the study is to advance the current understanding of the flow field near the blade tip in an axial transonic compressor rotor near the stall and peak-efficiency conditions. Flow interaction between the tip leakage vortex and the passage shock is inherently unsteady in a transonic compressor. Casing-mounted unsteady pressure transducers have been widely applied to investigate steady and unsteady flow behavior near the casing. Although many aspects of flow have been revealed, flow structures below the casing cannot be studied with casing-mounted pressure transducers. In the present study, unsteady velocity fields are measured with a PIV system and the measured unsteady flow fields are compared with LES simulations. The currently applied PIV measurements indicate that the flow near the tip region is not steady even at the design condition. This self-induced unsteadiness increases significantly as the compressor rotor operates near the stall condition. Measured data from PIV show that the tip clearance vortex oscillates substantially near stall. The calculated unsteady characteristics of the flow from LES agree well with the PIV measurements. Calculated unsteady flow fields show that the formation of the tip clearance vortex is intermittent and the concept of vortex breakdown from steady flow analysis does not seem to apply in the current flow field. Fluid with low momentum near the pressure side of the blade close to the leading edge periodically spills over into the adjacent blade passage. The present study indicates that stall inception is heavily dependent on unsteady behavior of the flow field near the leading edge of the blade tip section for the present transonic compressor rotor.

Cavitation noise studies on marine propellers

NASA Astrophysics Data System (ADS)

Sharma, S. D.; Mani, K.; Arakeri, V. H.

1990-04-01

Experimental observations are described of cavitation inception and noise from five model propellers, three basic and two modified, tested in the open jet section of the Indian Institute of Science high-speed water tunnel facility. Extensive experiments on the three basic propellers of different design, which included visualization of cavitation and measurements of noise, showed that the dominant type of cavitation was in the form of tip vortex cavitation, accompanied by leading edge suction side sheet cavitation in its close vicinity, and the resultant noise depended on parameters such as the advance coefficient, the cavitation number, and the propeller geometry. Of these, advance coefficient was found to have the maximum influence not only on cavitation noise but also on the inception of cavitation. Noise levels and frequencies of spectra obtained from all the three basic propellers at conditions near inception and different advance coefficient values, when plotted in the normalized form as suggested by Blake, resulted in a universal spectrum which would be useful for predicting cavitation noise at prototype scales when a limited extent of cavitation is expected in the same form as observed on the present models. In an attempt to delay the onset of tip vortex cavitation, the blades of two of the three basic propellers were modified by drilling small holes in the tip and leading edge areas. Studies on the modified propellers showed that the effectiveness of the blade modification was apparently stronger at low advance coefficient values and depended on the blade sectional profile. Measurements of cavitation noise indicated that the modification also improved the acoustic performance of the propellers as it resulted in a complete attenuation of the low-frequency spectral peaks, which were prominent with the basic propellers. In addition to the above studies, which were conducted under uniform flow conditions, one of the basic propellers was tested in the simulated wake of a typical single screw ship. The wake was simulated by using a wire screen technique. Observations of cavitation and measurement of noise clearly showed that the presence of the wake had a strong influence on the propeller cavitation and noise performance. Cavitation was found to be of the cloud type, which generated very intense noise compared to that generated by tip vortex cavitation along with leading edge suction side sheet cavitation in the uniform flow conditions. The noise spectra obtained with wake simulation also are presented in a normalized form to be of general utility.

Variable geometry Darrieus wind machine

NASA Astrophysics Data System (ADS)

Pytlinski, J. T.; Serrano, D.

1983-08-01

A variable geometry Darrieus wind machine is proposed. The lower attachment of the blades to the rotor can move freely up and down the axle allowing the blades of change shape during rotation. Experimental data for a 17 m. diameter Darrieus rotor and a theoretical model for multiple streamtube performance prediction were used to develop a computer simulation program for studying parameters that affect the machine's performance. This new variable geometry concept is described and interrelated with multiple streamtube theory through aerodynamic parameters. The computer simulation study shows that governor behavior of a Darrieus turbine can not be attained by a standard turbine operating within normally occurring rotational velocity limits. A second generation variable geometry Darrieus wind turbine which uses a telescopic blade is proposed as a potential improvement on the studied concept.

Design of 9.271-pressure-ratio 5-stage core compressor and overall performance for first 3 stages

NASA Technical Reports Server (NTRS)

Steinke, Ronald J.

1986-01-01

Overall aerodynamic design information is given for all five stages of an axial flow core compressor (74A) having a 9.271 pressure ratio and 29.710 kg/sec flow. For the inlet stage group (first three stages), detailed blade element design information and experimental overall performance are given. At rotor 1 inlet tip speed was 430.291 m/sec, and hub to tip radius ratio was 0.488. A low number of blades per row was achieved by the use of low-aspect-ratio blading of moderate solidity. The high reaction stages have about equal energy addition. Radial energy varied to give constant total pressure at the rotor exit. The blade element profile and shock losses and the incidence and deviation angles were based on relevant experimental data. Blade shapes are mostly double circular arc. Analysis by a three-dimensional Euler code verified the experimentally measured high flow at design speed and IGV-stator setting angles. An optimization code gave an optimal IGV-stator reset schedule for higher measured efficiency at all speeds.

Rotor Design Options for Improving XV-15 Whirl-Flutter Stability Margins

NASA Technical Reports Server (NTRS)

Acree, C. W., Jr.; Peyran, R. J.; Johnson, Wayne

2004-01-01

Rotor design changes intended to improve tiltrotor whirl-flutter stability margins were analyzed. A baseline analytical model of the XV-15 was established, and then a thinner, composite wing was designed to be representative of a high-speed tiltrotor. The rotor blade design was modified to increase the stability speed margin for the thin-wing design. Small rearward offsets of the aerodynamic-center locus with respect to the blade elastic axis created large increases in the stability boundary. The effect was strongest for offsets at the outboard part of the blade, where an offset of the aerodynamic center by 10% of tip chord improved the stability margin by over 100 knots. Forward offsets of the blade center of gravity had similar but less pronounced effects. Equivalent results were seen for swept-tip blades. Appropriate combinations of sweep and pitch stiffness completely eliminated whirl flutter within the speed range examined; alternatively, they allowed large increases in pitch-flap coupling (delta-three) for a given stability margin. A limited investigation of the rotor loads in helicopter and airplane configuration showed only minor increases in loads.

Influence of the conservative rotor loads on the near wake of a wind turbine

NASA Astrophysics Data System (ADS)

Herráez, I.; Micallef, D.; van Kuik, G. A. M.

2017-05-01

The presence of conservative forces on rotor blades is neglected in the blade element theory and all the numerical methods derived from it (like e.g. the blade element momentum theory and the actuator line technique). This might seem a reasonable simplification of the real flow of rotor blades, since conservative loads, by definition, do not contribute to the power conversion. However, conservative loads originating from the chordwise bound vorticity might affect the tip vortex trajectory, as we discussed in a previous work. In that work we also hypothesized that this effect, in turn, could influence the wake induction and correspondingly the rotor performance. In the current work we extend a standard actuator line model in order to account for the conservative loads at the blade tip. This allows to isolate the influence of conservative forces from other effects. The comparison of numerical results with and without conservative loads enables to confirm qualitatively their relevance for the near wake and the rotor performance. However, an accurate quantitative assessment of the effect still remains out of reach due to the inherent uncertainty of the numerical model.

A Comparison of Measured Tone Modes for Two Low Noise Propulsion Fans

NASA Technical Reports Server (NTRS)

Heidelberg, Laurence J.; Elliott, David M.

2000-01-01

The acoustic modes for two low tip speed propulsion fans were measured to examine the effects of fan tip speed, at constant pressure ratio. A continuously rotating microphone method was used that provided the complete modal structure (circumferential and radial order) at the fundamental and second harmonic of the blade passing tone as well as most of the third harmonic modes. The fans are compared in terms of their rotor/stator interaction modal power, and total tone power. It was hoped that the lower tip speed might produce less noise. This was not the case. The higher tip speed fan, at both takeoff and cutback speeds, had lower tone and interaction levels. This could be an indication that the higher aerodynamic loading required to produce the same pressure ratio for the lower tip speed fan resulted in a greater velocity deficit in the blade wakes and thus more noise. Results consistent with expected rotor transmission effects were noted in the inlet modal structures of both fans.

Data Reduction Procedures for Laser Velocimeter Measurements in Turbomachinery Rotors

NASA Technical Reports Server (NTRS)

Lepicovsky, Jan

1994-01-01

Blade-to-blade velocity distributions based on laser velocimeter data acquired in compressor or fan rotors are increasingly used as benchmark data for the verification and calibration of turbomachinery computational fluid dynamics (CFD) codes. Using laser Doppler velocimeter (LDV) data for this purpose, however, must be done cautiously. Aside from the still not fully resolved issue of the seed particle response in complex flowfields, there is an important inherent difference between CFD predictions and LDV blade-to-blade velocity distributions. CFD codes calculate velocity fields for an idealized rotor passage. LDV data, on the other hand, stem from the actual geometry of all blade channels in a rotor. The geometry often varies from channel to channel as a result of manufacturing tolerances, assembly tolerances, and incurred operational damage or changes in the rotor individual blades.

Effect of casing treatment on overall and blade element performance of a compressor rotor

NASA Technical Reports Server (NTRS)

Moore, R. D.; Kovich, G.; Blade, R. J.

1971-01-01

An axial flow compressor rotor was tested at design speed with six different casing treatments across the rotor tip. Radial surveys of pressure, temperature, and flow angle were taken at the rotor inlet and outlet. Surveys were taken at several weight flows for each treatment. All the casings treatments decreased the weight flow at stall over that for the solid casing. Radial surveys indicate that the performance over the entire radial span of the blade is affected by the treatment across the rotor tip.

Wind-tunnel measurement of noise emitted by helicopter rotors at high speed

NASA Astrophysics Data System (ADS)

Prieur, J.

Measurements of high-speed impulsive helicopter rotor noise in a wind-tunnel are presented. High-speed impulsive noise measurements have been performed in 1988 in the ONERA S2ch wind-tunnel, fitted with an acoustic lining, on two types of rotors. They show that substantial noise reduction is obtained with sweptback tips, initially designed for aerodynamic purposes, which lower transonic effects on the advancing blade tip. Emphasis is placed on the necessity of taking into account the acoustic annoyance problem, using noise prediction tools, when designing new helicopter blades.

Studies of blade-vortex interaction noise reduction by rotor blade modification

NASA Technical Reports Server (NTRS)

Brooks, Thomas F.

1993-01-01

Blade-vortex interaction (BVI) noise is one of the most objectionable types of helicopter noise. This impulsive blade-slap noise can be particularly intense during low-speed landing approach and maneuvers. Over the years, a number of flight and model rotor tests have examined blade tip modification and other blade design changes to reduce this noise. Many times these tests have produced conflicting results. In the present paper, a number of these studies are reviewed in light of the current understanding of the BVI noise problem. Results from one study in particular are used to help establish the noise reduction potential and to shed light on the role of blade design. Current blade studies and some new concepts under development are also described.

A Numerical Analysis on the Effects of Self-Excited Tip Flow Unsteadiness and Upstream Blade Row Interactions on the Performance Predictions of a Transonic Compressor

NASA Astrophysics Data System (ADS)

Heberling, Brian

Computational fluid dynamics (CFD) simulations can offer a detailed view of the complex flow fields within an axial compressor and greatly aid the design process. However, the desire for quick turnaround times raises the question of how exact the model must be. At design conditions, steady CFD simulating an isolated blade row can accurately predict the performance of a rotor. However, as a compressor is throttled and mass flow rate decreased, axial flow becomes weaker making the capturing of unsteadiness, wakes, or other flow features more important. The unsteadiness of the tip clearance flow and upstream blade wake can have a significant impact on a rotor. At off-design conditions, time-accurate simulations or modeling multiple blade rows can become necessary in order to receive accurate performance predictions. Unsteady and multi- bladerow simulations are computationally expensive, especially when used in conjunction. It is important to understand which features are important to model in order to accurately capture a compressor's performance. CFD simulations of a transonic axial compressor throttling from the design point to stall are presented. The importance of capturing the unsteadiness of the rotor tip clearance flow versus capturing upstream blade-row interactions is examined through steady and unsteady, single- and multi-bladerow computations. It is shown that there are significant differences at near stall conditions between the different types of simulations.

Turbine blade cooling

DOEpatents

Staub, F.W.; Willett, F.T.

1999-07-20

A turbine rotor blade comprises a shank portion, a tip portion and an airfoil. The airfoil has a pressure side wall and a suction side wall that are interconnected by a plurality of partition sidewalls, defining an internal cooling passageway within the airfoil. The internal cooling passageway includes at least one radial outflow passageway to direct a cooling medium flow from the shank portion towards the tip portion and at least one radial inflow passageway to direct a cooling medium flow from the tip portion towards the shank portion. A number of mixing ribs are disposed on the partition sidewalls within the radial outflow passageways so as to enhance the thermal mixing of the cooling medium flow, thereby producing improved heat transfer over a broad range of the Buoyancy number. 13 figs.

Effects of Swept Tips on V-22 Whirl Flutter and Loads

NASA Technical Reports Server (NTRS)

Acree, C. W., Jr.

2005-01-01

A CAMRAD II model of the V-22 Osprey tiltrotor was constructed for the purpose of analyzing the effects of blade design changes on whirl flutter. The model incorporated a dual load-path grip/yoke assembly, a swashplate coupled to the transmission case, and a drive train. A multiple-trailer free wake was used for loads calculations. The effects of rotor design changes on whirl-mode stability were calculated for swept blades and offset tip masses. A rotor with swept tips and inboard tuning masses was examined in detail to reveal the mechanisms by which these design changes affect stability and loads. Certain combinations of design features greatly increased whirl-mode stability, with (at worst) moderate increases to loads.

Turbine blade cooling

DOEpatents

Staub, Fred Wolf; Willett, Fred Thomas

1999-07-20

A turbine rotor blade comprises a shank portion, a tip portion and an airfoil. The airfoil has a pressure side wall and a suction side wall that are interconnected by a plurality of partition sidewalls, defining an internal cooling passageway within the airfoil. The internal cooling passageway includes at least one radial outflow passageway to direct a cooling medium flow from the shank portion towards the tip portion and at least one radial inflow passageway to direct a cooling medium flow from the tip portion towards the shank portion. A number of mixing ribs are disposed on the partition sidewalls within the radial outflow passageways so as to enhance the thermal mixing of the cooling medium flow, thereby producing improved heat transfer over a broad range of the Buoyancy number.

Turbine blade cooling

DOEpatents

Staub, Fred Wolf; Willett, Fred Thomas

2000-01-01

A turbine rotor blade comprises a shank portion, a tip portion and an airfoil. The airfoil has a pressure side wall and a suction side wall that are interconnected by a plurality of partition sidewalls, defining an internal cooling passageway within the airfoil. The internal cooling passageway includes at least one radial outflow passageway to direct a cooling medium flow from the shank portion towards the tip portion and at least one radial inflow passageway to direct a cooling medium flow from the tip portion towards the shank portion. A number of mixing ribs are disposed on the partition sidewalls within the radial outflow passageways so as to enhance the thermal mixing of the cooling medium flow, thereby producing improved heat transfer over a broad range of the Buoyancy number.

Experimental and theoretical investigation of HT-S/PMR-PI composites for application to advanced aircraft engines. [High-Tip-Speed/Polymerization of Monomeric Reactant

NASA Technical Reports Server (NTRS)

Hanson, M. P.; Chamis, C. C.

1974-01-01

A combined experimental and theoretical investigation was performed in order to: (1) demonstrate that high quality angleplied laminates can be made from HT-S/PMR-PI (PMR in situ polymerization of monomeric reactants), (2) characterize the PMR-PI material and to determine the HT-S unidirectional composite properties required for composite micro and macromechanics and laminate analyses, (3) select HT-S/PMR laminate configurations to meet the general design requirements for high-tip-speed compressor blades. The results of the investigation showed that: HT-S/PMR laminate configurations can be fabricated which satisfy the high-tip-speed compressor blade design requirements when operating within the temperature capability of the polymide matrix.

Feasibility Study for a Practical High Rotor Tip Clearance Turbine.

DTIC Science & Technology

GAS TURBINE BLADES ), (* TURBINE BLADES , TOLERANCES(MECHANICS)), (* TURBOFAN ENGINES , GAS TURBINES , AXIAL FLOW TURBINES , AXIAL FLOW TURBINE ROTORS...AERODYNAMIC CONFIGURATIONS, LEAKAGE(FLUID), MEASUREMENT, TEST METHODS, PERFORMANCE( ENGINEERING ), MATHEMATICAL PREDICTION, REDUCTION, PRESSURE, PREDICTIONS, NOZZLE GAS FLOW, COMBUSTION CHAMBER GASES, GAS FLOW.

A method to estimate wind turbine blade damage and to design damage-resilient blades

NASA Astrophysics Data System (ADS)

Fiore, Giovanni

Wind turbine blades are affected by continuous impacts with airborne particles that deteriorate the blade surface and yield to a drop in output power. Based on the climatic conditions and geographic locations of a given wind farm, multiple types of particles are observed in air. The present study focuses on simulating the impact of four types of particles, namely insects, sand grains, hailstones, and rain drops with the blade surface. A numerical inviscid flowfield code, coupled with a particle position predictor code was used. Upon impact, the damaging effect to the blade surface was evaluated. Each type of particle was associated with a damage mode, which depends on the mass, size, and hardness of the particle. It was found that insects strike and adhere to the blade in a region close to the leading edge. On the other hand, it was seen that sand grains promote erosion just downstream of the leading edge, where local velocity reaches a maximum and the impact angle is shallow. Moreover, particles such as rain drops are associated with fatigue and erosion at the very leading edge and on the upper side of the blade section. Finally, hailstones promote delamination and fatigue in the composite panels of the blade surface. Photographic evidence of damaged blade surfaces was used in the present research as a comparison with the simulations performed for various types of particle and different initial conditions. Based on such observations, a theorization of the damage pattern and evolution was proposed. Finally, given a set of well-established blade section geometries, such as the Delft University and NREL S airfoil families, a comparison of airfoil damage fitness was proposed and possible means of shape optimization were discussed. The investigation of blade geometry features to mitigate damage was performed. Based on previous results, it was argued that a viable blade section optimization may be performed for the lightest and smallest particles considered in the study, the sand grains. A pool of airfoils was analyzed regarding the sand erosion rate. It was shown that a bulbous leading edge coupled with airfoil aft camber is beneficial toward the erosion rate due to sand grains. An optimization algorithm was written to improve the damage resilience toward sand erosion of wind turbine airfoils. A direct and inverse approach were integrated in a genetic algorithm code, and it was confirmed that bulbous leading edges, coupled with aft cambers allowed for a reduction in blade erosion rates. Lastly, a time-stepping code was developed to predict the blade section geometry when sand erosion is present. It was found that three main phases occur during the erosive life of a blade. A parametric study allowed to find the most relevant drivers to the blade lifespan with respect to erosion. Beneficial effects come from an increase in turbine hub height, turbine rated power, increase in lift coefficient, and a reduction in average particle diameter. A parametric study was also performed by investigating different airfoil geometries. Again, it was found that bulbous leading edges coupled with aft cambered geometries allow for longer blade lifespan.

Automated eddy current inspection of Space Shuttle APU turbine wheel blades

NASA Technical Reports Server (NTRS)

Fisher, Jay L.; Rowland, Stephen N.; Stolte, Jeffrey S.; Salkowski, Charles

1991-01-01

An automated inspection system based on eddy current testing (ET) techniques has been developed to inspect turbine wheel blades on the APU used in NASA's Space Transportation system. The APU is a hydrazine-powered gas turbine with a 15-cm diameter Rene 41 turbine wheel, which has 123 first-stage blades and 123 second-stage blades. The flaw detection capability of the ET system is verified through comparison with fluorescent penetrant test results. Results of the comparison indicate that ET is capable of inspecting surfaces with very restrictive geometries. The ET capability requires development of probes with extremely small coils to allow inspection within 0.4 mm of the blade root and the leading and trailing edges of the blade and within a height restriction of less than 1 mm. The color 2D presentation of the ET data provided crack-growth pattern and length information similar to those found with visual techniques. It also provided visual clues to minimize geometry effects such as generated from blade edges, a neighoring blade, and changes in the blade thickness.

Design and analysis of flow velocity distribution inside a raceway pond using computational fluid dynamics.

PubMed

Pandey, Ramakant; Premalatha, M

2017-03-01

Open raceway ponds are widely adopted for cultivating microalgae on a large scale. Working depth of the raceway pond is the major component to be analysed for increasing the volume to surface area ratio. The working depth is limited up to 5-15 cm in conventional ponds but in this analysis working depth of raceway pond is considered as 25 cm. In this work, positioning of the paddle wheel is analysed and corresponding Vertical Mixing Index are calculated using CFD. Flow pattern along the length of the raceway pond, at three different paddle wheel speeds are analysed for L/W ratio of 6, 8 and 10, respectively. Effect of clearance (C) between rotor blade tip and bottom surface is also analysed by taking four clearance conditions i.e. C = 2, 5, 10 and 15. Moving reference frame method of Fluent is used for the modeling of six blade paddle wheel and realizable k-ε model is used for capturing turbulence characteristics. Overall objective of this work is to analyse the required geometry for maintaining a minimum flow velocity to avoid settling of algae corresponding to 25 cm working depth. Geometry given in [13] is designed using ANSYS Design modular and CFD results are generated using ANSYS FLUENT for the purpose of validation. Good agreement of results is observed between CFD and experimental Particle image velocimetry results with the deviation of 7.23%.

Fundamental Understanding of Rotor Aeromechanics at High Advance Ratio Through Wind Tunnel Testing

NASA Astrophysics Data System (ADS)

Berry, Benjamin

The purpose of this research is to further the understanding of rotor aeromechanics at advance ratios (mu) beyond the maximum of 0.5 (ratio of forward airspeed to rotor tip speed) for conventional helicopters. High advance ratio rotors have applications in high speed compound helicopters. In addition to one or more conventional main rotors, these aircraft employ either thrust compounding (propellers), lift compounding (fixed-wings), or both. An articulated 4-bladed model rotor was constructed, instrumented, and tested up to a maximum advance ratio of mu=1.6 in the Glenn L. Martin Wind Tunnel at the University of Maryland. The data set includes steady and unsteady rotor hub forces and moments, blade structural loads, blade flapping angles, swashplate control angles, and unsteady blade pressures. A collective-thrust control reversal--where increasing collective pitch results in lower rotor thrust--was observed and is a unique phenomenon to the high advance ratio flight regime. The thrust reversal is explained in a physical manner as well as through an analytical formulation. The requirements for the occurrence of the thrust reversal are enumerated. The effects of rotor geometry design on the thrust reversal onset are explored through the formulation and compared to the measured data. Reverse-flow dynamic stall was observed to extend the the lifting capability of the edgewise rotor well beyond the expected static stall behavior of the airfoil sections. Through embedded unsteady blade surface pressure transducers, the normal force, pitching moment, and shed dynamic stall vortex time histories at a blade section in strong reverse flow were analyzed. Favorable comparisons with published 2-D pitching airfoil reverse flow dynamic stall data indicate that the 3-D stall environment can likely be predicted using models developed from such 2-D experiments. Vibratory hub loads were observed to increase with advance ratio. Maximum amplitude was observed near mu=1, with a reduction in vibratory loads at higher advance ratios. Blade load 4/rev harmonics dominated due to operation near a 4/rev fanplot crossing of the 2nd flap bending mode natural frequency. Oscillatory loads sharply increase in the presence of retreating blade reverse flow dynamic stall, and are evident in blade torsion, pitch link, and hub load measurements. The blades exhibited torsion moment vibrations at the frequency of the 1st torsion mode in response to the reverse flow pitching moment loading.

Airfoil family design for large offshore wind turbine blades

NASA Astrophysics Data System (ADS)

Méndez, B.; Munduate, X.; San Miguel, U.

2014-06-01

Wind turbine blades size has scaled-up during last years due to wind turbine platform increase especially for offshore applications. The EOLIA project 2007-2010 (Spanish Goverment funded project) was focused on the design of large offshore wind turbines for deep waters. The project was managed by ACCIONA Energia and the wind turbine technology was designed by ACCIONA Windpower. The project included the design of a wind turbine airfoil family especially conceived for large offshore wind turbine blades, in the order of 5MW machine. Large offshore wind turbines suffer high extreme loads due to their size, in addition the lack of noise restrictions allow higher tip speeds. Consequently, the airfoils presented in this work are designed for high Reynolds numbers with the main goal of reducing blade loads and mantainig power production. The new airfoil family was designed in collaboration with CENER (Spanish National Renewable Energy Centre). The airfoil family was designed using a evolutionary algorithm based optimization tool with different objectives, both aerodynamic and structural, coupled with an airfoil geometry generation tool. Force coefficients of the designed airfoil were obtained using the panel code XFOIL in which the boundary layer/inviscid flow coupling is ineracted via surface transpiration model. The desing methodology includes a novel technique to define the objective functions based on normalizing the functions using weight parameters created from data of airfoils used as reference. Four airfoils have been designed, here three of them will be presented, with relative thickness of 18%, 21%, 25%, which have been verified with the in-house CFD code, Wind Multi Block WMB, and later validated with wind tunnel experiments. Some of the objectives for the designed airfoils concern the aerodynamic behavior (high efficiency and lift, high tangential coefficient, insensitivity to rough conditions, etc.), others concern the geometry (good for structural design, compatibility for the different airfoil family members, etc.) and with the ultimate objective that the airfoils will reduce the blade loads. In this paper the whole airfoil design process and the main characteristics of the airfoil family are described. Some force coefficients for the design Reynolds number are also presented. The new designed airfoils have been studied with computational calculations (panel method code and CFD) and also in a wind tunnel experimental campaign. Some of these results will be also presented in this paper.

Probabilistic structural analysis to quantify uncertainties associated with turbopump blades

NASA Technical Reports Server (NTRS)

Nagpal, Vinod K.; Rubinstein, Robert; Chamis, Christos C.

1988-01-01

A probabilistic study of turbopump blades has been in progress at NASA Lewis Research Center for over the last two years. The objectives of this study are to evaluate the effects of uncertainties in geometry and material properties on the structural response of the turbopump blades to evaluate the tolerance limits on the design. A methodology based on probabilistic approach was developed to quantify the effects of the random uncertainties. The results indicate that only the variations in geometry have significant effects.

Integrated Aeromechanics with Three-Dimensional Solid-Multibody Structures

NASA Technical Reports Server (NTRS)

Datta, Anubhav; Johnson, Wayne

2014-01-01

A full three-dimensional finite element-multibody structural dynamic solver is coupled to a three-dimensional Reynolds-averaged Navier-Stokes solver for the prediction of integrated aeromechanical stresses and strains on a rotor blade in forward flight. The objective is to lay the foundations of all major pieces of an integrated three-dimensional rotor dynamic analysis - from model construction to aeromechanical solution to stress/strain calculation. The primary focus is on the aeromechanical solution. Two types of three-dimensional CFD/CSD interfaces are constructed for this purpose with an emphasis on resolving errors from geometry mis-match so that initial-stage approximate structural geometries can also be effectively analyzed. A three-dimensional structural model is constructed as an approximation to a UH-60A-like fully articulated rotor. The aerodynamic model is identical to the UH-60A rotor. For preliminary validation measurements from a UH-60A high speed flight is used where CFD coupling is essential to capture the advancing side tip transonic effects. The key conclusion is that an integrated aeromechanical analysis is indeed possible with three-dimensional structural dynamics but requires a careful description of its geometry and discretization of its parts.

Tests of Full-Scale Helicopter Rotors at High Advancing Tip Mach Numbers and Advance Ratios

NASA Technical Reports Server (NTRS)

Biggers, James C.; McCloud, John L., III; Stroub, Robert H.

2015-01-01

As a continuation of the studies of reference 1, three full-scale helicopter rotors have been tested in the Ames Research Center 40- by SO-foot wind tunnel. All three of them were two-bladed, teetering rotors. One of the rotors incorporated the NACA 0012 airfoil section over the entire length of the blade. This rotor was tested at advance ratios up to 1.05. Both of the other rotors were tapered in thickness and incorporated leading-edge camber over the outer 20 percent of the blade radius. The larger of these rotors was tested at advancing tip Mach numbers up to 1.02. Data were obtained for a wide range of lift and propulsive force, and are presented without discussion.

Nonlinear piezoelectric devices for broadband air-flow energy harvesting

NASA Astrophysics Data System (ADS)

Bai, Y.; Havránek, Z.; Tofel, P.; Meggs, C.; Hughes, H.; Button, T. W.

2015-11-01

This paper presents preliminary work on an investigation of a nonlinear air-flow energy harvester integrating magnets and a piezoelectric cantilever array. Two individual piezoelectric cantilevers with the structure of free-standing multi-layer thick-films have been fabricated and assembled with a free-spinning fan. The cantilevers were attached with different tip masses thereby achieving separated resonant frequencies. Also, permanent magnets were fixed onto the blades of the fan as well as the tips of the cantilevers, in order to create nonlinear coupling and transfer fluidic movement into mechanical oscillation. The device has been tested in a wind tunnel. Bifurcations in the spectra of the blade rotation speed of the fan as a function of output voltage have been observed, and a bandwidth (blade rotation speed range) widening effect has been achieved.

An improved computer model for prediction of axial gas turbine performance losses

NASA Technical Reports Server (NTRS)

Jenkins, R. M.

1984-01-01

The calculation model performs a rapid preliminary pitchline optimization of axial gas turbine annular flowpath geometry, as well as an initial estimate of blade profile shapes, given only a minimum of thermodynamic cycle requirements. No geometric parameters need be specified. The following preliminary design data are determined: (1) the optimum flowpath geometry, within mechanical stress limits; (2) initial estimates of cascade blade shapes; and (3) predictions of expected turbine performance. The model uses an inverse calculation technique whereby blade profiles are generated by designing channels to yield a specified velocity distribution on the two walls. Velocity distributions are then used to calculate the cascade loss parameters. Calculated blade shapes are used primarily to determine whether the assumed velocity loadings are physically realistic. Model verification is accomplished by comparison of predicted turbine geometry and performance with an array of seven NASA single-stage axial gas turbine configurations.

A comprehensive method for preliminary design optimization of axial gas turbine stages

NASA Technical Reports Server (NTRS)

Jenkins, R. M.

1982-01-01

A method is presented that performs a rapid, reasonably accurate preliminary pitchline optimization of axial gas turbine annular flowpath geometry, as well as an initial estimate of blade profile shapes, given only a minimum of thermodynamic cycle requirements. No geometric parameters need be specified. The following preliminary design data are determined: (1) the optimum flowpath geometry, within mechanical stress limits; (2) initial estimates of cascade blade shapes; (3) predictions of expected turbine performance. The method uses an inverse calculation technique whereby blade profiles are generated by designing channels to yield a specified velocity distribution on the two walls. Velocity distributions are then used to calculate the cascade loss parameters. Calculated blade shapes are used primarily to determine whether the assumed velocity loadings are physically realistic. Model verification is accomplished by comparison of predicted turbine geometry and performance with four existing single stage turbines.

Effect of tip flange on tip leakage flow of small axial flow fans

NASA Astrophysics Data System (ADS)

Zhang, Li; Jin, Yingzi; Jin, Yuzhen

2014-02-01

Aerodynamic performance of an axial flow fan is closely related to its tip clearance leakage flow. In this paper, the hot-wire anemometer is used to measure the three dimensional mean velocity near the blade tips. Moreover, the filtered N-S equations with finite volume method and RNG k-ɛ turbulence model are adopted to carry out the steady simulation calculation of several fans that differ only in tip flange shape and number. The large eddy simulation and the FW-H noise models are adopted to carry out the unsteady numerical calculation and aerodynamic noise prediction. The results of simulation calculation agree roughly with that of tests, which proves the numerical calculation method is feasible.The effects of tip flange shapes and numbers on the blade tip vortex structure and the characteristics are analyzed. The results show that tip flange of the fan has a certain influence on the characteristics of the fan. The maximum efficiencies for the fans with tip flanges are shifted towards partial flow with respect to the design point of the datum fan. Furthermore, the noise characteristics for the fans with tip flanges have become more deteriorated than that for the datum fan. Tip flange contributes to forming tip vortex shedding and the effect of the half-cylinder tip flange on tip vortex shedding is obvious. There is a distinct relationship between the characteristics of the fan and tip vortex shedding. The research results provide the profitable reference for the internal flow mechanism of the performance optimization of small axial flow fans.

Probabilistic structural analysis methods for improving Space Shuttle engine reliability

NASA Technical Reports Server (NTRS)

Boyce, L.

1989-01-01

Probabilistic structural analysis methods are particularly useful in the design and analysis of critical structural components and systems that operate in very severe and uncertain environments. These methods have recently found application in space propulsion systems to improve the structural reliability of Space Shuttle Main Engine (SSME) components. A computer program, NESSUS, based on a deterministic finite-element program and a method of probabilistic analysis (fast probability integration) provides probabilistic structural analysis for selected SSME components. While computationally efficient, it considers both correlated and nonnormal random variables as well as an implicit functional relationship between independent and dependent variables. The program is used to determine the response of a nickel-based superalloy SSME turbopump blade. Results include blade tip displacement statistics due to the variability in blade thickness, modulus of elasticity, Poisson's ratio or density. Modulus of elasticity significantly contributed to blade tip variability while Poisson's ratio did not. Thus, a rational method for choosing parameters to be modeled as random is provided.

Energy efficient engine high pressure turbine test hardware detailed design report

NASA Technical Reports Server (NTRS)

Halila, E. E.; Lenahan, D. T.; Thomas, T. T.

1982-01-01

The high pressure turbine configuration for the Energy Efficient Engine is built around a two-stage design system. Moderate aerodynamic loading for both stages is used to achieve the high level of turbine efficiency. Flowpath components are designed for 18,000 hours of life, while the static and rotating structures are designed for 36,000 hours of engine operation. Both stages of turbine blades and vanes are air-cooled incorporating advanced state of the art in cooling technology. Direct solidification (DS) alloys are used for blades and one stage of vanes, and an oxide dispersion system (ODS) alloy is used for the Stage 1 nozzle airfoils. Ceramic shrouds are used as the material composition for the Stage 1 shroud. An active clearance control (ACC) system is used to control the blade tip to shroud clearances for both stages. Fan air is used to impinge on the shroud casing support rings, thereby controlling the growth rate of the shroud. This procedure allows close clearance control while minimizing blade tip to shroud rubs.

Simultaneous Boundary-Layer Transition, Tip Vortex, and Blade Deformation Measurements of a Rotor in Hover

NASA Technical Reports Server (NTRS)

Heineck, James; Schairer, Edward; Ramasamy, Manikandan; Roozeboom, Nettie

2016-01-01

This paper describes simultaneous optical measurements of a sub-scale helicopter rotor in the U.S. Army Hover Chamber at NASA Ames Research Center. The measurements included thermal imaging of the rotor blades to detect boundary layer transition; retro-reflective background-oriented schlieren (RBOS) to visualize vortices; and stereo photogrammetry to measure displacements of the rotor blades, to compute spatial coordinates of the vortices from the RBOS data, and to map the thermal imaging data to a three-dimensional surface grid. The test also included an exploratory effort to measure flow near the rotor tip by tomographic particle image velocimetry (tomo PIV)an effort that yielded valuable experience but little data. The thermal imaging was accomplished using an image-derotation method that allowed long integration times without image blur. By mapping the thermal image data to a surface grid it was possible to accurately locate transition in spatial coordinates along the length of the rotor blade.

Main rotor free wake geometry effects on blade air loads and response for helicopters in steady maneuvers. Volume 1: Theoretical formulation and analysis of results

NASA Technical Reports Server (NTRS)

Sadler, S. G.

1972-01-01

A mathematical model and computer program were implemented to study the main rotor free wake geometry effects on helicopter rotor blade air loads and response in steady maneuvers. The theoretical formulation and analysis of results are presented.

Mitigation of tip vortex cavitation by means of air injection on a Kaplan turbine scale model

NASA Astrophysics Data System (ADS)

Rivetti, A.; Angulo, M.; Lucino, C.; Liscia, S.

2014-03-01

Kaplan turbines operating at full-load conditions may undergo excessive vibration, noise and cavitation. In such cases, damage by erosion associated to tip vortex cavitation can be observed at the discharge ring. This phenomenon involves design features such as (1) overhang of guide vanes; (2) blade profile; (3) gap increasing size with blade opening; (4) suction head; (5) operation point; and (6) discharge ring stiffness, among others. Tip vortex cavitation may cause erosion at the discharge ring and draft tube inlet following a wavy pattern, in which the number of vanes can be clearly identified. Injection of pressurized air above the runner blade centerline was tested as a mean to mitigate discharge ring cavitation damage on a scale model. Air entrance was observed by means of a high-speed camera in order to track the air trajectory toward its mergence with the tip vortex cavitation core. Post-processing of acceleration signals shows that the level of vibration and the RSI frequency amplitude decrease proportionally with air flow rate injected. These findings reveal the potential mitigating effect of air injection in preventing cavitation damage and will be useful in further tests to be performed on prototype, aiming at determining the optimum air flow rate, size and distribution of the injectors.

Computer program for aerodynamic and blading design of multistage axial-flow compressors

NASA Technical Reports Server (NTRS)

Crouse, J. E.; Gorrell, W. T.

1981-01-01

A code for computing the aerodynamic design of a multistage axial-flow compressor and, if desired, the associated blading geometry input for internal flow analysis codes is presented. Compressible flow, which is assumed to be steady and axisymmetric, is the basis for a two-dimensional solution in the meridional plane with viscous effects modeled by pressure loss coefficients and boundary layer blockage. The radial equation of motion and the continuity equation are solved with the streamline curvature method on calculation stations outside the blade rows. The annulus profile, mass flow, pressure ratio, and rotative speed are input. A number of other input parameters specify and control the blade row aerodynamics and geometry. In particular, blade element centerlines and thicknesses can be specified with fourth degree polynomials for two segments. The output includes a detailed aerodynamic solution and, if desired, blading coordinates that can be used for internal flow analysis codes.

Axial compressor gas path design for desensitization of aerodynamic performance and stability to tip clearance

NASA Astrophysics Data System (ADS)

Cevik, Mert

Tip clearance is the necessary small gap left between the moving rotor tip and stationary shroud of a turbomachine. In a compressor, the pressure driven flow through this gap, called tip clearance flow, has a major and generally detrimental impact on compressor performance (pressure ratio and efficiency) and aerodynamic stability (stall margin). The increase in tip clearance, either temporary during transient engine operations or permanent from wear, leads to a drop in compressor performance and aerodynamic stability which results in a fuel consumption increase and a reduced operating envelope for a gas turbine engine. While much research has looked into increasing compressor performance and stall margin at the design (minimum or nominal) tip clearance, very little attention has been paid for reducing the sensitivity of these parameters to tip clearance size increase. The development of technologies that address this issue will lead to aircraft engines whose performance and operating envelope are more robust to operational demands and wear. The current research is the second phase of a research programme to develop design strategies to reduce the sensitivity of axial compressor performance and aerodynamic stability to tip clearance. The first phase had focused on blade design strategies and had led to the discovery and explanation of two flow features that reduces tip sensitivity, namely increased incoming meridional momentum in the rotor tip region and reduction/elimination of double leakage. Double leakage is the flow that exits one tip clearance and enters the tip clearance of the adjacent blade instead of convecting downstream out of the rotor passage. This flow was shown to be very detrimental to compressor performance and stall margin. Two rotor design strategies involving sweep and tip stagger reduction were proposed and shown by CFD simulations to exploit these features to reduce sensitivity. As the second phase, the objectives of the current research project are to develop gas path design strategies for axial compressors to achieve the same goal, to assess their ability to be combined with desensitizing axial compressor blade design strategies and to be applied to non-axial compressors. The search for gas path design strategies was based on the exploitation of the two flow desensitizing features listed above. Two gas path design strategies were proposed and analyzed. The first was gas path contouring in the form of a concave gas path to increase incoming tip meridional momentum.

Cost analysis of composite fan blade manufacturing processes

NASA Technical Reports Server (NTRS)

Stelson, T. S.; Barth, C. F.

1980-01-01

The relative manufacturing costs were estimated for large high technology fan blades prepared by advanced composite fabrication methods using seven candidate materials/process systems. These systems were identified as laminated resin matrix composite, filament wound resin matrix composite, superhybrid solid laminate, superhybrid spar/shell, metal matrix composite, metal matrix composite with a spar and shell, and hollow titanium. The costs were calculated utilizing analytical process models and all cost data are presented as normalized relative values where 100 was the cost of a conventionally forged solid titanium fan blade whose geometry corresponded to a size typical of 42 blades per disc. Four costs were calculated for each of the seven candidate systems to relate the variation of cost on blade size. Geometries typical of blade designs at 24, 30, 36 and 42 blades per disc were used. The impact of individual process yield factors on costs was also assessed as well as effects of process parameters, raw materials, labor rates and consumable items.

JT15D simulated flight data evaluation

NASA Technical Reports Server (NTRS)

Holm, R. G.

1984-01-01

The noise characteristics of the JT15D turbofan engine was analyzed with the objectives of: (1) assessing the state-of-art ability to simulate flight acoustic data using test results acquired in wind tunnel and outdoor (turbulence controlled) environments; and (2) predicting the farfield noise directivity of the blade passage frequency (BPF) tonal components using results from rotor blade mounted dynamic pressure instrumentation. Engine rotor tip speeds at subsonic, transonic, and supersonic conditions were evaluated. The ability to simulate flight results was generally within 2-3 dB for both outdoor and wind tunnel acoustic results. Some differences did occur in the broadband noise level and in the multiple-pure-tone harmonics at supersonic tip speeds. The prediction of blade passage frequency tone directivity from dynamic pressure measurements was accomplished for the three tip speed conditions. Predictions were made of the random and periodic components of the tone directivity. The technique for estimating the random tone component used hot wire data to establish a correlation between dynamic pressure and turbulence intensity. This prediction overestimated the tone level by typically 10 dB with the greatest overestimates occurring at supersonic conditions.

Computational Analysis of a Wells Turbine with Flexible Trailing Edges

NASA Astrophysics Data System (ADS)

Kincaid, Kellis; Macphee, David

2017-11-01

The Wells turbine is often used to produce a net positive power from an oscillating air column excited by ocean waves. It has been parametrically studied quite thoroughly in the past, both experimentally and numerically. The effects of various characteristics such as blade count and profile, solidity, and tip gap are well known. Several three-dimensional computational studies have been carried out using commercial code to investigate many phenomena detected in experiments: hysteresis, tip-gap drag, and post-stall behavior for example. In this work, the open-source code Foam-Extend is used to examine the effect of flexible blades on the performance of the Wells turbine. A new solver is created to integrate fluid-structure interaction into the code, allowing an accurate solution for both the solid and fluid domains. Reynolds-averaged governing equations are employed in a fully transient solution model. The elastic modulus of the flexible portion of the blade and the tip-gap width are varied, and the resulting flow fields are investigated to determine the cause of any performance differences. NSF Grant EEC 1659710.

Electrospark Deposition for Depot- and Field-Level Component Repair and Replacement of Hard Chromium Plating

DTIC Science & Technology

2006-09-07

aircraft repairs, including: 1. Single crystal turbine blade for two gas turbine engines ( FAA approved repair) 2. Second stage gas turbine blade ...gas turbine engine components. These include (a) application of corrosion resistant coatings to turbine blade tips where protective diffusion...base materials on many functional components (e.g., Ni-base superalloys, stainless steels, Monel, titanium alloys), thus allowing for self-repair

High Bypass Turbofan Component Development. Phase II. Fan Detail Design.

DTIC Science & Technology

1979-12-01

Vane metal angles ........ ..................... ... 18 22 Vane conical airfoil sections ..... ............... ... 19 23 Principal blade stresses at...31.25 deg. The number of rotor airfoils is 20 while the stator has 42 vanes . The number of vanes and the vane - blade spacing were consequences of...effect of radius change are accounted for. Figure 16 shows the blade hub, mean, and tip conical airfoil sections in engine orientation. For

An Architecture for On-Line Measurement of the Tip Clearance and Time of Arrival of a Bladed Disk of an Aircraft Engine

PubMed Central

Solís, Alejandro; Aranguren, Gerardo; Zubia, Joseba

2017-01-01

Safety and performance of the turbo-engine in an aircraft is directly affected by the health of its blades. In recent years, several improvements to the sensors have taken place to monitor the blades in a non-intrusive way. The parameters that are usually measured are the distance between the blade tip and the casing, and the passing time at a given point. Simultaneously, several techniques have been developed that allow for the inference—from those parameters and under certain conditions—of the amplitude and frequency of the blade vibration. These measurements are carried out on engines set on a rig, before being installed in an airplane. In order to incorporate these methods during the regular operation of the engine, signal processing that allows for the monitoring of those parameters at all times should be developed. This article introduces an architecture, based on a trifurcated optic sensor and a hardware processor, that fulfills this need. The proposed architecture is scalable and allows several sensors to be simultaneously monitored at different points around a bladed disk. Furthermore, the results obtained by the electronic system will be compared with the results obtained by the validation of the optic sensor. PMID:28934105

Aeroacoustic effects of reduced aft tip speed at constant thrust for a model counterrotation turboprop at takeoff conditions

NASA Technical Reports Server (NTRS)

Woodward, Richard P.; Hughes, Christopher E.

1990-01-01

A model high-speed, advanced counterrotation propeller, F7/A7, was tested in the NASA Lewis Research Center's 9- by 15-foot anechoic wind tunnel at simulated takeoff and approach conditions of Mach 0.2. The propeller was operated in a baseline configuration with the forward and aft rotor blade setting angles (36.2deg and 35.4 deg) and forward and aft rotational speeds essentially equal. Two additional configurations were tested with the aft rotor at increased blade setting angles and the rotational speed reduced to achieve overall performance similar to that of the baseline configuration. The aft rotor blade angles were adjusted such that the thrust and power absorption for each rotor remained the same as for the baseline configuration. Acoustic data were taken with an axially translating microphone probe that was attached to the tunnel floor. Concurrent aerodynamic data were taken to define propeller operating conditions. The aft rotor fundamental tone was about 6 dB lower with the 36.2 deg and 38.4 deg blade setting angles, and about 9 dB lower with the 36.2 and 41.4 deg blade setting angles. Predicted noise reductions based on tip speed considerations were 5 and 9.5 dB, respectively, for the two altered blade setting angles.

An Architecture for On-Line Measurement of the Tip Clearance and Time of Arrival of a Bladed Disk of an Aircraft Engine.

PubMed

Gil-García, José Miguel; Solís, Alejandro; Aranguren, Gerardo; Zubia, Joseba

2017-09-21

Safety and performance of the turbo-engine in an aircraft is directly affected by the health of its blades. In recent years, several improvements to the sensors have taken place to monitor the blades in a non-intrusive way. The parameters that are usually measured are the distance between the blade tip and the casing, and the passing time at a given point. Simultaneously, several techniques have been developed that allow for the inference-from those parameters and under certain conditions-of the amplitude and frequency of the blade vibration. These measurements are carried out on engines set on a rig, before being installed in an airplane. In order to incorporate these methods during the regular operation of the engine, signal processing that allows for the monitoring of those parameters at all times should be developed. This article introduces an architecture, based on a trifurcated optic sensor and a hardware processor, that fulfills this need. The proposed architecture is scalable and allows several sensors to be simultaneously monitored at different points around a bladed disk. Furthermore, the results obtained by the electronic system will be compared with the results obtained by the validation of the optic sensor.

Inducer analysis/pump model development

NASA Astrophysics Data System (ADS)

Cheng, Gary C.

1994-03-01

Current design of high performance turbopumps for rocket engines requires effective and robust analytical tools to provide design information in a productive manner. The main goal of this study was to develop a robust and effective computational fluid dynamics (CFD) pump model for general turbopump design and analysis applications. A finite difference Navier-Stokes flow solver, FDNS, which includes an extended k-epsilon turbulence model and appropriate moving zonal interface boundary conditions, was developed to analyze turbulent flows in turbomachinery devices. In the present study, three key components of the turbopump, the inducer, impeller, and diffuser, were investigated by the proposed pump model, and the numerical results were benchmarked by the experimental data provided by Rocketdyne. For the numerical calculation of inducer flows with tip clearance, the turbulence model and grid spacing are very important. Meanwhile, the development of the cross-stream secondary flow, generated by curved blade passage and the flow through tip leakage, has a strong effect on the inducer flow. Hence, the prediction of the inducer performance critically depends on whether the numerical scheme of the pump model can simulate the secondary flow pattern accurately or not. The impeller and diffuser, however, are dominated by pressure-driven flows such that the effects of turbulence model and grid spacing (except near leading and trailing edges of blades) are less sensitive. The present CFD pump model has been proved to be an efficient and robust analytical tool for pump design due to its very compact numerical structure (requiring small memory), fast turnaround computing time, and versatility for different geometries.

Inducer analysis/pump model development

NASA Technical Reports Server (NTRS)

Cheng, Gary C.

1994-01-01

Current design of high performance turbopumps for rocket engines requires effective and robust analytical tools to provide design information in a productive manner. The main goal of this study was to develop a robust and effective computational fluid dynamics (CFD) pump model for general turbopump design and analysis applications. A finite difference Navier-Stokes flow solver, FDNS, which includes an extended k-epsilon turbulence model and appropriate moving zonal interface boundary conditions, was developed to analyze turbulent flows in turbomachinery devices. In the present study, three key components of the turbopump, the inducer, impeller, and diffuser, were investigated by the proposed pump model, and the numerical results were benchmarked by the experimental data provided by Rocketdyne. For the numerical calculation of inducer flows with tip clearance, the turbulence model and grid spacing are very important. Meanwhile, the development of the cross-stream secondary flow, generated by curved blade passage and the flow through tip leakage, has a strong effect on the inducer flow. Hence, the prediction of the inducer performance critically depends on whether the numerical scheme of the pump model can simulate the secondary flow pattern accurately or not. The impeller and diffuser, however, are dominated by pressure-driven flows such that the effects of turbulence model and grid spacing (except near leading and trailing edges of blades) are less sensitive. The present CFD pump model has been proved to be an efficient and robust analytical tool for pump design due to its very compact numerical structure (requiring small memory), fast turnaround computing time, and versatility for different geometries.

Examination of propeller sound production using large eddy simulation

NASA Astrophysics Data System (ADS)

Keller, Jacob; Kumar, Praveen; Mahesh, Krishnan

2018-06-01

The flow field of a five-bladed marine propeller operating at design condition, obtained using large eddy simulation, is used to calculate the resulting far-field sound. The results of three acoustic formulations are compared, and the effects of the underlying assumptions are quantified. The integral form of the Ffowcs-Williams and Hawkings (FW-H) equation is solved on the propeller surface, which is discretized into a collection of N radial strips. Further assumptions are made to reduce FW-H to a Curle acoustic analogy and a point-force dipole model. Results show that although the individual blades are strongly tonal in the rotor plane, the propeller is acoustically compact at low frequency and the tonal sound interferes destructively in the far field. The propeller is found to be acoustically compact for frequencies up to 100 times the rotation rate. The overall far-field acoustic signature is broadband. Locations of maximum sound of the propeller occur along the axis of rotation both up and downstream. The propeller hub is found to be a source of significant sound to observers in the rotor plane, due to flow separation and interaction with the blade-root wakes. The majority of the propeller sound is generated by localized unsteadiness at the blade tip, which is caused by shedding of the tip vortex. Tonal blade sound is found to be caused by the periodic motion of the loaded blades. Turbulence created in the blade boundary layer is convected past the blade trailing edge leading to generation of broadband noise along the blade. Acoustic energy is distributed among higher frequencies as local Reynolds number increases radially along the blades. Sound source correlation and spectra are examined in the context of noise modeling.

A non-uniformly under-sampled blade tip-timing signal reconstruction method for blade vibration monitoring.

PubMed

Hu, Zheng; Lin, Jun; Chen, Zhong-Sheng; Yang, Yong-Min; Li, Xue-Jun

2015-01-22

High-speed blades are often prone to fatigue due to severe blade vibrations. In particular, synchronous vibrations can cause irreversible damages to the blade. Blade tip-timing methods (BTT) have become a promising way to monitor blade vibrations. However, synchronous vibrations are unsuitably monitored by uniform BTT sampling. Therefore, non-equally mounted probes have been used, which will result in the non-uniformity of the sampling signal. Since under-sampling is an intrinsic drawback of BTT methods, how to analyze non-uniformly under-sampled BTT signals is a big challenge. In this paper, a novel reconstruction method for non-uniformly under-sampled BTT data is presented. The method is based on the periodically non-uniform sampling theorem. Firstly, a mathematical model of a non-uniform BTT sampling process is built. It can be treated as the sum of certain uniform sample streams. For each stream, an interpolating function is required to prevent aliasing in the reconstructed signal. Secondly, simultaneous equations of all interpolating functions in each sub-band are built and corresponding solutions are ultimately derived to remove unwanted replicas of the original signal caused by the sampling, which may overlay the original signal. In the end, numerical simulations and experiments are carried out to validate the feasibility of the proposed method. The results demonstrate the accuracy of the reconstructed signal depends on the sampling frequency, the blade vibration frequency, the blade vibration bandwidth, the probe static offset and the number of samples. In practice, both types of blade vibration signals can be particularly reconstructed by non-uniform BTT data acquired from only two probes.

Laser-based gluing of diamond-tipped saw blades

NASA Astrophysics Data System (ADS)

Hennigs, Christian; Lahdo, Rabi; Springer, André; Kaierle, Stefan; Hustedt, Michael; Brand, Helmut; Wloka, Richard; Zobel, Frank; Dültgen, Peter

2016-03-01

To process natural stone such as marble or granite, saw blades equipped with wear-resistant diamond grinding segments are used, typically joined to the blade by brazing. In case of damage or wear, they must be exchanged. Due to the large energy input during thermal loosening and subsequent brazing, the repair causes extended heat-affected zones with serious microstructure changes, resulting in shape distortions and disadvantageous stress distributions. Consequently, axial run-out deviations and cutting losses increase. In this work, a new near-infrared laser-based process chain is presented to overcome the deficits of conventional brazing-based repair of diamond-tipped steel saw blades. Thus, additional tensioning and straightening steps can be avoided. The process chain starts with thermal debonding of the worn grinding segments, using a continuous-wave laser to heat the segments gently and to exceed the adhesive's decomposition temperature. Afterwards, short-pulsed laser radiation removes remaining adhesive from the blade in order to achieve clean joining surfaces. The third step is roughening and activation of the joining surfaces, again using short-pulsed laser radiation. Finally, the grinding segments are glued onto the blade with a defined adhesive layer, using continuous-wave laser radiation. Here, the adhesive is heated to its curing temperature by irradiating the respective grinding segment, ensuring minimal thermal influence on the blade. For demonstration, a prototype unit was constructed to perform the different steps of the process chain on-site at the saw-blade user's facilities. This unit was used to re-equip a saw blade with a complete set of grinding segments. This saw blade was used successfully to cut different materials, amongst others granite.

A Non-Uniformly Under-Sampled Blade Tip-Timing Signal Reconstruction Method for Blade Vibration Monitoring

PubMed Central

Hu, Zheng; Lin, Jun; Chen, Zhong-Sheng; Yang, Yong-Min; Li, Xue-Jun

2015-01-01

High-speed blades are often prone to fatigue due to severe blade vibrations. In particular, synchronous vibrations can cause irreversible damages to the blade. Blade tip-timing methods (BTT) have become a promising way to monitor blade vibrations. However, synchronous vibrations are unsuitably monitored by uniform BTT sampling. Therefore, non-equally mounted probes have been used, which will result in the non-uniformity of the sampling signal. Since under-sampling is an intrinsic drawback of BTT methods, how to analyze non-uniformly under-sampled BTT signals is a big challenge. In this paper, a novel reconstruction method for non-uniformly under-sampled BTT data is presented. The method is based on the periodically non-uniform sampling theorem. Firstly, a mathematical model of a non-uniform BTT sampling process is built. It can be treated as the sum of certain uniform sample streams. For each stream, an interpolating function is required to prevent aliasing in the reconstructed signal. Secondly, simultaneous equations of all interpolating functions in each sub-band are built and corresponding solutions are ultimately derived to remove unwanted replicas of the original signal caused by the sampling, which may overlay the original signal. In the end, numerical simulations and experiments are carried out to validate the feasibility of the proposed method. The results demonstrate the accuracy of the reconstructed signal depends on the sampling frequency, the blade vibration frequency, the blade vibration bandwidth, the probe static offset and the number of samples. In practice, both types of blade vibration signals can be particularly reconstructed by non-uniform BTT data acquired from only two probes. PMID:25621612

Prediction of dynamic and aerodynamic characteristics of the centrifugal fan with forward curved blades

NASA Astrophysics Data System (ADS)

Polanský, Jiří; Kalmár, László; Gášpár, Roman

2013-12-01

The main aim of this paper is determine the centrifugal fan with forward curved blades aerodynamic characteristics based on numerical modeling. Three variants of geometry were investigated. The first, basic "A" variant contains 12 blades. The geometry of second "B" variant contains 12 blades and 12 semi-blades with optimal length [1]. The third, control variant "C" contains 24 blades without semi-blades. Numerical calculations were performed by CFD Ansys. Another aim of this paper is to compare results of the numerical simulation with results of approximate numerical procedure. Applied approximate numerical procedure [2] is designated to determine characteristics of the turbulent flow in the bladed space of a centrifugal-flow fan impeller. This numerical method is an extension of the hydro-dynamical cascade theory for incompressible and inviscid fluid flow. Paper also partially compares results from the numerical simulation and results from the experimental investigation. Acoustic phenomena observed during experiment, during numerical simulation manifested as deterioration of the calculation stability, residuals oscillation and thus also as a flow field oscillation. Pressure pulsations are evaluated by using frequency analysis for each variant and working condition.

Adjoint Airfoil Optimization of Darrieus-Type Vertical Axis Wind Turbine

NASA Astrophysics Data System (ADS)

Fuchs, Roman; Nordborg, Henrik

2012-11-01

We present the feasibility of using an adjoint solver to optimize the torque of a Darrieus-type vertical axis wind turbine (VAWT). We start with a 2D cross section of a symmetrical airfoil and restrict us to low solidity ratios to minimize blade vortex interactions. The adjoint solver of the ANSYS FLUENT software package computes the sensitivities of airfoil surface forces based on a steady flow field. Hence, we find the torque of a full revolution using a weighted average of the sensitivities at different wind speeds and angles of attack. The weights are computed analytically, and the range of angles of attack is given by the tip speed ratio. Then the airfoil geometry is evolved, and the proposed methodology is evaluated by transient simulations.

Aerodynamic analysis of the Darrieus rotor including secondary effects

NASA Astrophysics Data System (ADS)

Paraschivoiu, I.; Delclaux, F.; Fraunie, P.; Beguier, C.

1983-10-01

An aerodynamic analysis is made of two variants of the two-actuator-disk theory for modeling the Darrieus wind turbine. The double-multiple-streamtube model with constant and variable interference factors, including secondary effects, is examined for a Darrieus rotor. The influence of the secondary effects, namely, the blade geometry and profile type, the rotating tower, and the presence of struts and aerodynamic spoilers, is relatively significant, especially at high tip-speed ratios. Variation of the induced velocity as a function of the azimuthal angle allows a more accurate calculation of the aerodynamic loads on the downwind zone of the rotor with respect to the assumed constant interference factors. The theoretical results were compared with available experimental data for the Magdalen Islands wind turbine and Sandia-type machines (straight-line/circular-arc shape).

Design and performance of controlled-diffusion stator compared with original double-circular-arc stator

NASA Technical Reports Server (NTRS)

Gelder, Thomas F.; Schmidt, James F.; Suder, Kenneth L.; Hathaway, Michael D.

1987-01-01

The capabilities of two stators, one with controlled-diffusion (CD) blade sections and one with double-circular-arc (DCA) blade sections, were compared. A CD stator was designed and tested that had the same chord length but half the blades of the DCA stator. The same fan rotor (tip speed, 429 m/sec; pressure ratio, 1.65) was used with each stator row. The design and analysis system is briefly described. The overall stage and rotor performances with each stator are compared, as are selected blade element data. The minimum overall efficiency decrement across the stator was approximately 1 percentage point greater with the CD blade sections than with the DCA blade sections.

Vibration and loads in hingeless rotors. Volume 1: Theoretical analyses

NASA Technical Reports Server (NTRS)

Watts, G. A.; London, R. J.

1972-01-01

Analytic methods are developed for calculating blade loads and shaft-transmitted vibratory forces in stiff bladed hingeless rotors operating at advance ratios from mu = .3 to mu = 2.0. Calculated shaft harmonic moments compared well with experimental values when the blade first flap frequency was in the region of two-per-revolution harmonic excitation. Calculated blade bending moment azimuthal distributions due to changes in cyclic pitch agreed well with experiment at radial stations near the blade root at values of the ratio of first flap frequency to rotor rotation rate from 1.5 to 5.0. At stations near the blade tip good agreement was only obtained at the higher values of frequency ratio.

Development and application of a method for predicting rotor free wake positions and resulting rotor blade air loads. Volume 1: Model and results

NASA Technical Reports Server (NTRS)

Sadler, S. G.

1971-01-01

Rotor wake geometries are predicted by a process similar to the startup of a rotor in a free stream. An array of discrete trailing and shed vortices is generated with vortex strengths corresponding to stepwise radial and azimuthal blade circulations. The array of shed and trailing vortices is limited to an arbitrary number of azimuthal steps behind each blade. The remainder of the wake model of each blade is an arbitrary number of trailing vortices. Vortex element end points were allowed to be transported by the resultant velocity of the free stream and vortex-induced velocities. Wake geometry, wake flow, and wake-induced velocity influence coefficients are generated by this program for use in the blade loads portion of the calculations. Blade loads computations include the effects of nonuniform inflow due to a free wake, nonlinear airfoil characteristics, and response of flexible blades to the applied loads. Computed wake flows and blade loads are compared with experimentally measured data. Predicted blade loads, response and shears and moments are obtained for a model rotor system having two independent rotors. The effects of advance ratio, vertical separation of rotors, different blade radius ratios, and different azimuthal spacing of the blades of one rotor with respect to the other are investigated.

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

NASA Technical Reports Server (NTRS)

Hah, Chunill; Katz, Joseph

2012-01-01

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.

Flow fields behind a variable-area nozzle for radial turbines

NASA Astrophysics Data System (ADS)

Hayami, Hiroshi; Hyun, Yong-Ik; Senoo, Yasutoshi; Yamaguchi, Michiteru

The flow fields behind a variable-area nozzle for radial turbines were measured in detail using a three-hole cobra probe in 15 cases, which are a combination of three nozzle throat areas (0.8, 1.0, and 1.4 times the rated area) and five values of the tip-clearance to blade-height ratio (between 0.0 to 0.099). The flow fields at different tip clearances are presented in contour maps, and the pitch mean values are discussed as spanwise distributions of total pressure loss, flow angle, and radial and tangential velocity components. It is shown that the intensity of swirl behind the nozzle is decreased and the pressure loss is increased with the tip clearance, and the effect is magnified as the blade loading is higher.

Research turbine for high-temperature core engine application. 2: Effect of rotor tip clearance on overall performance

NASA Technical Reports Server (NTRS)

Szanca, E. M.; Behning, F. P.; Schum, H. J.

1974-01-01

A 25.4-cm (10-in) tip diameter turbine was tested to determine the effect of rotor radial tip clearance on turbine overall performance. The test turbine was a half-scale model of a 50.8-cm-(20-in.-) diameter research turbine designed for high-temperature core engine application. The test turbine was fabricated with solid vanes and blades with no provision for cooling air and tested at much reduced inlet conditions. The tests were run at design speed over a range of pressure ratios for three different rotor clearances ranging from 2.3 to 6.7 percent of the annular blade passage height. The results obtained are compared to the results obtained with three other turbines of varying amounts of reaction.

The effect of forward skewed rotor blades on aerodynamic and aeroacoustic performance of axial-flow fan

NASA Astrophysics Data System (ADS)

Wei, Jun; Zhong, Fangyuan

Based on comparative experiment, this paper deals with using tangentially skewed rotor blades in axial-flow fan. It is seen from the comparison of the overall performance of the fan with skewed bladed rotor and radial bladed rotor that the skewed blades operate more efficiently than the radial blades, especially at low volume flows. Meanwhile, decrease in pressure rise and flow rate of axial-flow fan with skewed rotor blades is found. The rotor-stator interaction noise and broadband noise of axial-flow fan are reduced with skewed rotor blades. Forward skewed blades tend to reduce the accumulation of the blade boundary layer in the tip region resulting from the effect of centrifugal forces. The turning of streamlines from the outer radius region into inner radius region in blade passages due to the radial component of blade forces of skewed blades is the main reason for the decrease in pressure rise and flow rate.

Design and performance of a 427-meter-per-second-tip-speed two-stage fan having a 2.40 pressure ratio

NASA Technical Reports Server (NTRS)

Cunnan, W. S.; Stevans, W.; Urasek, D. C.

1978-01-01

The aerodynamic design and the overall and blade-element performances are presented of a 427-meter-per-second-tip-speed two-stage fan designed with axially spaced blade rows to reduce noise transmitted upstream of the fan. At design speed the highest recorded adiabatic efficiency was 0.796 at a pressure of 2.30. Peak efficiency was not established at design speed because of a damper failure which terminated testing prematurely. The overall efficiencies, at 60 and 80 percent of design speed, peaked at approximately 0.83.

Rotating stall investigation of 0.72 hub-tip ratio single-stage compressor

NASA Technical Reports Server (NTRS)

Graham, Robert W; Prian, Vasily D

1954-01-01

The rotating stall characteristics of a 0.72 hub-tip ratio, single-stage compressor were investigated. The stage was a 14-inch-diameter replica of the fourth stage of an experimental multistage compressor. No similarity existed between the frequency and propagation rate of the stall patterns observed in the single-stage replica and those observed in the multistage compressor after the fourth stage. A fatigue failure of the rotor blades occurred during the testing which was attributed to a resonance between the stall frequency and the natural bending frequency of the blades.

The effects of free stream turbulence on the flow field through a compressor cascade

NASA Astrophysics Data System (ADS)

Muthanna Kolera, Chittiappa

The flow through a compressor cascade with tip leakage has been studied experimentally. The cascade of GE rotor B section blades had an inlet angle of 65.1°, a stagger angle of 56.9°, and a solidity of 1.08. The final turning angle of the cascade was 11.8°. This compressor configuration was representative of the core compressor of an aircraft engine. The cascade was operated with a tip gap of 1.65%, and operated at a Reynolds number based on the chord length (0.254 m) of 388,000. Measurements were made at 8 axial locations to reveal the structure of the flow as it evolved through the cascade. Measurements were also made to reveal the effects of grid generated turbulence on this flow. The data set is unique in that not only does it give a comparison of elevated free stream turbulence effects, but also documents the developing flow through the blade row of a compressor cascade with tip leakage. Measurements were made at a total of 8 locations 0.8, 0.23 axial chords upstream and 0, 0.27, 0.48, 0.77, 0.98, and 1.26 axial chords downstream of the leading edge of the blade row for both inflow turbulence cases. The measurements revealed the formation and development of the tip leakage vortex within the passage. The tip leakage vortex becomes apparent at approximately X/ca = 0.27 and dominated much of the endwall flow. The tip leakage vortex is characterized by high streamwise velocity deficits, high vorticity and high turbulence kinetic energy levels. The result showed that between 0.77 and 0.98 axial chords downstream of the leading edge, the vortex structure and behavior changes. The effects of grid generated turbulence were also documented. The results revealed significant effects on the flow field. The results showed a 4% decrease in the blade loading and a 20% reduction in the vorticity levels within tip leakage vortex. There was also a shift in the vortex path, showing a shift close to the suction side with grid generated turbulence, indicating the strength of the vortex was decreased. Circulation calculations showed this reduction, and also indicated that the tip leakage vortex increased in size by about 30%. The results revealed that overall, the turbulence kinetic energy levels in the tip leakage vortex were increased, with the most drastic change occurring at X/ca = 0.77.

Probabilistic structural analysis to quantify uncertainties associated with turbopump blades

NASA Technical Reports Server (NTRS)

Nagpal, Vinod K.; Rubinstein, Robert; Chamis, Christos C.

1987-01-01

A probabilistic study of turbopump blades has been in progress at NASA Lewis Research Center for over the last two years. The objectives of this study are to evaluate the effects of uncertainties in geometry and material properties on the structural response of the turbopump blades to evaluate the tolerance limits on the design. A methodology based on probabilistic approach has been developed to quantify the effects of the random uncertainties. The results of this study indicate that only the variations in geometry have significant effects.

The Spectral and Statistical Properties of Turbulence Generated by a Vortex/Blade-Tip Interaction

NASA Technical Reports Server (NTRS)

Devenport, William J.; Wittmer, Kenneth S.; Wenger, Christian W.

1997-01-01

The perpendicular interaction of a streamwise vortex with the tip of a lifting blade was studied in incompressible flow to provide information useful to the accurate prediction of helicopter rotor noise and the understanding of vortex dominated turbulent flows. The vortex passed 0.3 chord lengths to the suction side of the blade tip, providing a weak interaction. Single and two-point turbulence measurements were made using sub-miniature four sensor hot-wire probes 15 chord lengths downstream of the blade trailing edge; revealing the mean velocity and Reynolds stress tensor distributions of the turbulence, as well as its spanwise length scales as a function of frequency. The single point measurements show the flow downstream of the blade to be dominated by the interaction of the original tip vortex and the vortex shed by the blade. These vortices rotate about each other under their mutual induction, winding up the turbulent wakes of the blades. This interaction between the vortices appears to be the source of new turbulence in their cores and in the region between them. This turbulence appears to be responsible for some decay in the core of the original vortex, not seen when the blade is removed. The region between the vortices is not only a region of comparatively large stresses, but also one of intense turbulence production. Velocity autospectra measured near its center suggests the presence quasi-periodic large eddies with axes roughly parallel to a line joining the vortex cores. Detailed two-point measurements were made on a series of spanwise cuts through the flow so as to reveal the turbulence scales as they would be seen along the span of an intersecting airfoil. The measurements were made over a range of probe separations that enabled them to be analyzed not only in terms of coherence and phase spectra but also in terms of wave-number frequency (kappa-omega) spectra, computed by transforming the measured cross-spectra with respect to the spanwise separation of the probes. These data clearly show the influence of the coherent eddies in the spiral wake and the turbulent region between the cores. These eddies produce distinct peaks in the upwash velocity kappa-omega spectra, and strong anisotropy manifested both in the decay of the kappa-omega spectrum at larger wave-numbers and in differences between the kappa-omega spectra of different components. None of these features are represented in the von Karman spectrum for isotropic turbulence that is often used in broadband noise computations. Wave-number frequency spectra measured in the cores appear to show some evidence that the turbulence outside sets tip core waves, as has previously been hypothesized. These spectra also provide for the first time a truly objective method for distinguishing velocity fluctuations produced by core wandering from other motions.

Two-dimensional compressible flow in centrifugal compressors with straight blades

NASA Technical Reports Server (NTRS)

Stanitz, John D; Ellis, Gaylord O

1950-01-01

Six numerical examples are presented for steady, two-dimensional, compressible, nonviscous flow in centrifugal compressors with thin straight blades, the center lines of which generate the surface of a right circular cone when rotated about the axis of the compressor. A seventh example is presented for incompressible flow. The solutions were obtained in a region of the compressors, including the impeller tip, that was considered to be unaffected by the diffuser vanes or by the impeller-inlet configuration. Each solution applies to radial and mixed flow compressors with various cone angles but with the same angle between blades on the conic flow surface. The solution also apply to radial and mixed flow turbines with the rotation and the flow direction reversed. The effects of variations in the following parameters were investigated: (1) flow rate, (2) impeller-tip speed, (3) variation of passage height with radius, and (4) angle between blades on conic flow surface. The numerical results are presented in plots of the streamlines and constant Mach number lines. Correlation equations are developed whereby the flow conditions in any impeller with straight blades can be determined (in the region investigated by this analysis) for all operating conditions.

JT8D revised high-pressure turbine cooling and other outer air seal program

NASA Technical Reports Server (NTRS)

Gaffin, W. O.

1979-01-01

The JT8D high pressure turbine was revised to reduce leakage between the blade tip shrouds and the outer air seal, and engine testing was performed to determine the effect on performance. The addition of a second knife-edge on the blade tip shroud, the extension of the honeycomb seal land to cover the added knife-edge and an existing spoiler on the shroud, and a material substitution in the seal support ring to improve thermal growth characteristics are included. A relocation of the blade cooling air discharge to insure adequate cooling flow is required. Significant specific fuel consumption and exhaust gas temperature improvements were demonstrated with the revised turbine in sea level and simulated altitude engine tests. Inspection of the revised seal hardware after these tests showed no unusual wear or degradation.

Tip Vortices of Isolated Wings and Helicopter Rotor Blades.

DTIC Science & Technology

1987-12-01

root to tip, as expected due to the induced downwash of the tip vor- tex and wake vortex sheet. Although the three different tip-caps produce very...the inherent limitation of not being able to model the vortex wake with these equations, although the Euler formulation has in it the necessary...physics to model vorticity transport correctly. These equations basically lack the physical mecha- nism needed to generate the vortex wake . However, in

Single-stage experimental evaluation of tandem-airfoil rotor stator blading for compressors. Part 6: Data and performance for stage D

NASA Technical Reports Server (NTRS)

Clemmons, D. R.

1973-01-01

An axial flow compressor stage, having single-airfoil blading, was designed for zero rotor prewhirl, constant rotor work across the span, and axial discharge flow. The stage was designed to produce a pressure ratio of 1.265 at a rotor tip velocity of 757 ft/sec. The rotor had an inlet hub/tip ratio of 0.8. The design procedure accounted for the rotor inlet boundary layer and included the effects of axial velocity ratio and secondary flow on blade row performance. The objectives of this experimental program were: (1) to obtain performance with uniform and distorted inlet flow for comparison with the performance of a stage consisting of tandem-airfoil blading designed for the same vector diagrams; and (2) to evaluate the effectiveness of accounting for the inlet boundary layer, axial velocity ratio, and secondary flows in the stage design. With uniform inlet flow, the rotor achieved a maximum adiabatic efficiency of 90.1% at design equivalent rotor speed and a pressure ratio of 1.281. The stage maximum adiabatic efficiency at design equivalent rotor speed with uniform inlet flow was 86.1% at a pressure ratio of 1.266. Hub radial, tip radial, and circumferential distortion of the inlet flow caused reductions in surge pressure ratio of approximately 2, 10 and 5%, respectively, at design rotor speed.

Experimental investigation of turbine blade-tip excitation forces

NASA Technical Reports Server (NTRS)

Martinez-Sanchez, Manuel; Jaroux, Belgacem; Song, Seung Jin; Yoo, Soom-Yung; Palczynski, Taras

1994-01-01

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.

Ultra high tip speed (670.6 m/sec) fan stage with composite rotor: Aerodynamic and mechanical design

NASA Technical Reports Server (NTRS)

Halle, J. E.; Burger, G. D.; Dundas, R. E.

1977-01-01

A highly loaded, single-stage compressor having a tip speed of 670.6 m/sec was designed for the purpose of investigating very high tip speeds and high aerodynamic loadings to obtain high stage pressure ratios at acceptable levels of efficiency. The design pressure ratio is 2.8 at an adiabatic efficiency of 84.4%. Corrected design flow is 83.4 kg/sec; corrected design speed is 15,200 rpm; and rotor inlet tip diameter is 0.853 m. The rotor uses multiple-circular-arc airfoils from 0 to 15% span, precompression airfoils assuming single, strong oblique shocks from 21 to 43% span, and precompression airfoils assuming multiple oblique shocks from 52% span to the tip. Because of the high tip speeds, the rotor blades are designed to be fabricated of composite materials. Two composite materials were investigated: Courtaulds HTS graphite fiber in a Kerimid 601 polyimide matrix and the same fibers in a PMR polyimide matrix. In addition to providing a description of the aerodynamic and mechanical design of the 670.0 m/sec fan, discussion is presented of the results of structural tests of blades fabricated with both types of matrices.