Theoretical analysis of linearized acoustics and aerodynamics of advanced supersonic propellers

NASA Technical Reports Server (NTRS)

Farassat, F.

1985-01-01

The derivation of a formula for prediction of the noise of supersonic propellers using time domain analysis is presented. This formula is a solution of the Ffowcs Williams-Hawkings equation and does not have the Doppler singularity of some other formulations. The result presented involves some surface integrals over the blade and line integrals over the leading and trailing edges. The blade geometry, motion and surface pressure are needed for noise calculation. To obtain the blade surface pressure, the observer is moved onto the blade surface and a linear singular integral equation is derived which can be solved numerically. Two examples of acoustic calculations using a computer program are currently under development.

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.

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

NASA Technical Reports Server (NTRS)

Bunker, Ronald S.; Bailey, Jeremy C.; Ameri, Ali A.

1999-01-01

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

Deployment of a Pressure Sensitive Paint System for Measuring Global Surface Pressures on Rotorcraft Blades in Simulated Forward Flight: Preliminary PSP Results from Test 581 in the 14- by 22-Foot Subsonic Tunnel

NASA Technical Reports Server (NTRS)

Watkins, Anthony Neal; Leighty, Bradley D.; Lipford, William E.; Wong, Oliver D.; Goodman, Kyle Z.; Crafton, James; Forlines, Alan; Goss, Larry; Gregory, James W.; Juliano, Thomas J.

2011-01-01

This report will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on the tips of rotorcraft blades in simulated forward flight at the 14- x 22-Foot Subsonic Tunnel. The system was designed to use a pulsed laser as an excitation source and PSP data was collected using the lifetime-based approach. With the higher intensity of the laser, this allowed PSP images to be acquired during a single laser pulse, resulting in the collection of crisp images that can be used to determine blade pressure at a specific instant in time. This is extremely important in rotorcraft applications as the blades experience dramatically different flow fields depending on their position in the rotor disk. Testing of the system was performed using the U.S. Army General Rotor Model System equipped with four identical blades. Two of the blades were instrumented with pressure transducers to allow for comparison of the results obtained from the PSP. This report will also detail possible improvements to the system.

The validation and application of a rotor acoustic prediction computer program

NASA Technical Reports Server (NTRS)

Gallman, Judith M.

1990-01-01

An essential prerequisite to reducing the acoustic detectability of military rotorcraft is a better understanding of main rotor noise which is the major contributor to the overall noise. A simple, yet accurate, Rotor Acoustic Prediction Program (RAPP) was developed to advance the understanding of main rotor noise. This prediction program uses the Ffowcs Williams and Hawkings (FW-H) equation. The particular form of the FW-H equation used is well suited for the coupling of the measured blade surface pressure to the prediction of acoustic pressure. The FW-H equation is an inhomogeneous wave equation that is valid in all space and governs acoustic pressure generated by thin moving bodies. The nonhomogeneous terms describe mass displacement due to surface motion and forces due to local surface stresses, such as viscous stress and pressure distribution on the surface. This paper examines two of the four types of main rotor noise: BVI noise and low-frequency noise. Blade-vortex interaction noise occurs when a tip vortex, previously shed by a rotor blade, passes close enough to a rotor blade to cause large variations in the blade surface pressures. This event is most disturbing when it happens on the advancing side of the rotor disk. Low-frequency noise includes hover and low to moderate speed forward flight. For these flight conditions, the low frequency components of the acoustic signal dominate.

Incidence loss for a core turbine rotor blade in a two-dimensional cascade

NASA Technical Reports Server (NTRS)

Stabe, R. G.; Kline, J. F.

1974-01-01

The effect of incidence angle on the aerodynamic performance of an uncooled core turbine rotor blade was investigated experimentally in a two-dimensional cascade. The cascade test covered a range of incidence angles from minus 15 deg to 15 deg in 5-degree increments and a range of pressure ratios corresponding to ideal exit critical velocity ratios of 0.6 to 0.95. The principal measurements were blade-surface static pressures and cross-channel surveys of exit total pressure, static pressure, and flow angle. The results of the investigation include blade-surface velocity distribution and overall performance in terms of weight flow and loss for the range of incidence angles and exit velocity ratios investigated. The measured losses are also compared with two common methods of predicting incidence loss.

Study on the Pressure Pulsation inside Runner with Splitter Blades in Ultra-High Head Turbine

NASA Astrophysics Data System (ADS)

Meng, L.; Zhang, S. P.; Zhou, L. J.; Wang, Z. W.

2014-03-01

Runners with splitter blades were used widely for the high efficiency and stability. In this paper, the unsteady simulation of an ultra-high head turbine at the best efficiency point, 50% and 75% discharge points were established, to analyze the pressure pulsation in the vaneless space, rotating domain and the draft tube. First of all, runners with different length splitter blades and without splitter blades were compared to learn the efficiency and the pressure distribution on the blade surface. And then the amplitude of the pressure pulsation was analysed. The peak efficiency of the runner with splitter blades is remarkably higher than that of the corresponding impeller without splitter blades. And the efficiency of the turbine is the highest when the length ratio of the splitter blades is 0.75 times the main blades. The pressure pulsation characteristics were also influenced, because the amplitudes of the pulsation induced by the RSI phenomenon were changed as a result of more blades. At last, the best design plan of the length of the splitter blades (length ratio=0.825) was obtained, which improved the pressure pulsation characteristics without significant prejudice to the efficiency.

Steady and unsteady blade stresses within the SSME ATD/HPOTP inducer

NASA Technical Reports Server (NTRS)

Gross, R. Steven

1994-01-01

There were two main goals of the ATD HPOTP (alternate turbopump development)(high pressure oxygen turbopump). First, determine the steady and unsteady inducer blade surface strains produced by hydrodynamic sources as a function of flow capacity (Q/N), suction specific speed (Nss), and Reynolds number (Re). Second, to identify the hydrodynamic source(s) of the unsteady blade strains. The reason the aforementioned goals are expressed in terms of blade strains as opposed to blade hydrodynamic pressures is because of the interest regarding the high cycle life of the inducer blades. This report focuses on the first goal of the test program which involves the determination of the steady and unsteady strain (stress) values at various points within the inducer blades. Strain gages were selected as the strain measuring devices. Concurrent with the experimental program, an analytical study was undertaken to produce a complete NASTRAN finite-element model of the inducer. Computational fluid dynamics analyses were utilized to provide the estimated steady-state blade surface pressure loading needed as load input to the NASTRAN inducer model.

Deployment of a Pressure Sensitive Paint System for Measuring Global Surface Pressures on Rotorcraft Blades in Simulated Forward Flight

NASA Technical Reports Server (NTRS)

Watkins, A. Neal; Leighty, Bradley; Lipford, William E.; Wong, Oliver D.; Goodman, Kyle Z.; Crafton, Jim; Forlines, Alan; Goss, Larry P.; Gregory, James W.; Juliano, Thomas J.

2012-01-01

This paper will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on the tips of rotorcraft blades in simulated forward flight at the 14- x 22-Foot Subsonic Tunnel at the NASA Langley Research Center. The system was designed to use a pulsed laser as an excitation source and PSP data was collected using the lifetime-based approach. With the higher intensity of the laser, this allowed PSP images to be acquired during a single laser pulse, resulting in the collection of crisp images that can be used to determine blade pressure at a specific instant in time. This is extremely important in rotorcraft applications as the blades experience dramatically different flow fields depending on their position in the rotor disk. Testing of the system was performed using the U.S. Army General Rotor Model System equipped with four identical blades. Two of the blades were instrumented with pressure transducers to allow for comparison of the results obtained from the PSP. Preliminary results show that the PSP agrees both qualitatively and quantitatively with both the expected results as well as with the pressure taps. Several areas of improvement have been indentified and are currently being developed.

Design and two dimensional cascade test of a jet-flap turbine stator blade with ratio of axial chord to spacing of 0.5

NASA Technical Reports Server (NTRS)

Stabe, R. G.

1971-01-01

A jet-flap blade was designed for a velocity diagram typical of the first-stage stator of a jet engine turbine and was tested in a simple two-dimensional cascade of six blades. The principal measurements were blade surface static pressure and cross-channel surveys of exit total pressure, static pressure, and flow angle. The results of the experimental investigation include blade loading, exit angle, flow, and loss data for a range of exit critical velocity ratios and three jet flow conditions.

Measurement of Unsteady Blade Surface Pressure on a Single Rotation Large Scale Advanced Prop-fan with Angular and Wake Inflow at Mach Numbers from 0.02 to 0.70

NASA Technical Reports Server (NTRS)

Bushnell, P.; Gruber, M.; Parzych, D.

1988-01-01

Unsteady blade surface pressure data for the Large-Scale Advanced Prop-Fan (LAP) blade operation with angular inflow, wake inflow and uniform flow over a range of inflow Mach numbers of 0.02 to 0.70 is provided. The data are presented as Fourier coefficients for the first 35 harmonics of shaft rotational frequency. Also presented is a brief discussion of the unsteady blade response observed at takeoff and cruise conditions with angular and wake inflow.

Unsteady blade pressures on a propfan at takeoff: Euler analysis and flight data

NASA Technical Reports Server (NTRS)

Nallasamy, M.

1991-01-01

The unsteady blade pressures due to the operation of the propfan at an angle to the direction of the mean flow are obtained by solving the unsteady three dimensional Euler equations. The configuration considered is the eight bladed SR7L propfan at takeoff conditions and the inflow angles considered are 6.3 deg, 8.3 deg, 11.3 deg. The predicted blade pressure waveforms are compared with inflight measurements. At the inboard radial station (r/R = 0.68) the phase of the predicted waveforms show reasonable agreement with the measurements while the amplitudes are over predicted in the leading edge region of the blade. At the outboard radial station (r/R = 0.95), the predicted amplitudes of the waveforms on the pressure surface are in good agreement with flight data for all inflow angles. The measured (installed propfan) waveforms show a relative phase lag compared to the computed (propfan alone) waveforms. The phase lag depends on the axial location of the transducer and the surface of the blade. On the suction surface, in addition to the relative phase lag, the measurements show distortion (widening and steepening) of the waveforms. The extent of distortion increases with increase in inflow angle. This distortion seems to be due to viscous separation effects which depend on the azimuthal location of the blade and the axial location of the transducer.

Heat transfer and pressure measurements and comparison with prediction for the SSME two-stage turbine

NASA Technical Reports Server (NTRS)

Dunn, M. G.; Kim, J.

1992-01-01

Time averaged Stanton number and surface pressure distributions are reported for the first stage vane row, the first stage blade row, and the second stage vane row of the Rocketdyne Space Shuttle Main Engine (SSME) two-stage fuel-side turbine. Unsteady pressure envelope measurements for the first blade are also reported. These measurements were made at 10 percent, 50 percent, and 90 percent span on both the pressure and suction surfaces of the first stage components. Additional Stanton number measurements were made on the first stage blade platform, blade tip, and shroud, and at 50 percent span on the second vane. A shock tube was used as a short duration source of heated and pressurized air to which the turbine was subjected. Platinum thin film heat flux gages were used to obtain the heat flux measurements, while miniature silicon diaphragm flush-mounted pressure transducers were used to obtain the pressure measurements. The first stage vane Stanton number distributions are compared with predictions obtained using a version of STAN5 and quasi-3D Navier-Stokes solution. This same quasi-3D N-S code was also used to obtain predictions for the first blade and the second vane.

Prediction of Relaminarization Effects on Turbine Blade Heat Transfer

NASA Technical Reports Server (NTRS)

Boyle, R. J.; Giel, P. W.

2001-01-01

An approach to predicting turbine blade heat transfer when turbulent flow relaminarizes due to strong favorable pressure gradients is described. Relaminarization is more likely to occur on the pressure side of a rotor blade. While stators also have strong favorable pressure gradients, the pressure surface is less likely to become turbulent at low to moderate Reynolds numbers. Accounting for the effects of relaminarization for blade heat transfer can substantially reduce the predicted rotor surface heat transfer. This in turn can lead to reduced rotor cooling requirements. Two-dimensional midspan Navier-Stokes analyses were done for each of eighteen test cases using eleven different turbulence models. Results showed that including relaminarization effects generally improved the agreement with experimental data. The results of this work indicate that relatively small changes in rotor shape can be utilized to extend the likelihood of relaminarization to high Reynolds numbers. Predictions showing how rotor blade heat transfer at a high Reynolds number can be reduced through relaminarization are given.

Heat transfer and pressure measurements for the SSME fuel turbine

NASA Technical Reports Server (NTRS)

Dunn, Michael G.; Kim, Jungho

1991-01-01

A measurement program is underway using the Rocketdyne two-stage Space Shuttle Main Engine (SSME) fuel turbine. The measurements use a very large shock tunnel to produce a short-duration source of heated and pressurized gas which is subsequently passed through the turbine. Within this environment, the turbine is operated at the design values of flow function, stage pressure ratio, stage temperature ratio, and corrected speed. The first stage vane row and the first stage blade row are instrumented in both the spanwise and chordwise directions with pressure transducers and heat flux gages. The specific measurements to be taken include time averaged surface pressure and heat flux distributions on the vane and blade, flow passage static pressure, flow passage total pressure and total temperature distributions, and phase resolved surface pressure and heat flux on the blade.

Limits on the prediction of helicopter rotor noise using thickness and loading sources: Validation of helicopter noise prediction techniques

NASA Technical Reports Server (NTRS)

Succi, G. P.

1983-01-01

The techniques of helicopter rotor noise prediction attempt to describe precisely the details of the noise field and remove the empiricisms and restrictions inherent in previous methods. These techniques require detailed inputs of the rotor geometry, operating conditions, and blade surface pressure distribution. The Farassat noise prediction techniques was studied, and high speed helicopter noise prediction using more detailed representations of the thickness and loading noise sources was investigated. These predictions were based on the measured blade surface pressures on an AH-1G rotor and compared to the measured sound field. Although refinements in the representation of the thickness and loading noise sources improve the calculation, there are still discrepancies between the measured and predicted sound field. Analysis of the blade surface pressure data indicates shocks on the blades, which are probably responsible for these discrepancies.

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.

Effect of Reynolds number, turbulence level and periodic wake flow on heat transfer on low pressure turbine blades.

PubMed

Suslov, D; Schulz, A; Wittig, S

2001-05-01

The development of effective cooling methods is of major importance for the design of new gas turbines blades. The conception of optimal cooling schemes requires a detailed knowledge of the heat transfer processes on the blade's surfaces. The thermal load of turbine blades is predominantly determined by convective heat transfer which is described by the local heat transfer coefficient. Heat transfer is closely related to the boundary layer development along the blade surface and hence depends on various flow conditions and geometrical parameters. Particularly Reynolds number, pressures gradient and turbulence level have great impact on the boundary layer development and the according heat transfer. Therefore, in the present study, the influence of Reynolds number, turbulence intensity, and periodic unsteady inflow on the local heat transfer of a typical low pressure turbine airfoil is experimentally examined in a plane cascade.

Rocket Engine Turbine Blade Surface Pressure Distributions Experiment and Computations

NASA Technical Reports Server (NTRS)

Hudson, Susan T.; Zoladz, Thomas F.; Dorney, Daniel J.; Turner, James (Technical Monitor)

2002-01-01

Understanding the unsteady aspects of turbine rotor flow fields is critical to successful future turbine designs. A technology program was conducted at NASA's Marshall Space Flight Center to increase the understanding of unsteady environments for rocket engine turbines. The experimental program involved instrumenting turbine rotor blades with miniature surface mounted high frequency response pressure transducers. The turbine model was then tested to measure the unsteady pressures on the rotor blades. The data obtained from the experimental program is unique in two respects. First, much more unsteady data was obtained (several minutes per set point) than has been possible in the past. Also, an extensive steady performance database existed for the turbine model. This allowed an evaluation of the effect of the on-blade instrumentation on the turbine's performance. A three-dimensional unsteady Navier-Stokes analysis was also used to blindly predict the unsteady flow field in the turbine at the design operating conditions and at +15 degrees relative incidence to the first-stage rotor. The predicted time-averaged and unsteady pressure distributions show good agreement with the experimental data. This unique data set, the lessons learned for acquiring this type of data, and the improvements made to the data analysis and prediction tools are contributing significantly to current Space Launch Initiative turbine airflow test and blade surface pressure prediction efforts.

Helicopter rotor rotational noise predictions based on measured high-frequency blade loads

NASA Technical Reports Server (NTRS)

Hosier, R. N.; Ramakrishnan, R.

1974-01-01

In tests conducted at the Langley helicopter rotor test facility, simultaneous measurements of up to 200 harmonics of the fluctuating aerodynamic blade surface pressures and far-field radiated noise were made on a full-scale nontranslating rotor system. After their characteristics were determined, the measured blade surface pressures were converted to loading coefficients and used in an existing theory to predict the far-field rotational noise. A comparison of the calculated and measured noise shows generally good agreement up to 300 to 600 Hz, depending on the discreteness of the loading spectrum. Specific attention is given to the effects of the blade loading coefficients, chordwise loading distributions, blade loading phases, and observer azimuthal position on the calculations.

Comprehensive helicopter rotor instrumentation - A retrofit approach using miniature transducers

NASA Technical Reports Server (NTRS)

Jacklin, Stephen A.; Mort, Ray; Morrison, Dwayne

1992-01-01

This paper reports an approach used to retrofit a set of full-scale main rotor blades with 290 miniature pressure transducers, 46 strain gages, and 24 miniature accelerometers. Normally, in order to avoid disturbing the aerodynamics of the rotor flow field, the pressure instrumentation must be integrally built into the body of the rotor blades. However, using a method developed with NASA, miniature pressure transducers are mounted to the blade exterior surface without degrading the quality of the blade aerodynamics. Moreover, it is estimated that this approach reduced costs by more than 50 percent over building a set of pressure instrumented blades. The aerodynamic measurement objectives are presented as are instrumentation design considerations, type of instrumentation used, assembly process, and the installed instrumentation characteristics.

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.

Development of a locally mass flux conservative computer code for calculating 3-D viscous flow in turbomachines

NASA Technical Reports Server (NTRS)

Walitt, L.

1982-01-01

The VANS successive approximation numerical method was extended to the computation of three dimensional, viscous, transonic flows in turbomachines. A cross-sectional computer code, which conserves mass flux at each point of the cross-sectional surface of computation was developed. In the VANS numerical method, the cross-sectional computation follows a blade-to-blade calculation. Numerical calculations were made for an axial annular turbine cascade and a transonic, centrifugal impeller with splitter vanes. The subsonic turbine cascade computation was generated in blade-to-blade surface to evaluate the accuracy of the blade-to-blade mode of marching. Calculated blade pressures at the hub, mid, and tip radii of the cascade agreed with corresponding measurements. The transonic impeller computation was conducted to test the newly developed locally mass flux conservative cross-sectional computer code. Both blade-to-blade and cross sectional modes of calculation were implemented for this problem. A triplet point shock structure was computed in the inducer region of the impeller. In addition, time-averaged shroud static pressures generally agreed with measured shroud pressures. It is concluded that the blade-to-blade computation produces a useful engineering flow field in regions of subsonic relative flow; and cross-sectional computation, with a locally mass flux conservative continuity equation, is required to compute the shock waves in regions of supersonic relative flow.

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.

Blade Surface Pressure Distributions in a Rocket Engine Turbine: Experimental Work With On-Blade Pressure Transducers

NASA Technical Reports Server (NTRS)

Hudson, Susan T.; Zoladz, Thomas F.; Griffin, Lisa W.; Turner, James E. (Technical Monitor)

2000-01-01

Understanding the unsteady aspects of turbine rotor flowfields is critical to successful future turbine designs. A technology program was conducted at NASA's Marshall Space Flight Center to increase the understanding of unsteady environments for rocket engine turbines. The experimental program involved instrumenting turbine rotor blades with surface-mounted high frequency response pressure transducers. The turbine model was then tested to measure the unsteady pressures on the rotor blades. The data obtained from the experimental program is unique in three respects. First, much more unsteady data was obtained (several minutes per set point) than has been possible in the past. Also, two independent unsteady data acquisition systems and fundamental signal processing approaches were used. Finally, an extensive steady performance database existed for the turbine model. This allowed an evaluation of the effect of the on-blade instrumentation on the turbine's performance. This unique data set, the lessons learned for acquiring this type of data, and the improvements made to the data analysis and prediction tools will contribute to future turbine programs such as those for reusable launch vehicles.

Pressure fluctuation generated by the interaction of blade and tongue

NASA Astrophysics Data System (ADS)

Zheng, Lulu; Dou, Hua-Shu; Chen, Xiaoping; Zhu, Zuchao; Cui, Baoling

2018-02-01

Pressure fluctuation around the tongue has large effect on the stable operation of a centrifugal pump. In this paper, the Reynolds averaged Navier-Stokes equations (RANS) and the RNG k-epsilon turbulence model is employed to simulate the flow in a pump. The flow field in the centrifugal pump is computed for a range of flow rate. The simulation results have been compared with the experimental data and good agreement has been achieved. In order to study the interaction of the tongue with the impeller, fifteen monitor probes are evenly distributed circumferentially at three radii around the tongue. Pressure distribution is investigated at various blade positions while the blade approaches to and leaves the tongue region. Results show that pressure signal fluctuates largely around the tongue, and it is more intense near the tongue surface. At design condition, standard deviation of pressure fluctuation is the minimum. At large flow rate, the increased low pressure region at the blade trailing edge results in the increases of pressure fluctuation amplitude and pressure spectra at the monitor probes. Minimum pressure is obtained when the blade is facing to the tongue. It is found that the amplitude of pressure fluctuation strongly depends on the blade positions at large flow rate, and pressure fluctuation is caused by the relative movement between blades and tongue. At small flow rate, the rule of pressure fluctuation is mainly depending on the structure of vortex flow at blade passage exit besides the influence from the relative position between the blade and the tongue.

A FORTRAN program for calculating three dimensional, inviscid and rotational flows with shock waves in axial compressor blade rows: User's manual

NASA Technical Reports Server (NTRS)

Thompkins, W. T., Jr.

1982-01-01

A FORTRAN-IV computer program was developed for the calculation of the inviscid transonic/supersonic flow field in a fully three dimensional blade passage of an axial compressor rotor or stator. Rotors may have dampers (part span shrouds). MacCormack's explicit time marching method is used to solve the unsteady Euler equations on a finite difference mesh. This technique captures shocks and smears them over several grid points. Input quantities are blade row geometry, operating conditions and thermodynamic quanities. Output quantities are three velocity components, density and internal energy at each mesh point. Other flow quanities are calculated from these variables. A short graphics package is included with the code, and may be used to display the finite difference grid, blade geometry and static pressure contour plots on blade to blade calculation surfaces or blade suction and pressure surfaces. The flow in a low aspect ratio transonic compressor was analyzed and compared with high response total pressure probe measurements and gas fluorescence static density measurements made in the MIT blowdown wind tunnel. These comparisons show that the computed flow fields accurately model the measured shock wave locations and overall aerodynamic performance.

Analysis and improvement of gas turbine blade temperature measurement error

NASA Astrophysics Data System (ADS)

Gao, Shan; Wang, Lixin; Feng, Chi; Daniel, Ketui

2015-10-01

Gas turbine blade components are easily damaged; they also operate in harsh high-temperature, high-pressure environments over extended durations. Therefore, ensuring that the blade temperature remains within the design limits is very important. In this study, measurement errors in turbine blade temperatures were analyzed, taking into account detector lens contamination, the reflection of environmental energy from the target surface, the effects of the combustion gas, and the emissivity of the blade surface. In this paper, each of the above sources of measurement error is discussed, and an iterative computing method for calculating blade temperature is proposed.

Measurement of unsteady surface pressure on rotor blades of fans by pressure-sensitive paint

NASA Astrophysics Data System (ADS)

Yokoyama, Hiroshi; Miura, Kouhei; Iida, Akiyoshi

2017-01-01

To clarify the unsteady pressure distributions on the rotor blades of an axial fan, a pressure-sensitive paint (PSP) technique was used. To capture the image of the rotating fan as a static image, an optical derotator method with a dove prism was adopted. It was confirmed by preliminary experiments with a resonator and a speaker that the pressure fluctuations with 347 Hz can be measured by the present PSP. The measured mean pressure distributions were compared with the predicted results based on large-eddy simulations. The measured instantaneous surface pressure is instrumental to identify acoustic source of fan noise in the design stage.

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

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

An experimental study of near wall flow parameters in the blade end-wall corner region

NASA Technical Reports Server (NTRS)

Bhargava, Rakesh K.; Raj, Rishi S.

1989-01-01

The near wall flow parameters in the blade end-wall corner region is investigated. The blade end-wall corner region was simulated by mounting an airfoil section (NACA 65-015 base profile) symmetric blades on both sides of the flat plate with semi-circular leading edge. The initial 7 cm from the leading edge of the flat plate was roughened by gluing No. 4 floor sanding paper to artificially increase the boundary layer thickness on the flat plate. The initial flow conditions of the boundary layer upstream of the corner region are expected to dictate the behavior of flow inside the corner region. Therefore, an experimental investigation was extended to study the combined effect of initial roughness and increased level of free stream turbulence on the development of a 2-D turbulent boundary layer in the absence of the blade. The measurement techniques employed in the present investigation included, the conventional pitot and pitot-static probes, wall taps, the Preston tube, piezoresistive transducer and the normal sensor hot-wire probe. The pitot and pitot-static probes were used to obtain mean velocity profile measurements within the boundary layer. The measurements of mean surface static pressure were obtained with the surface static tube and the conventional wall tap method. The wall shear vector measurements were made with a specially constructed Preston tube. The flush mounted piezoresistive type pressure transducer were employed to measure the wall pressure fluctuation field. The velocity fluctuation measurements, used in obtaining the wall pressure-velocity correlation data, were made with normal single sensor hot-wire probe. At different streamwise stations, in the blade end-wall corner region, the mean values of surface static pressure varied more on the end-wall surface in the corner region were mainly caused by the changes in the curvature of the streamlines. The magnitude of the wall shear stress in the blade end-wall corner region increased significantly in the close vicinity of the corner line. The maximum value of the wall shear stress and its location from the corner line, on both the surfaces forming the corner region, were observed to change along the corner. These observed changes in the maximum values of the wall shear stress and its location from the corner line could be associated with the stretching and attenuation of the horseshoe vortex. The wall shear stress vectors in the blade end-wall corner region were observed to be more skewed on the end-wall surface as compared to that on the blade surface. The differences in the wall shear stress directions obtained with the Preston tube and flow visualization method were within the range in which the Preston tube was found to be insensitive to the yaw angle.

Flow in a centrifugal fan impeller at off-design conditions

NASA Astrophysics Data System (ADS)

Wright, T.; Tzou, K. T. S.; Madhavan, S.

1984-06-01

A fully three-dimensional finite element analysis of inviscid, incompressible blade channel flow is the basis of the present study of both predicted and measured surface velocity and pressure distributions in the internal flow channels of a centrifugal fan impeller, for volume flow rates of 80-125 percent the design flow rate. The experimental results made extensive use of blade and sidewall surface pressure taps installed in a scale model of an airfoil-bladed centrifugal fan impeller. The results obtained illustrate the ability of both flow analyses to predict the dominant features of the impeller flow field, including peak blade surface velocities and adverse gradients at flows far from the design point. Insight is also gained into the limiting channel diffusion values for typical centrifugal cascade performance, together with the influence of viscous effects, as seen in deviations from ideal flow predictions.

Unsteady blade-surface pressures on a large-scale advanced propeller: Prediction and data

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Groeneweg, J. F.

1990-01-01

An unsteady 3-D Euler analysis technique is employed to compute the flow field of an advanced propeller operating at an angle of attack. The predicted blade pressure waveforms are compared with wind tunnel data at two Mach numbers, 0.5 and 0.2. The inflow angle is three degrees. For an inflow Mach number of 0.5, the predicted pressure response is in fair agreement with data: the predicted phases of the waveforms are in close agreement with data while the magnitudes are underpredicted. At the low Mach number of 0.2 (takeoff), the numerical solution shows the formation of a leading edge vortex which is in qualitative agreement with measurements. However, the highly nonlinear pressure response measured on the blade suction surface is not captured in the present inviscid analysis.

Unsteady blade surface pressures on a large-scale advanced propeller - Prediction and data

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Groeneweg, J. F.

1990-01-01

An unsteady three dimensional Euler analysis technique is employed to compute the flowfield of an advanced propeller operating at an angle of attack. The predicted blade pressure waveforms are compared with wind tunnel data at two Mach numbers, 0.5 and 0.2. The inflow angle is three degrees. For an inflow Mach number of 0.5, the predicted pressure response is in fair agreement with data: the predicted phases of the waveforms are in close agreement with data while the magnitudes are underpredicted. At the low Mach number of 0.2 (take-off) the numerical solution shows the formation of a leading edge vortex which is in qualitative agreement with measurements. However, the highly nonlinear pressure response measured on the blade suction surface is not captured in the present inviscid analysis.

Applying Pressure Sensitive Paint Technology to Rotor Blades

NASA Technical Reports Server (NTRS)

Watkins, A. Neal; Leighty, Bradley D.; Lipford, William E.; Goodman, Kyle Z.; Crafton, Jim; Gregory, James W.

2014-01-01

This report will present details of a Pressure Sensitive Paint (PSP) system for measuring global surface pressures on rotorcrtaft blades in simulated forward flight at the 14- by 22-Foot Subsonic Tunnel at the NASA Langley Research Center. The basics of the PSP method will be discussed and the modifications that were needed to extend this technology for use on rotor blades. Results from a series of tests will also be presented as well as several areas of improvement that have been identified and are currently being developed for future testing.

An analysis of the viscous flow through a compact radial turbine by the average passage approach

NASA Technical Reports Server (NTRS)

Heidmann, James D.; Beach, Timothy A.

1990-01-01

A steady, three-dimensional viscous average passage computer code is used to analyze the flow through a compact radial turbine rotor. The code models the flow as spatially periodic from blade passage to blade passage. Results from the code using varying computational models are compared with each other and with experimental data. These results include blade surface velocities and pressures, exit vorticity and entropy contour plots, shroud pressures, and spanwise exit total temperature, total pressure, and swirl distributions. The three computational models used are inviscid, viscous with no blade clearance, and viscous with blade clearance. It is found that modeling viscous effects improves correlation with experimental data, while modeling hub and tip clearances further improves some comparisons. Experimental results such as a local maximum of exit swirl, reduced exit total pressures at the walls, and exit total temperature magnitudes are explained by interpretation of the flow physics and computed secondary flows. Trends in the computed blade loading diagrams are similarly explained.

Determination and Use of the Local Recovery Factor for Calculating the Effectiveness Gas Temperature for Turbine Blades / Jack B. Esgar and Alfred L. Lea

NASA Technical Reports Server (NTRS)

Esgar, Jack B; Lea, Alfred L

1951-01-01

In an in experimental investigation of local recovery factors for a blade having a pressure distribution similar to that of a typical reaction-type turbine blade, it a was found that the recovery factors were essentially independent of Mach number, Reynolds number, pressure gradient, and position on the blade surface except for regions where the boundary layer was probably in the transition range from laminar to turbulent. The recommended value of local subsonic recovery factor for use in calculating the effective gas temperature for gas turbine blades was 0.89.

2-D and 3-D oscillating wing aerodynamics for a range of angles of attack including stall

NASA Technical Reports Server (NTRS)

Piziali, R. A.

1994-01-01

A comprehensive experimental investigation of the pressure distribution over a semispan wing undergoing pitching motions representative of a helicopter rotor blade was conducted. Testing the wing in the nonrotating condition isolates the three-dimensional (3-D) blade aerodynamic and dynamic stall characteristics from the complications of the rotor blade environment. The test has generated a very complete, detailed, and accurate body of data. These data include static and dynamic pressure distributions, surface flow visualizations, two-dimensional (2-D) airfoil data from the same model and installation, and important supporting blockage and wall pressure distributions. This body of data is sufficiently comprehensive and accurate that it can be used for the validation of rotor blade aerodynamic models over a broad range of the important parameters including 3-D dynamic stall. This data report presents all the cycle-averaged lift, drag, and pitching moment coefficient data versus angle of attack obtained from the instantaneous pressure data for the 3-D wing and the 2-D airfoil. Also presented are examples of the following: cycle-to-cycle variations occurring for incipient or lightly stalled conditions; 3-D surface flow visualizations; supporting blockage and wall pressure distributions; and underlying detailed pressure results.

CMC blade with pressurized internal cavity for erosion control

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

Garcia-Crespo, Andres; Goike, Jerome Walter

A ceramic matrix composite blade for use in a gas turbine engine having an airfoil with leading and trailing edges and pressure and suction side surfaces, a blade shank secured to the lower end of each airfoil, one or more interior fluid cavities within the airfoil having inlet flow passages at the lower end which are in fluid communication with the blade shank, one or more passageways in the blade shank corresponding to each one of the interior fluid cavities and a fluid pump (or compressor) that provides pressurized fluid (nominally cool, dry air) to each one of the interiormore » fluid cavities in each airfoil. The fluid (e.g., air) is sufficient in pressure and volume to maintain a minimum fluid flow to each of the interior fluid cavities in the event of a breach due to foreign object damage.« less

Effect of Favorable Pressure Gradients on Turbine Blade Pressure Surface Heat Transfer

NASA Technical Reports Server (NTRS)

Boyle, Robert J.; Giel, P. W.

2002-01-01

Recent measurements on a turbine rotor showed significant relaminarization effects. These effects were evident on the pressure surface heat transfer measurements. The character of the heat transfer varied with Reynolds number. Data were obtained for exit Reynolds numbers between 500,000 and 880,000. Tests were done with a high level of inlet turbulence, 7.5%. At lower Reynolds numbers the heat transfer was similar to that for laminar flow, but at a level higher than for laminar flow. At higher Reynolds numbers the heat transfer was similar to turbulent flow, when the acceleration parameter, K, was sufficiently small. The proposed paper discusses the experimental results, and also discusses approaches to calculating the surface heat transfer for the blade surface. Calculations were done using a three-dimensional Navier-Stokes CFD analysis. The results of these tests, when compared with previous blade tests in the same facility, illustrate modeling difficulties that were encountered in CFD predictions. The two blades were in many ways similar. However, the degree of agreement between the same analysis and the experimental data was significantly different. These differences are highlighted to illustrate where improvements in modeling approaches are needed for transitional flows.

Unsteady blade pressure measurements for the SR-7A propeller at cruise conditions

NASA Technical Reports Server (NTRS)

Heidelberg, L. J.; Nallasamy, M.

1990-01-01

The unsteady blade surface pressures were measured on the SR-7A propeller. The freestream Mach no., inflow angle, and advance ratio were varied while measurements were made at nine blade stations. At a freestream Mach no. of 0.8, the data in terms of unsteady pressure coefficient vs. azimuth angle are compared to an unsteady 3-D Euler solution, yielding very encouraging results. The code predicts the shape (phase) of the waveform very well, while the magnitude is over-predicted in many cases. At tunnel Mach nos. below 0.6, an unusually large response on the suction surface at 0.15 chord and 0.88 radius was observed. The behavior of this response suggests the presence of a leading edge vortex. The midchord measuring stations on the suction surface exhibit a response that leads the forcing function while most other locations show a phase lag.

Experimental and analytical investigation of fan flow interaction with downstream struts

NASA Technical Reports Server (NTRS)

Olsen, T. L.; Ng, W. F.; Obrien, W. F., Jr.

1985-01-01

An investigation which was designed to provide insight into the fundamental aspects of fan rotor-downstream strut interaction was undertaken. High response, miniature pressure transducers were embedded in the rotor blades of an experimental fan rig. Five downstream struts were placed at several downstream locations in the discharge flow annulus of the single-stage machine. Significant interaction of the rotor blade surface pressures with the flow disturbance produced by the downstream struts was measured. Several numerical procedures for calculating the quasi-steady rotor response due to downstream flow obstructions were developed. A preliminary comparison of experimental and calculated fluctuating blade pressures on the rotor blades shows general agreement between the experimental and calculated values.

Design of Aspirated Compressor Blades Using Three-dimensional Inverse Method

NASA Technical Reports Server (NTRS)

Dang, T. Q.; Rooij, M. Van; Larosiliere, L. M.

2003-01-01

A three-dimensional viscous inverse method is extended to allow blading design with full interaction between the prescribed pressure-loading distribution and a specified transpiration scheme. Transpiration on blade surfaces and endwalls is implemented as inflow/outflow boundary conditions, and the basic modifications to the method are outlined. This paper focuses on a discussion concerning an application of the method to the design and analysis of a supersonic rotor with aspiration. Results show that an optimum combination of pressure-loading tailoring with surface aspiration can lead to a minimization of the amount of sucked flow required for a net performance improvement at design and off-design operations.

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.

Wall shear stress measurement in blade end-wall corner region

NASA Technical Reports Server (NTRS)

Bhargava, R.; Raj, R.; Boldman, D. R.

1987-01-01

The magnitude and the direction of wall shear stress and surface pressure in the blade end-wall corner region were investigated. The measurements were obtained on a specially designed Preston tube, the tip of which could be concentrically rotated about its axis of rotation at the measurement location. The magnitude of wall shear stress in the vicinity of the corner was observed to increase significantly (170 percent) compared to its far-upstream value; the increase was consistently higher on the blade surface compared to the value on the plate surface of the blade end-wall corner. On both surfaces in the blade end-wall corner, the variation of the wall shear stress direction was found to be more predominant in the vicinity of the blade leading-edge location. The trend of the measured wall shear stress direction showed good agreement with the limiting streamline directions obtained from the flow visualization studies.

Study of secondary-flow patterns in an annular cascade of turbine nozzle blades with vortex design

NASA Technical Reports Server (NTRS)

Rohlik, Harold E; Allen, Hubert W; Herzig, Howard Z

1953-01-01

In order to increase understanding of the origin of losses in a turbine, the secondary-flow components in the boundary layers and the blade wakes of an annular cascade of turbine nozzle blades (vortex design) was investigated. A detailed study was made of the total-pressure contours and, particularly, of the inner-wall loss cores downstream of the blades. The inner-wall loss core associated with a blade of the turbine-nozzle cascade is largely the accumulation of low-momentum fluids originating elsewhere in the cascade. This accumulation is effected by a secondary-flow mechanism which acts to transport the low-momentum fluids across the channels on the walls and radially in the blade wakes and boundary layers. The patterns of secondary flow were determined by use of hydrogen sulfide traces, paint, flow fences, and total pressure surveys. At one flow condition investigated, the radial transport of low-momentum fluid in the blade wake and on the suction surface near the trailing edge accounted for 65 percent of the loss core; 30 percent resulted from flow in the thickened boundary layer on the suction surface and 35 percent from flow in the blade wake.

The Influence of Forward Flight on Propeller Noise

NASA Technical Reports Server (NTRS)

Magliozzi, B.

1977-01-01

The effect of flight on blade surface pressures and propeller noise was reported. There were significant differences in blade surface pressures and far-field noise between static and flight conditions. The static data showed many high-intensity, tone-like peaks whereas the flight data was generally free from tones. The turbulence ingested by the propeller operating statically was dominated by long, thin eddies. In flight the scale of the turbulence was greately reduced from that observed statically.

Experimental Pressure Measurements on Hydropower Turbine Runners

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

Harding, Samuel F.; Richmond, Marshall C.

The range of hydrodynamic operating conditions to which the turbine is exposed results in significant pressure fluctuations on both the pressure and suction sides of the blades. Understanding these dynamic pressures has a range of applications. Structurally, the resulting dynamic loads are significant in understanding the design life and maintenance schedule of the bearing, shafts and runner components. The pulsing pressures have also been seen to have a detrimental effect on the surface condition of the blades. Biologically, the pressure gradients and pressure extremes are the primary driver of barotrauma for fish passing through hydroturbines. Improvements in computational fluid dynamicsmore » (CFD) can be used to simulate such unsteady pressures in the regions of concern. High frequency model scale and prototype measurements of pressures at the blade are important in the validation of these models. Experimental characterization of pressure fields over hydroturbine blades has been demonstrated by a number of studies which using multiple pressure transducers to map the pressure contours on the runner blades. These have been performed at both model and prototype scales, often to validate computational models of the pressure and flow fields over the blades. This report provides a review of existing studies in which the blade pressure was measured in situ. The report assesses the technology for both model and prototype scale testing. The details of the primary studies in this field are reported and used to inform the types of hardware required for similar experiments based on the Ice Harbor Dam owned by the US Corps of Engineers on the Snake River, WA, USA. Such a study would be used to validate the CFD performed for the BioPA.« less

Full-field inspection of a wind turbine blade using three-dimensional digital image correlation

NASA Astrophysics Data System (ADS)

LeBlanc, Bruce; Niezrecki, Christopher; Avitabile, Peter; Chen, Julie; Sherwood, James; Hughes, Scott

2011-04-01

Increasing demand and deployment of wind power has led to a significant increase in the number of wind-turbine blades manufactured globally. As the physical size and number of turbines deployed grows, the probability of manufacturing defects being present in composite turbine blade fleets also increases. As both capital blade costs, and operational and maintenance costs, increase for larger turbine systems the need for large-scale inspection and monitoring of the state of structural health of turbine blades during manufacturing and operation critically increase. One method for locating and quantifying manufacturing defects, while also allowing for the in-situ measurement of the structural health of blades, is through the observation of the full-field state of deformation and strain of the blade. Static tests were performed on a nine-meter CX-100 composite turbine blade to extract full-field displacement and strain measurements using threedimensional digital image correlation (3D DIC). Measurements were taken at several angles near the blade root, including along the high-pressure surface, low-pressure surface, and along the trailing edge of the blade. The overall results indicate that the measurement approach can clearly identify failure locations and discontinuities in the blade curvature under load. Post-processing of the data using a stitching technique enables the shape and curvature of the entire blade to be observed for a large-scale wind turbine blade for the first time. The experiment demonstrates the feasibility of the approach and reveals that the technique readily can be scaled up to accommodate utility-scale blades. As long as a trackable pattern is applied to the surface of the blade, measurements can be made in-situ when a blade is on a manufacturing floor, installed in a test fixture, or installed on a rotating turbine. The results demonstrate the great potential of the optical measurement technique and its capability for use in the wind industry for large-area inspection.

Fully three-dimensional and viscous semi-inverse method for axial/radial turbomachine blade design

NASA Astrophysics Data System (ADS)

Ji, Min

2008-10-01

A fully three-dimensional viscous semi-inverse method for the design of turbomachine blades is presented in this work. Built on a time marching Reynolds-Averaged Navier-Stokes solver, the inverse scheme is capable of designing axial/radial turbomachinery blades in flow regimes ranging from very low Mach number to transonic/supersonic flows. In order to solve flow at all-speed conditions, the preconditioning technique is incorporated into the basic JST time-marching scheme. The accuracy of the resulting flow solver is verified with documented experimental data and commercial CFD codes. The level of accuracy of the flow solver exhibited in those verification cases is typical of CFD analysis employed in the design process in industry. The inverse method described in the present work takes pressure loading and blade thickness as prescribed quantities and computes the corresponding three-dimensional blade camber surface. In order to have the option of imposing geometrical constraints on the designed blade shapes, a new inverse algorithm is developed to solve the camber surface at specified spanwise pseudo stream-tubes (i.e. along grid lines), while the blade geometry is constructed through ruling (e.g. straight-line element) at the remaining spanwise stations. The new inverse algorithm involves re-formulating the boundary condition on the blade surfaces as a hybrid inverse/analysis boundary condition, preserving the full three-dimensional nature of the flow. The new design procedure can be interpreted as a fully three-dimensional viscous semi-inverse method. The ruled surface design ensures the blade surface smoothness and mechanical integrity as well as achieves cost reduction for the manufacturing process. A numerical target shooting experiment for a mixed flow impeller shows that the semi-inverse method is able to accurately recover the target blade composed of straightline element from a different initial blade. The semi-inverse method is proved to work well with various loading strategies for the mixed flow impeller. It is demonstrated that uniformity of impeller exit flow and performance gain can be achieved with appropriate loading combinations at hub and shroud. An application of this semi-inverse method is also demonstrated through a redesign of an industrial shrouded subsonic centrifugal impeller. The redesigned impeller shows improved performance and operating range from the original one. Preliminary studies of blade designs presented in this work show that through the choice of the prescribed pressure loading profiles, this semi-inverse method can be used to design blade with the following objectives: (1) Various operating envelope. (2) Uniformity of impeller exit flow. (3) Overall performance improvement. By designing blade geometry with the proposed semi-inverse method whereby the blade pressure loading is specified instead of the conventional design approach of manually adjusting the blade angle to achieve blade design objectives, designers can discover blade geometry design space that has not been explored before.

Active Flow Separation Control of a Stator Vane Using Surface Injection in a Multistage Compressor Experiment

NASA Technical Reports Server (NTRS)

Culley, Dennis E.; Bright, Michelle M.; Prahst, Patricia S.; Strazisar, Anthony J.

2003-01-01

Micro-flow control actuation embedded in a stator vane was used to successfully control separation and improve near stall performance in a multistage compressor rig at NASA Glenn. Using specially designed stator vanes configured with internal actuation to deliver pulsating air through slots along the suction surface, a research study was performed to identify performance benefits using this microflow control approach. Pressure profiles and unsteady pressure measurements along the blade surface and at the shroud provided a dynamic look at the compressor during microflow air injection. These pressure measurements lead to a tracking algorithm to identify the onset of separation. The testing included steady air injection at various slot locations along the vane. The research also examined the benefit of pulsed injection and actively controlled air injection along the stator vane. Two types of actuation schemes were studied, including an embedded actuator for on-blade control. Successful application of an online detection and flow control scheme will be discussed. Testing showed dramatic performance benefit for flow reattachment and subsequent improvement in diffusion through the use of pulsed controlled injection. The paper will discuss the experimental setup, the blade configurations, and preliminary CFD results which guided the slot location along the blade. The paper will also show the pressure profiles and unsteady pressure measurements used to track flow control enhancement, and will conclude with the tracking algorithm for adjusting the control.

Application of the results of experimental and numerical turbulent flow researches based on pressure pulsations analysis

NASA Astrophysics Data System (ADS)

Kovalnogov, Vladislav N.; Fedorov, Ruslan V.; Khakhalev, Yuri A.; Khakhaleva, Larisa V.; Chukalin, Andrei V.

2017-07-01

The numerical investigation of the turbulent flow with the impacts, based on a modified Prandtl mixing-length model with using of the analysis of pulsations of pressure, calculation of structure and a friction factor of a turbulent flow is made. These results under the study allowed us to propose a new design of a cooled turbine blade and gas turbine mobile. The turbine blade comprises a combined cooling and cylindrical cavity on the blade surface, and on the inner surfaces of the cooling channels too damping cavity located on the guide vanes of the compressor of a gas turbine engine, increase the supply of gas-dynamic stability of the compressor of a gas turbine engine, reduce the resistance of the guide blades, and increase the efficiency of the turbine engine.

Prediction of Unsteady Blade Surface Pressures on an Advanced Propeller at an Angle of Attack

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Groeneweg, J. F.

1989-01-01

The numerical solution of the unsteady, three-dimensional, Euler equations is considered in order to obtain the blade surface pressures of an advanced propeller at an angle of attack. The specific configuration considered is the SR7L propeller at cruise conditions with a 4.6 deg inflow angle corresponding to the plus 2 deg nacelle tilt of the Propeller Test Assessment (PTA) flight test condition. The results indicate nearly sinusoidal response of the blade loading, with angle of attack. For the first time, detailed variations of the chordwise loading as a function of azimuthal angle are presented. It is observed that the blade is lightly loaded for part of the revolution and shocks appear from hub to about 80 percent radial station for the highly loaded portion of the revolution.

Prediction of unsteady blade surface pressures on an advanced propeller at an angle of attack

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Groeneweg, J. F.

1989-01-01

The paper considers the numerical solution of the unsteady, three-dimensional, Euler equations to obtain the blade surface pressures of an advanced propeller at an angle of attack. The specific configuration considered is the SR7L propeller at cruise conditions with a 4.6 deg inflow angle corresponding to the +2 deg nacelle tilt of the Propeller Test Assessment (PTA) flight test condition. The results indicate nearly sinusoidal response of the blade loading, with angle of attack. For the first time, detailed variations of the chordwise loading as a function of azimuthal angle are presented. It is observed that the blade is lightly loaded for part of the revolution and shocks appear from hub to about 80 percent radial station for the highly loaded portion of the revolution.

Periodic acoustic radiation from a low aspect ratio propeller

NASA Astrophysics Data System (ADS)

Muench, John David

An experimental program was conducted with the objective of providing high fidelity measurements of propeller inflow, unsteady blade surface pressures, and discrete acoustic radiation over a wide range of speeds. Anechoic wind tunnel experiments were preformed using the SISUP propeller. The upstream stator blades generate large wake deficits that result in periodic unsteady blade forces that acoustically radiate at blade passing frequency and higher harmonics. The experimental portion of this research successfully measured the inflow velocity, blade span unsteady pressures and directive characteristics of the blade-rate radiated noise associated with this complex propeller geometry while the propeller was operating on design. The spatial harmonic decomposition of the inflow revealed significant coefficients at 8, 16 and 24. The magnitude of the unsteady blade forces scale as U4 and linearly shift in frequency with speed. The magnitude of the discrete frequency acoustic levels associated with blade rate scale as U6 and also shift linearly with speed. At blade-rate, the far-field acoustic directivity has a dipole-like directivity oriented perpendicular to the inflow. At the first harmonic of blade-rate, the far-field directivity is not as well defined. The experimental inflow and blade surface pressure results were used to generate an acoustic prediction at blade rate based on a blade strip theory model developed by Blake (1986). The predicted acoustic levels were compared to the experimental results. The model adequately predicts the measured sound field at blade rate at 120 ft/sec. Radiated noise at blade-rate for 120 ft/s can be described by a dipole, whose orientation is perpendicular to the flow and is generated by the interaction of the rotating propeller with the 8th harmonic of the inflow. At blade-rate for 60 ft/s, the model under predicts measured levels. At the first harmonic of blade-rate, for 120 ft/s, the sound field is described as a combination of dipole sources, one generated by the 16 th harmonic, perpendicular to the inflow, and the other generated by the 12th harmonic of the inflow parallel to the inflow. At the first harmonic of blade-rate for 60 ft/s, the model under predicts measured levels.

An unsteady lifting surface method for single rotation propellers

NASA Technical Reports Server (NTRS)

Williams, Marc H.

1990-01-01

The mathematical formulation of a lifting surface method for evaluating the steady and unsteady loads induced on single rotation propellers by blade vibration and inflow distortion is described. The scheme is based on 3-D linearized compressible aerodynamics and presumes that all disturbances are simple harmonic in time. This approximation leads to a direct linear integral relation between the normal velocity on the blade (which is determined from the blade geometry and motion) and the distribution of pressure difference across the blade. This linear relation is discretized by breaking the blade up into subareas (panels) on which the pressure difference is treated as approximately constant, and constraining the normal velocity at one (control) point on each panel. The piece-wise constant loads can then be determined by Gaussian elimination. The resulting blade loads can be used in performance, stability and forced response predictions for the rotor. Mathematical and numerical aspects of the method are examined. A selection of results obtained from the method is presented. The appendices include various details of the derivation that were felt to be secondary to the main development in Section 1.

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.

Three Dimensional Viscous Flow Field in an Axial Flow Turbine Nozzle Passage

NASA Technical Reports Server (NTRS)

Ristic, D.; Lakshminarayana, B.

1997-01-01

The objective of this investigation is experimental and computational study of three dimensional viscous flow field in the nozzle passage of an axial flow turbine stage. The nozzle passage flow field has been measured using a two sensor hot-wire probe at various axial and radial stations. In addition, two component LDV measurements at one axial station (x/c(sum m) = 0.56) were performed to measure the velocity field. Static pressure measurements and flow visualization, using a fluorescent oil technique, were also performed to obtain the location of transition and the endwall limiting streamlines. A three dimensional boundary layer code, with a simple intermittency transition model, was used to predict the viscous layers along the blade and endwall surfaces. The boundary layers on the blade surface were found to be very thin and mostly laminar, except on the suction surface downstream of 70% axial chord. Strong radial pressure gradient, especially close to the suction surface, induces strong cross flow components in the trailing edge regions of the blade. On the end-walls the boundary layers were much thicker, especially near the suction corner of the casing surface, caused by secondary flow. The secondary flow region near the suction-casing surface corner indicates the presence of the passage vortex detached from the blade surface. The corner vortex is found to be very weak. The presence of a closely spaced rotor downstream (20% of the nozzle vane chord) introduces unsteadiness in the blade passage. The measured instantaneous velocity signal was filtered using FFT square window to remove the periodic unsteadiness introduced by the downstream rotor and fans. The filtering decreased the free stream turbulence level from 2.1% to 0.9% but had no influence on the computed turbulence length scale. The computation of the three dimensional boundary layers is found to be accurate on the nozzle passage blade surfaces, away from the end-walls and the secondary flow region. On the nozzle passage endwall surfaces the presence of strong pressure gradients and secondary flow limit the validity of the boundary layer code.

Active control of turbomachine discrete tones

NASA Technical Reports Server (NTRS)

Fleeter, Sanford

1994-01-01

This paper was directed at active control of discrete frequency noise generated by subsonic blade rows through cancellation of the blade row interaction generated propagating acoustic waves. First discrete frequency noise generated by a rotor and stator in a duct was analyzed to determine the propagating acoustic pressure waves. Then a mathematical model was developed to analyze and predict the active control of discrete frequency noise generated by subsonic blade rows through cancellation of the propagating acoustic waves, accomplished by utilizing oscillating airfoil surfaces to generate additional control propagating pressure waves. These control waves interact with the propagating acoustic waves, thereby, in principle, canceling the acoustic waves and thus, the far field discrete frequency tones. This model was then applied to a fan exit guide vane to investigate active airfoil surface techniques for control of the propagating acoustic waves, and thus the far field discrete frequency tones, generated by blade row interactions.

Active control of turbomachine discrete tones

NASA Astrophysics Data System (ADS)

Fleeter, Sanford

This paper was directed at active control of discrete frequency noise generated by subsonic blade rows through cancellation of the blade row interaction generated propagating acoustic waves. First discrete frequency noise generated by a rotor and stator in a duct was analyzed to determine the propagating acoustic pressure waves. Then a mathematical model was developed to analyze and predict the active control of discrete frequency noise generated by subsonic blade rows through cancellation of the propagating acoustic waves, accomplished by utilizing oscillating airfoil surfaces to generate additional control propagating pressure waves. These control waves interact with the propagating acoustic waves, thereby, in principle, canceling the acoustic waves and thus, the far field discrete frequency tones. This model was then applied to a fan exit guide vane to investigate active airfoil surface techniques for control of the propagating acoustic waves, and thus the far field discrete frequency tones, generated by blade row interactions.

Multi-spectral temperature measurement method for gas turbine blade

NASA Astrophysics Data System (ADS)

Gao, Shan; Feng, Chi; Wang, Lixin; Li, Dong

2016-02-01

One of the basic methods to improve both the thermal efficiency and power output of a gas turbine is to increase the firing temperature. However, gas turbine blades are easily damaged in harsh high-temperature and high-pressure environments. Therefore, ensuring that the blade temperature remains within the design limits is very important. There are unsolved problems in blade temperature measurement, relating to the emissivity of the blade surface, influences of the combustion gases, and reflections of radiant energy from the surroundings. In this study, the emissivity of blade surfaces has been measured, with errors reduced by a fitting method, influences of the combustion gases have been calculated for different operational conditions, and a reflection model has been built. An iterative computing method is proposed for calculating blade temperatures, and the experimental results show that this method has high precision.

Experimental and numerical analysis on noise reduction in a multi-blade centrifugal fan

NASA Astrophysics Data System (ADS)

Chen, X. J.; Y Cao, T.; Su, J.; Qin, G. L.

2013-12-01

In this work, analysis on noise source and reduction in a multi-blade centrifugal fan used for air-conditioners was carried out by experimental and numerical methods. Firstly, an experimental system using microphone mounted on volute surface for measuring surface pressure fluctuations of volute was designed and introduced, then surface pressure fluctuations of the whole volute for a multi-blade centrifugal fan were measured by this system, and the inlet noise for this fan was also obtained. And then, based on the experimental results, the aerodynamic noise source of the studied fan was analysed. The surface pressure fluctuations of the volute showed that there were largest surface pressure fluctuations near the volute tongue, and peaks appeared at the Blade Passing Frequency (BPF). The spectra of fan inlet noise showed that the peaks also appeared at BPF, and noise levels in a wide range of frequency were also larger. Secondly, the internal flow of the fan was simulated by commercial software under the same conditions with the experiment, and then the fluid flow and acoustic power field were obtained and discussed. The contours of acoustic power level showed that the larger noise was generated at the impeller area close to the outlet of scroll and at the volute tongue, which is same as that from experiment. Based on all of the results, we can find that the vortex noise is an important part of fan noise for the studied fan, and the rotation noise also cannot be neglected. Finally, several reduction methods that are thought to be effective based on experimental and numerical results were suggested.

Effect of velocity and temperature distribution at the hole exit on film cooling of turbine blades

NASA Technical Reports Server (NTRS)

Garg, Vijay K.; Gaugler, Raymond E.

1995-01-01

An existing three-dimensional Navier-Stokes code, modified to include film cooling considerations, has been used to study the effect of coolant velocity and temperature distribution at the hole exit on the heat transfer coefficient on three-film-cooled turbine blades, namely, the C3X vane, the VKI rotor, and the ACE rotor. Results are also compared with the experimental data for all the blades. Moreover, Mayle's transition criterion, Forest's model for augmentation of leading edge heat transfer due to freestream turbulence, and Crawford's model for augmentation of eddy viscosity due to film cooling are used. Use of Mayle's and Forest's models is relevant only for the ACE rotor due to the absence of showerhead cooling on this rotor. It is found that, in some cases, the effect of distribution of coolant velocity and temperature at the hole exit can be as much as 60% on the heat transfer coefficient at the blade suction surface, and 50% at the pressure surface. Also, different effects are observed on the pressure and suction surface depending upon the blade as well as upon the hole shape, conical or cylindrical.

Laser Anemometer Measurements of the Flow Field in a 4:1 Pressure Ratio Centrifugal Impeller

NASA Technical Reports Server (NTRS)

Skoch, G. J.; Prahst, P. S.; Wernet, M. P.; Wood, J. R.; Strazisar, A. J.

1997-01-01

A laser-doppler anemometer was used to obtain flow-field velocity measurements in a 4:1 pressure ratio, 4.54 kg/s (10 lbm/s), centrifugal impeller, with splitter blades and backsweep, which was configured with a vaneless diffuser. Measured through-flow velocities are reported for ten quasi-orthogonal survey planes at locations ranging from 1% to 99% of main blade chord. Measured through-flow velocities are compared to those predicted by a 3-D viscous steady flow analysis (Dawes) code. The measurements show the development and progression through the impeller and vaneless diffuser of a through-flow velocity deficit which results from the tip clearance flow and accumulation of low momentum fluid centrifuged from the blade and hub surfaces. Flow traces from the CFD analysis show the origin of this deficit which begins to grow in the inlet region of the impeller where it is first detected near the suction surface side of the passage. It then moves toward the pressure side of the channel, due to the movement of tip clearance flow across the impeller passage, where it is cut by the splitter blade leading edge. As blade loading increases toward the rear of the channel the deficit region is driven back toward the suction surface by the cross-passage pressure gradient. There is no evidence of a large wake region that might result from flow separation and the impeller efficiency is relatively high. The flow field in this impeller is quite similar to that documented previously by NASA Lewis in a large low-speed backswept impeller.

Characteristics of Boundary Layer Transition in a Multi-Stage Low-Pressure Turbine

NASA Technical Reports Server (NTRS)

Wisler, Dave; Halstead, David E.; Okiishi, Ted

2007-01-01

An experimental investigation of boundary layer transition in a multi-stage turbine has been completed using surface-mounted hot-film sensors. Tests were carried out using the two-stage Low Speed Research Turbine of the Aerodynamics Research Laboratory of GE Aircraft Engines. Blading in this facility models current, state-of-the-art low pressure turbine configurations. The instrumentation technique involved arrays of densely-packed hot-film sensors on the surfaces of second stage rotor and nozzle blades. The arrays were located at mid-span on both the suction and pressure surfaces. Boundary layer measurements were acquired over a complete range of relevant Reynolds numbers. Data acquisition capabilities provided means for detailed data interrogation in both time and frequency domains. Data indicate that significant regions of laminar and transitional boundary layer flow exist on the rotor and nozzle suction surfaces. Evidence of relaminarization both near the leading edge of the suction surface and along much of the pressure surface was observed. Measurements also reveal the nature of the turbulent bursts occuring within and between the wake segments convecting through the blade row. The complex character of boundary layer transition resulting from flow unsteadiness due to nozzle/nozzle, rotor/nozzle, and nozzle/rotor wake interactions are elucidated using these data. These measurements underscore the need to provide turbomachinery designers with models of boundary layer transition to facilitate accurate prediction of aerodynamic loss and heat transfer.

Measurement of the steady surface pressure distribution on a single rotation large scale advanced prop-fan blade at Mach numbers from 0.03 to 0.78

NASA Technical Reports Server (NTRS)

Bushnell, Peter

1988-01-01

The aerodynamic pressure distribution was determined on a rotating Prop-Fan blade at the S1-MA wind tunnel facility operated by the Office National D'Etudes et de Recherches Aerospatiale (ONERA) in Modane, France. The pressure distributions were measured at thirteen radial stations on a single rotation Large Scale Advanced Prop-Fan (LAP/SR7) blade, for a sequence of operating conditions including inflow Mach numbers ranging from 0.03 to 0.78. Pressure distributions for more than one power coefficient and/or advanced ratio setting were measured for most of the inflow Mach numbers investigated. Due to facility power limitations the Prop-Fan test installation was a two bladed version of the eight design configuration. The power coefficient range investigated was therefore selected to cover typical power loading per blade conditions which occur within the Prop-Fan operating envelope. The experimental results provide an extensive source of information on the aerodynamic behavior of the swept Prop-Fan blade, including details which were elusive to current computational models and do not appear in the two-dimensional airfoil data.

Observed acoustic and aeroelastic spectral responses of a MOD-2 turbine blade to turbulence excitation

NASA Technical Reports Server (NTRS)

Kelley, N. D.; Mckenna, H. E.; Jacobs, E. W.

1995-01-01

Early results from a recent experiment designed to directly evaluate the aeroacoustic/elastic spectral responses of a MOD-2 turbine blade to turbulence-induced unsteady blade loads are discussed. The experimental procedure consisted of flying a hot-film anemometer from a tethered balloon in the turbine inflow and simultaneously measuring the fluctuating airload and aeroelastic response at two blade span stations (65% and 87% spans) using surface-mounted, subminiature pressure transducers and standard strain gage instrumentation. The radiated acoustic pressure field was measured with a triad of very-low-frequency microphones placed at ground level, 1.5 rotor diameters upwind of the disk. Initial transfer function estimates for acoustic radiation, blade normal forces, flapwise acceleration/displacement, and chord/flapwise moments are presented.

Conjugate heat transfer investigation on the cooling performance of air cooled turbine blade with thermal barrier coating

NASA Astrophysics Data System (ADS)

Ji, Yongbin; Ma, Chao; Ge, Bing; Zang, Shusheng

2016-08-01

A hot wind tunnel of annular cascade test rig is established for measuring temperature distribution on a real gas turbine blade surface with infrared camera. Besides, conjugate heat transfer numerical simulation is performed to obtain cooling efficiency distribution on both blade substrate surface and coating surface for comparison. The effect of thermal barrier coating on the overall cooling performance for blades is compared under varied mass flow rate of coolant, and spatial difference is also discussed. Results indicate that the cooling efficiency in the leading edge and trailing edge areas of the blade is the lowest. The cooling performance is not only influenced by the internal cooling structures layout inside the blade but also by the flow condition of the mainstream in the external cascade path. Thermal barrier effects of the coating vary at different regions of the blade surface, where higher internal cooling performance exists, more effective the thermal barrier will be, which means the thermal protection effect of coatings is remarkable in these regions. At the designed mass flow ratio condition, the cooling efficiency on the pressure side varies by 0.13 for the coating surface and substrate surface, while this value is 0.09 on the suction side.

Measuring unsteady pressure on rotating compressor blades

NASA Technical Reports Server (NTRS)

Englund, D. R.; Grant, H. P.; Lanati, G. A.

1979-01-01

Miniature semiconductor strain gage pressure transducers mounted in several arrangements were studied. Both surface mountings and recessed flush mountings were tested. Test parameters included mounting arrangement, blade material, temperature, local strain in the acceleration normal to the transducer diaphragm, centripetal acceleration, and pressure. Test results show no failures of transducers or mountings and indicate an uncertainty of unsteady pressure measurement of approximately + or - 6 percent + 0.1 kPa for a typical application. Two configurations were used on a rotating fan flutter program. Examples of transducer data and correction factors are presented.

Low-Pressure Turbine Separation Control: Comparison With Experimental Data

NASA Technical Reports Server (NTRS)

Garg, Vijay K.

2002-01-01

The present work details a computational study, using the Glenn HT code, that analyzes the use of vortex generator jets (VGJs) to control separation on a low-pressure turbine (LPT) blade at low Reynolds numbers. The computational results are also compared with the experimental data for steady VGJs. It is found that the code determines the proper location of the separation point on the suction surface of the baseline blade (without any VGJ) for Reynolds numbers of 50,000 or less. Also, the code finds that the separated region on the suction surface of the blade vanishes with the use of VGJs. However, the separated region and the wake characteristics are not well predicted. The wake width is generally over-predicted while the wake depth is under-predicted.

Revised FORTRAN program for calculating velocities and streamlines on the hub-shroud midchannel stream surface of an axial-, radial-, or mixed-flow turbomachine or annular duct. 2: Programmer's manual

NASA Technical Reports Server (NTRS)

Katsanis, T.; Mcnally, W. D.

1977-01-01

A FORTRAN IV computer program has been developed that obtains a detailed subsonic or shock free transonic flow solution on the hub-shroud midchannel stream surface of a turbomachine. The blade row may be fixed or rotating, and the blades may be twisted and leaned. Flow may be axial, mixed, or radial. Upstream and downstream flow variables may vary from hub to shroud, and provisions are made to correct for loss of stagnation pressure. The results include velocities, streamlines, and flow angles on the stream surface and approximate blade surface velocities.

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.

Prediction and analysis of infra and low-frequency noise of upwind horizontal axis wind turbine using statistical wind speed model

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

Lee, Gwang-Se; Cheong, Cheolung, E-mail: ccheong@pusan.ac.kr

Despite increasing concern about low-frequency noise of modern large horizontal-axis wind turbines (HAWTs), few studies have focused on its origin or its prediction methods. In this paper, infra- and low-frequency (the ILF) wind turbine noise are closely examined and an efficient method is developed for its prediction. Although most previous studies have assumed that the ILF noise consists primarily of blade passing frequency (BPF) noise components, these tonal noise components are seldom identified in the measured noise spectrum, except for the case of downwind wind turbines. In reality, since modern HAWTs are very large, during rotation, a single blade ofmore » the turbine experiences inflow with variation in wind speed in time as well as in space, breaking periodic perturbations of the BPF. Consequently, this transforms acoustic contributions at the BPF harmonics into broadband noise components. In this study, the ILF noise of wind turbines is predicted by combining Lowson’s acoustic analogy with the stochastic wind model, which is employed to reproduce realistic wind speed conditions. In order to predict the effects of these wind conditions on pressure variation on the blade surface, unsteadiness in the incident wind speed is incorporated into the XFOIL code by varying incident flow velocities on each blade section, which depend on the azimuthal locations of the rotating blade. The calculated surface pressure distribution is subsequently used to predict acoustic pressure at an observing location by using Lowson’s analogy. These predictions are compared with measured data, which ensures that the present method can reproduce the broadband characteristics of the measured low-frequency noise spectrum. Further investigations are carried out to characterize the IFL noise in terms of pressure loading on blade surface, narrow-band noise spectrum and noise maps around the turbine.« less

Effect of Velocity and Temperature Distribution at the Hole Exit on Film Cooling of Turbine Blades

NASA Technical Reports Server (NTRS)

Garg, V. K.; Gaugler, R. E.

1997-01-01

An existing three-dimensional Navier-Stokes code (Arnone et al, 1991), modified Turbine Branch, to include film cooling considerations (Garg and Gaugler, 1994), has been used to study the effect of coolant velocity and temperature distribution at the hole exit on the heat transfer coefficient on three film-cooled turbine blades, namely, the C3X vane, the VKI rotor, and the ACE rotor. Results are also compared with the experimental data for all the blades. Moreover, Mayle's transition criterion (1991), Forest's model for augmentation of leading edge heat transfer due to free-stream turbulence (1977), and Crawford's model for augmentation of eddy viscosity due to film cooling (Crawford et al, 1980) are used. Use of Mayle's and Forest's models is relevant only for the ACE rotor due to the absence of showerhead cooling on this rotor. It is found that, in some cases, the effect of distribution of coolant velocity and temperature at the hole exit can be as much as 60 percent on the heat transfer coefficient at the blade suction surface, and 50 percent at the pressure surface. Also, different effects are observed on the pressure and suction surface depending upon the blade as well as upon the hole shape, conical or cylindrical.

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.

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

Endwall flows and blading design for axial flow compressors

NASA Astrophysics Data System (ADS)

Robinson, Christopher J.

Literature relevant to blading design in the endwall region is reviewed, and important three dimensional flow phenomena occurring in embedded stages of axial compressors are described. A low speed axial flow four stage compressor rig is described and bladings studied are detailed: two conventional and two with end bends. The application of a three dimensional Navier-Stokes solver to the bladings' stators, to assess the effectiveness of the code, is reported. Calculation results of exit whirl angles, losses, and surface static pressures are compared with experiment.

Unsteady Flowfield in a High-Pressure Turbine Modeled by TURBO

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.; Mehmed, Oral

2003-01-01

Forced response, or resonant vibrations, in turbomachinery components can cause blades to crack or fail because of the large vibratory blade stresses and subsequent high-cycle fatigue. Forced-response vibrations occur when turbomachinery blades are subjected to periodic excitation at a frequency close to their natural frequency. Rotor blades in a turbine are constantly subjected to periodic excitations when they pass through the spatially nonuniform flowfield created by upstream vanes. Accurate numerical prediction of the unsteady aerodynamics phenomena that cause forced-response vibrations can lead to an improved understanding of the problem and offer potential approaches to reduce or eliminate specific forced-response problems. The objective of the current work was to validate an unsteady aerodynamics code (named TURBO) for the modeling of the unsteady blade row interactions that can cause forced response vibrations. The three-dimensional, unsteady, multi-blade-row, Reynolds-averaged Navier-Stokes turbomachinery code named TURBO was used to model a high-pressure turbine stage for which benchmark data were recently acquired under a NASA contract by researchers at the Ohio State University. The test article was an initial design for a high-pressure turbine stage that experienced forced-response vibrations which were eliminated by increasing the axial gap. The data, acquired in a short duration or shock tunnel test facility, included unsteady blade surface pressures and vibratory strains.

Unsteady Force Calculations in Turbomachinery

DTIC Science & Technology

1991-07-01

Engineering for Gas Turbines and Power, Vol. 107, pp. 945-952, October 1985. Lefcort, M. P., "An Investigation into Unsteady Blade Forces in...generated unsteady flow around a rotating turbine blade row .. ..... 43 7 The rotating coordinate system with skew, 0, and rake, zr, defined at midchord...while Kerrebrock and Mikolajczak [19701 5 proved it experimentally. For a turbine blade passage, the wake fluid moves from the pressure 3 surface to the

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.

A case study of the fluid structure interaction of a Francis turbine

NASA Astrophysics Data System (ADS)

Müller, C.; Staubli, T.; Baumann, R.; Casartelli, E.

2014-03-01

The Francis turbine runners of the Grimsel 2 pump storage power plant showed repeatedly cracks during the last decade. It is assumed that these cracks were caused by flow induced forces acting on blades and eventual resonant runner vibrations lead to high stresses in the blade root areas. The eigenfrequencies of the runner were simulated in water using acoustic elements and compared to experimental data. Unsteady blades pressure distribution determined by a transient CFD simulation of the turbine were coupled to a FEM simulation. The FEM simulation enabled analyzing the stresses in the runner and the eigenmodes of the runner vibrations. For a part-load operating point, transient CFD simulations of the entire turbine, including the spiral case, the runner and the draft tube were carried out. The most significant loads on the turbine runner resulted from the centrifugal forces and the fluid forces. Such forces effect temporally invariant runner blades loads, in contrast rotor stator interaction or draft tube instabilities induce pressure fluctuations which cause the temporally variable forces. The blades pressure distribution resulting from the flow simulation was coupled by unidirectional-harmonic FEM simulation. The dominant transient blade pressure distribution of the CFD simulation were Fourier transformed, and the static and harmonic portion assigned to the blade surfaces in the FEM model. The evaluation of the FEM simulation showed that the simulated part load operating point do not cause critical stress peaks in the crack zones. The pressure amplitudes and frequencies are very small and interact only locally with the runner blades. As the frequencies are far below the modal frequencies of the turbine runner, resonant vibrations obviously are not excited.

Gas turbine blade film cooling and blade tip heat transfer

NASA Astrophysics Data System (ADS)

Teng, Shuye

The detailed heat transfer coefficient and film cooling effectiveness distributions as well as the detailed coolant jet temperature profiles on the suction side of a gas turbine blade were measured using a transient liquid crystal image method and a traversing cold wire and thermocouple probe, respectively. The blade has only one row of film holes near the gill hole portion on the suction side of the blade. The hole geometries studied include standard cylindrical holes and holes with diffuser shaped exit portion (i.e. fanshaped holes and laidback fanshaped holes). Tests were performed on a five-blade linear cascade in a low-speed wind tunnel. The mainstream Reynolds number based on cascade exit velocity was 5.3 x 105. The upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 and 0.1. The coolant blowing ratio was varied from 0.4 to 1.2. Results show that both expanded holes have significantly improved thermal protection over the surface downstream of the ejection location, particularly at high blowing ratios. However, the expanded hole injections induce earlier boundary layer transition to turbulence and enhance heat transfer coefficients at the latter part of the blade suction surface. In general, the unsteady wake tends to reduce film cooling effectiveness. Measurements of detailed heat transfer coefficient distributions on a turbine blade tip were performed in the same wind tunnel facility as above. The central blade had a variable tip gap clearance. Measurements were made at three different tip gap clearances of about 1.1%, 2.1%, and 3% of the blade span. Static pressure distributions were measured in the blade mid-span and on the shroud surface. Detailed heat transfer coefficient distributions were measured on the blade tip surface. Results show that reduced tip clearance leads to reduced heat transfer coefficient over the blade tip surface. Results also show that reduced tip clearance tends to weaken the unsteady wake effect on blade tip heat transfer.

Experimental Study of the Effects of Periodic Unsteady Wakes on Flow Separation in Low Pressure Turbines

NASA Technical Reports Server (NTRS)

Ozturk, Burak; Schobeiri, Meinhard T.

2009-01-01

The present study, which is the first of a series of investigations of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary layer flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed on a large-scale, high-subsonic unsteady turbine cascade research facility with an integrated wake generator and test section unit. Blade Pak B geometry was used in the cascade. The wakes were generated by continuously moving cylindrical bars device. Boundary layer investigations were performed using hot wire anemometry at Reynolds number of 110,000, based on the blade suction surface length and the exit velocity, for one steady and two unsteady inlet flow conditions, with the corresponding passing frequencies, wake velocities, and turbulence intensities. The reduced frequencies cover the entire operation range of LP-turbines. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re = 50,000, 75,000, 100,000, 110,000, and 125,000. For each Reynolds number, surface pressure measurements are carried out at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extension of the separation zone as well as its behavior under unsteady wake flow. The results, presented in ensemble-averaged and contour plot forms, help to understand the physics of the separation phenomenon under periodic unsteady wake flow.

Wind-Tunnel Tests of a Portion of a PV-2 Helicopter Rotor Blade

NASA Technical Reports Server (NTRS)

Kemp, William B., Jr.

1945-01-01

A portion of a PV-2 helicopter rotor blade has been tested in the 6- by 6-foot test section of the Langley stability tunnel to determine if the aerodynamic characteristics were seriously affected by cross flow or fabric distortion. The outer portion of the blade was tested as a reflection plane model pivoted about the tunnel wall to obtain various angles of cross flow over the blade. Because the tunnel wall acts as a plane of sytry, the measured aerodynamic characteristics correspond to those of an airfoil having various angles of sweepforward and sweepback. Tests were made with the vents on the lower surface open and also with the vents sealed and the internal pressure held at -20 inches of water producing an internal pressure coefficient of -1.059. The change in contour resulting from the range of internal pressures used had very little effect on the aerodynamic characteristics of the blade. The test methods were considered to simulate inadequately the flow conditions over the rotor blade because the effects of cross flow were limited to conditions corresponding to sweep of the blade. The results indicated that this type of cross flow had only minor effects on the aerodynamic characteristics of the blade. It is believed, therefore, that future tests to determine the effects on the aerodynamic characteristics of cross flow should utilize complete rotors.

Minnowbrook IV: 2003 Workshop on Transition and Unsteady Aspects of Turbomachinery Flows

NASA Technical Reports Server (NTRS)

LaGraff, John E. (Editor); Ashpis, David E.

2004-01-01

This Minnowbrook IV 2003 workshop on Transition and Unsteady Aspects of Turbomachinery Flows includes the following topics: 1) Current Issues in Unsteady Turbomachinery Flows; 2) Global Instability and Control of Low-Pressure Turbine Flows; 3) Influence of End Wall Leakage on Secondary Flow Development in Axial Turbines; 4) Active and Passive Flow Control on Low Pressure Turbine Airfoils; 5) Experimental and Numerical Investigation of Transitional Flows as Affected by Passing Wakes; 6) Effects of Freestream Turbulence on Turbine Blade Heat Transfer; 7) Bypass Transition Via Continuous Modes and Unsteady Effects on Film Cooling; 8) High Frequency Surface Heat Flux Imaging of Bypass Transition; 9) Skin Friction and Heat Flux Oscillations in Upstream Moving Wave Packets; 10) Transition Mechanisms and Use of Surface Roughness to Enhance the Benefits of Wake Passing in LP Turbines; 11) Transient Growth Approach to Roughness-Induced Transition; 12) Roughness- and Freestream-Turbulence-Induced Transient Growth as a Bypass Transition Mechanism; 13) Receptivity Calculations as a Means to Predicting Transition; 14) On Streamwise Vortices in a Curved Wall Jet and Their Effect on the Mean Flow; 15) Plasma Actuators for Separation Control of Low Pressure Turbine Blades; 16) Boundary-Layer Separation Control Under Low-Pressure-Turbine Conditions Using Glow-Discharge Plasma Actuators; 17) Control of Separation for Low Pressure Turbine Blades: Numerical Simulation; 18) Effects of Elevated Free-Stream Turbulence on Active Control of a Separation Bubble; 19) Wakes, Calming and Transition Under Strong Adverse Pressure Gradients; 20) Transitional Bubble in Periodic Flow Phase Shift; 21) Modelling Spots: The Calmed Region, Pressure Gradient Effects and Background; 22) Modeling of Unsteady Transitional Flow on Axial Compressor Blades; 23) Challenges in Predicting Component Efficiencies in Turbomachines With Low Reynolds Number Blading; 24) Observations on the Causal Relationship Between Blade Count and Developing Rotating Stall in a Four Stage Axial Compressor; 25) Experimental and Numerical Study of Non-Linear Interactions in Transonic Nozzle Flow; 26) Clocking Effects on a Modern Stage and One-Half Transonic Turbine; 27) DNS and LES of Transition on Turbine Blades; 28) The Use of Cellular Automata in Modeling the Transition; 29) Predicting Unsteady Buffet Onset Using RANS Solutions; 30) Transition Modelling With the SST Turbulence Model and an Intermittency Transport; and 31) Equation Workshop Summary Transcript

Research of performance prediction to energy on hydraulic turbine

NASA Astrophysics Data System (ADS)

Quan, H.; Li, R. N.; Li, Q. F.; Han, W.; Su, Q. M.

2012-11-01

Refer to the low specific speed Francis turbine blade design principle and double-suction pump structure. Then, design a horizontal double-channel hydraulic turbine Francis. Through adding different guide vane airfoil and and no guide vane airfoil on the hydraulic conductivity components to predict hydraulic turbine energy and using Fluent software to numerical simulation that the operating conditions and point. The results show that the blade pressure surface and suction surface pressure is low when the hydraulic turbine installation is added standard positive curvature of the guide vane and modified positive curvature of guide vane. Therefore, the efficiency of energy recovery is low. However, the pressure of negative curvature guide vane and symmetric guide vane added on hydraulic turbine installations is larger than that of the former ones, and it is conducive to working of runner. With the decreasing of guide vane opening, increasing of inlet angle, flow state gets significantly worse. Then, others obvious phenomena are that the reflux and horizontal flow appeared in blade pressure surface. At the same time, the vortex was formed in Leaf Road, leading to the loss of energy. Through analyzing the distribution of pressure, velocity, flow lines of over-current flow in the the back hydraulic conductivity components in above programs we can known that the hydraulic turbine installation added guide vane is more reasonable than without guide vanes, it is conducive to improve efficiency of energy conversion.

Compressor cascade performance deterioration caused by sand ingestion

NASA Technical Reports Server (NTRS)

Tabakoff, W.; Balan, C.

1982-01-01

Airfoil cascade erosion and performance deterioration was investigated in a gas particle cascade tunnel. The cascade blades were made of 2024 aluminum alloy and the solid particles used were quartz sand. The results of the experimental measurements are presented to show the change in the blade surface erosion, pressure distribution and the total loss coefficient with erosion. The surface quality of the blades exposed to particulate flows are changing the material surfaces. With time, the surface roughness increases and leads to a decrease in engine performance. It was found that the surface roughness values increase asymptotically to a maximum value with increased erosion. The experimental results indicate that the roughness parameters correlate well against the mass of particles impacting unit area of the surface. Such a correlation is useful in aerodynamics and performance computations in turbomachinery.

Jet Flap Stator Blade Test in the High Reaction Turbine Blade Cascade Tunnel

NASA Image and Video Library

1970-03-21

A researcher examines the setup of a jet flap blade in the High Reaction Turbine Blade Cascade Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis researchers were seeking ways to increase turbine blade loading on aircraft engines in an effort to reduce the overall size and weight of engines. The ability of each blade to handle higher loads meant that fewer stages and fewer blades were required. This study analyzed the performance of a turbine blade using a jet flap and high loading. A jet of air was injected into the main stream from the pressure surface near the trailing edge. The jet formed an aerodynamic flap which deflected the flow and changed the circulation around the blade and thus increased the blade loading. The air jet also reduced boundary layer thickness. The jet-flap blade design was appealing because the cooling air may also be used for the jet. The performance was studied in a two-dimensional cascade including six blades. The researcher is checking the jet flat cascade with an exit survey probe. The probe measured the differential pressure that was proportional to the flow angle. The blades were tested over a range of velocity ratios and three jet flow conditions. Increased jet flow improved the turning and decreased both the weight flow and the blade loading. However, high blade loadings were obtained at all jet flow conditions.

Aero-thermal optimization of film cooling flow parameters on the suction surface of a high pressure turbine blade

NASA Astrophysics Data System (ADS)

El Ayoubi, Carole; Hassan, Ibrahim; Ghaly, Wahid

2012-11-01

This paper aims to optimize film coolant flow parameters on the suction surface of a high-pressure gas turbine blade in order to obtain an optimum compromise between a superior cooling performance and a minimum aerodynamic penalty. An optimization algorithm coupled with three-dimensional Reynolds-averaged Navier Stokes analysis is used to determine the optimum film cooling configuration. The VKI blade with two staggered rows of axially oriented, conically flared, film cooling holes on its suction surface is considered. Two design variables are selected; the coolant to mainstream temperature ratio and total pressure ratio. The optimization objective consists of maximizing the spatially averaged film cooling effectiveness and minimizing the aerodynamic penalty produced by film cooling. The effect of varying the coolant flow parameters on the film cooling effectiveness and the aerodynamic loss is analyzed using an optimization method and three dimensional steady CFD simulations. The optimization process consists of a genetic algorithm and a response surface approximation of the artificial neural network type to provide low-fidelity predictions of the objective function. The CFD simulations are performed using the commercial software CFX. The numerical predictions of the aero-thermal performance is validated against a well-established experimental database.

A Combined Experimental/Computational Study of Flow in Turbine Blade Cooling Passage

NASA Technical Reports Server (NTRS)

Tse, D. G. N.; Kreskovsky, J. P.; Shamroth, S. J.; Mcgrath, D. B.

1994-01-01

Laser velocimetry was utilized to map the velocity field in a serpentine turbine blade cooling passage at Reynolds and Rotation numbers of up to 25.000 and 0.48. These results were used to assess the combined influence of passage curvature and Coriolis force on the secondary velocity field generated. A Navier-Stokes code (NASTAR) was validated against incompressible test data and then used to simulate the effect of buoyancy. The measurements show a net convection from the low pressure surface to high pressure surface. The interaction of the secondary flows induced by the turns and rotation produces swirl at the turns, which persisted beyond 2 hydraulic diameters downstream of the turns. The incompressible flow field predictions agree well with the measured velocities. With radially outward flow, the buoyancy force causes a further increase in velocity on the high pressure surface and a reduction on the low pressure surface. The results were analyzed in relation to the heat transfer measurements of Wagner et al. (1991). Predicted heat transfer is enhanced on the high pressure surfaces and in turns. The incompressible flow simulation underpredicts heat transfer in these locations. Improvements observed in compressible flow simulation indicate that the buoyancy force may be important.

An analysis of pump cavitation damage. [Space Shuttle main engine high pressure oxidizer turbopump

NASA Technical Reports Server (NTRS)

Brophy, M. C.; Stinebring, D. R.; Billet, M. L.

1985-01-01

The cavitation assessment for the space shuttle main engine high pressure oxidizer turbopump is documented. A model of the flow through the pump was developed. Initially, a computational procedure was used to analyze the flow through the inlet casing including the prediction of wakes downstream of the casing vanes. From these flow calculations, cavitation patterns on the inducer blades were approximated and the damage rate estimated. The model correlates the heavy damage on the housing and over the inducer with unsteady blade surface cavitation. The unsteady blade surface cavitation is due to the large incidence changes caused by the wakes of the upstream vanes. Very high cavitation damage rates are associated with this type of cavitation. Design recommendations for reducing the unsteady cavitation include removing the set of vanes closest to the inducer and modifying the remaining vanes.

A study of pump cavitation damage. [space shuttle main engine high pressure oxidizer turbopump

NASA Technical Reports Server (NTRS)

Brophy, M. C.; Stinebring, D. R.; Billet, M. L.

1983-01-01

The cavitation assessment for the space shuttle main engine high pressure oxidizer turbopump is documented. A model of the flow through the pump was developed. Initially, a computational procedure was used to analyze the flow through the inlet casing including the prediction of wakes downstream of the casing vanes. From these flow calculations, cavitation patterns on the inducer blades were approximated and the damage rate estimated. The model correlates the heavy damage on the housing and over the inducer with unsteady blade surface cavitation. The unsteady blade surface cavitation is due to the large incidence changes caused by the wakes of the upstream vanes. Very high cavitation damage rates are associated with this type of cavitation. Design recommendations for reducing the unsteady cavitation include removing the set of vanes closest to the inducer and modifying the remaining vanes.

Numerical investigation on pressure fluctuations in centrifugal compressor with different inlet guide vanes pre-whirl angles

NASA Astrophysics Data System (ADS)

Wang, Y. C.; Shi, M.; Cao, S. L.; Li, Z. H.

2013-12-01

The pressure fluctuations in a centrifugal compressor with different inlet guide vanes (IGV) pre-whirl angles were investigated numerically, as well as the pre-stress model and static structural of blade. The natural frequency was evaluated by pre-stress model analysis. The results show that, the aero-dynamic pressure acting on blade surface is smaller than rotation pre-stress, which wouldn't result in large deformation of blade. The natural frequencies with rotation pre-stress are slightly higher than without rotation pre-stress. The leading mechanism of pressure fluctuations for normal conditions is the rotor-stator (IGVs) interaction, while is serious flow separations for conditions that are close to surge line. A few frequency components in spectra are close to natural frequency, which possibly result in resonant vibration if amplitude is large enough, which is dangerous for compressor working, and should be avoided.

Large-scale Advanced Prop-fan (LAP) high speed wind tunnel test report

NASA Technical Reports Server (NTRS)

Campbell, William A.; Wainauski, Harold S.; Arseneaux, Peter J.

1988-01-01

High Speed Wind Tunnel testing of the SR-7L Large Scale Advanced Prop-Fan (LAP) is reported. The LAP is a 2.74 meter (9.0 ft) diameter, 8-bladed tractor type rated for 4475 KW (6000 SHP) at 1698 rpm. It was designated and built by Hamilton Standard under contract to the NASA Lewis Research Center. The LAP employs thin swept blades to provide efficient propulsion at flight speeds up to Mach .85. Testing was conducted in the ONERA S1-MA Atmospheric Wind Tunnel in Modane, France. The test objectives were to confirm that the LAP is free from high speed classical flutter, determine the structural and aerodynamic response to angular inflow, measure blade surface pressures (static and dynamic) and evaluate the aerodynamic performance at various blade angles, rotational speeds and Mach numbers. The measured structural and aerodynamic performance of the LAP correlated well with analytical predictions thereby providing confidence in the computer prediction codes used for the design. There were no signs of classical flutter throughout all phases of the test up to and including the 0.84 maximum Mach number achieved. Steady and unsteady blade surface pressures were successfully measured for a wide range of Mach numbers, inflow angles, rotational speeds and blade angles. No barriers were discovered that would prevent proceeding with the PTA (Prop-Fan Test Assessment) Flight Test Program scheduled for early 1987.

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).

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.

Calculation and Correlation of the Unsteady Flowfield in a High Pressure Turbine

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.; Liu, Jong S.; Panovsky, Josef; Keith, Theo G., Jr.; Mehmed, Oral

2002-01-01

Forced vibrations in turbomachinery components can cause blades to crack or fail due to high-cycle fatigue. Such forced response problems will become more pronounced in newer engines with higher pressure ratios and smaller axial gap between blade rows. An accurate numerical prediction of the unsteady aerodynamics phenomena that cause resonant forced vibrations is increasingly important to designers. Validation of the computational fluid dynamics (CFD) codes used to model the unsteady aerodynamic excitations is necessary before these codes can be used with confidence. Recently published benchmark data, including unsteady pressures and vibratory strains, for a high-pressure turbine stage makes such code validation possible. In the present work, a three dimensional, unsteady, multi blade-row, Reynolds-Averaged Navier Stokes code is applied to a turbine stage that was recently tested in a short duration test facility. Two configurations with three operating conditions corresponding to modes 2, 3, and 4 crossings on the Campbell diagram are analyzed. Unsteady pressures on the rotor surface are compared with data.

Axial-Flow Turbine Rotor Discharge-Flow Overexpansion and Limit-Loading Condition, Part I: Computational Fluid Dynamics (CFD) Investigation

NASA Technical Reports Server (NTRS)

Chen, Shu-Cheng S.

2017-01-01

A Computational Fluid Dynamic (CFD) investigation is conducted over a two-dimensional axial-flow turbine rotor blade row to study the phenomena of turbine rotor discharge flow overexpansion at subcritical, critical, and supercritical conditions. Quantitative data of the mean-flow Mach numbers, mean-flow angles, the tangential blade pressure forces, the mean-flow mass flux, and the flow-path total pressure loss coefficients, averaged or integrated across the two-dimensional computational domain encompassing two blade-passages, are obtained over a series of 14 inlet-total to exit-static pressure ratios, from 1.5 (un-choked; subcritical condition) to 10.0 (supercritical with excessively high pressure ratio.) Detailed flow features over the full domain-of-computation, such as the streamline patterns, Mach contours, pressure contours, blade surface pressure distributions, etc. are collected and displayed in this paper. A formal, quantitative definition of the limit loading condition based on the channel flow theory is proposed and explained. Contrary to the comments made in the historical works performed on this subject, about the deficiency of the theoretical methods applied in analyzing this phenomena, using modern CFD method for the study of this subject appears to be quite adequate and successful. This paper describes the CFD work and its findings.

Structural and aerodynamic loads and performance measurements of an SA349/2 helicopter with an advanced geometry rotor

NASA Technical Reports Server (NTRS)

Heffernan, Ruth M.; Gaubert, Michel

1986-01-01

A flight test program was conducted to obtain data from an upgraded Gazelle helicopter with an advanced geometry, three bladed rotor. Data were acquired on upper and lower surface chordwise blade pressure, blade bending and torsion moments, and fuselage structural loads. Results are presented from 16 individual flight conditions, including level flights ranging from 10 to 77 m/sec at 50 to 3000 m altitude, turning flights up to 2.0 g, and autorotation. Rotor aerodynamic data include information from 51 pressure transducers distributed chordwise at 75, 88, and 97% radial stations. Individual tranducer pressure coefficients and airfoil section lift and pitching moment coefficients are presented, as are steady state flight condition parameters and time dependence rotor loads. All dynamic data are presented as harmonic analysis coefficients.

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).

Energy Efficient High-Pressure Turbine Leakage Technology Report

NASA Technical Reports Server (NTRS)

Gardner, W. B.

1980-01-01

The leakage test program was one of such supporting technology programs structured to provide guidance to the Energy Efficient Engine High Pressure Turbine Component Design Effort. Leakage reduction techniques were identified and evaluated. Test models were used to simulate component leak paths and to evaluate leakage reduction techniques. These models simulated the blade/disk attachment, the vane inner platform attachment, and the vane outer platform attachment combined with the blade outer airseal. Disk blade attachment testing indicated that leakage in this area could be reduced to very low levels by paying careful attention to the tolerances along the contact surface between the blade vibration damper and the blade platform contact surface. The aim of feather seal testing was to achieve a goal for an effective leakage gap of one mil (.001 inch) per inch of feather seal length. Results indicated that effective gaps even below the goal level were achievable by (1) maintaining close tolerances between feather seals and their slots to minimize end gaps and limit seal rotation, (2) avoiding feather seal overlap, and (3) minimizing feather seal intersections. W seals were shown to be effective leakage control devices. Wire rope, in its present state of development, was shown not to be an effective sealing concept for application to the component design.

Potential disturbance interactions with a single IGV in an F109 turbofan engine

NASA Astrophysics Data System (ADS)

Kirk, Joel F.

A common cause of aircraft engine failure is the high cycle fatigue of engine blades and stators. One of the primary causes of these failures is due to blade row interactions, which cause an aerodynamic excitation to be resonant with a mechanical natural frequency. Traditionally, the primary source of such aerodynamic excitations has been practically limited to viscous wakes from upstream components. However, more advanced designs require that blade rows be very highly loaded and closely spaced. This results in aerodynamic excitation from potential fields of down stream engine components, in addition to the known wake excitations. An experimental investigation of the potential field from the fan of a Honeywell F109 turbofan engine has been completed. The investigation included velocity measurements upstream of the fan, addition of an airfoil shaped probe upstream of the fan on which surface pressure measurements were acquired, and measurement of the velocity in the interaction region between the probe and the fan. This investigation sought to characterize the response on the upstream probe due to the fan potential field and the interaction between a viscous wake and the potential field; as such, all test conditions were for subsonic fan speeds. The results from the collected data show that fan-induced potential disturbances propagate upstream at acoustic velocities, to produce vane surface-pressure amplitudes as high as 40 percent Joel F. Kirk of the inlet, mean total pressure. Further, these fan-induced pressure amplitudes display large variations between the two vane surfaces. An argument is made that the structure of the pressure response is consistent with the presence of two distinct sources of unsteady forcing disturbances. The disturbances on the incoming-rotation-facing surface of the IGV propagated upstream at a different speed than those on the outgoing-rotation-facing surface, indicating that one originated from a rotating source and the other from a stationary source. An argument is made to suggest that the stationary source is due to the rotor blades cutting through the wake of the IGV.

Heat transfer measurements on an incidence-tolerant low pressure turbine blade in a high speed linear cascade at low to moderate Reynolds numbers

NASA Astrophysics Data System (ADS)

Moualeu, Leolein Patrick Gouemeni

Runway-independent aircraft are expected to be the future for short-haul flights by improving air transportation and reducing area congestion encountered in airports. The Vehicle Systems Program of NASA identified a Large Civil Tilt-Rotor, equipped with variable-speed power-turbine engines, as the best concept. At cruise altitude, the engine rotor-speed will be reduced by as much as the 50% of take-off speed. The large incidence variation in the low pressure turbine associated with the change in speed can be detrimental to the engine performance. Low pressure turbine blades in cruise altitude are more predisposed to develop regions of boundary layer separation. Typical phenomenon such as impinging wakes on downstream blades and mainstream turbulences enhance the complexity of the flow in low pressure turbines. It is therefore important to be able to understand the flow behavior to accurately predict the losses. Research facilities are seldom able to experimentally reproduce low Reynolds numbers at relevant engine Mach number. Having large incidence swing as an additional parameter in the investigation of the boundary layer development, on a low pressure turbine blade, makes this topic unique and as a consequence requires a unique facility to conduct the experimental research. The compressible flow wind tunnel facility at the University of North Dakota had been updated to perform steady state experiments on a modular-cascade, designed to replicate a large variation of the incidence angles. The high speed and low Reynolds number facility maintained a sealed and closed loop configuration for each incidence angle. The updated facility is capable to produce experimental Reynolds numbers as low as 45,000 and as high as 570,000 at an exit Mach number of 0.72. Pressure and surface temperature measurements were performed at these low pressure turbine conditions. The present thesis investigates the boundary layer development on the surface of an Incidence-tolerant blade. The heat transfer approach is the method used to obtain knowledge of the state of the boundary layer on the surface of the blade. Pressure and temperature distributions are acquired for Reynolds numbers of 50,000, 66,000, 228,000, and 568,000 at an exit Mach number of 0.72, and Reynolds numbers of 228,000, and 568,000 at an exit Mach number of 0.35. These experimental flow conditions are conducted at different flow inlet angles of 40°, 34.2°, 28°, 18°, 8°, -2.6°, -12°, and -17°, and at two free-stream turbulence levels. Results of the analyses performed show that as the incidence angle decreases, a region of laminar separation bubble forms on the pressure surface and grows toward the trailing-edge. It is also noted that the position of the leading-edge moves as the incidence angle varies. A transitional flow is observed on both the pressure and suction surfaces, mainly at the two highest incidence angles, for the high turbulence case. This investigation also reveals that the Stanton number increases as the mainstream turbulence increases, and that the Stanton number at the leading-edge increases as the Reynolds number decreases, as it is documented in the literature.

Water droplet erosion of stainless steel steam turbine blades

NASA Astrophysics Data System (ADS)

Kirols, H. S.; Kevorkov, D.; Uihlein, A.; Medraj, M.

2017-08-01

Steam turbine blades are highly subjected to water droplet erosion (WDE) caused by high energy impingement of liquid water droplets. However, most of the published research on this wear phenomenon is performed on laboratory test rigs, instead of addressing WDE of actual steam turbine blades. In this work, the progression of erosion on the surface of ex-service low pressure steam turbine blades was investigated using scanning electron microscopy. The erosion appearance and mechanisms are compared with laboratory test rig results that are carried out using a rotating disk rig according to ASTM G73 standard. Initial and advanced erosion stages could be observed on the steam turbine blades. Similar to the WDE rig coupons, initial pits and cracks were preceded by blade surface roughening through the formation of asperities and depressions. In addition, it was also observed that the twist angle of the turbine blade around its diagonal, is an important parameter that influences its WDE. Twist angle has an effect on: impact angle, erosion appearance, impact speed, and the affected area. Furthermore, according to the current experimental results, multi-ray rig erosion test results are considered the closest simulation to the actual ex-service blade in terms of damage appearance.

Effects of Freestream Turbulence on Turbine Blade Heat Transfer

NASA Technical Reports Server (NTRS)

Boyle, Robert J.; Giel, Paul W.; Ames, Forrest E.

2004-01-01

Experiments have shown that moderate turbulence levels can nearly double turbine blade stagnation region heat transfer. Data have also shown that heat transfer is strongly affected by the scale of turbulence as well as its level. In addition to the stagnation region, turbulence is often seen to increase pressure surface heat transfer. This is especially evident at low to moderate Reynolds numbers. Vane and rotor stagnation region, and vane pressure surface heat transfer augmentation is often seen in a pre-transition environment. Accurate predictions of transition and relaminarization are critical to accurately predicting blade surface heat transfer. An approach is described which incorporates the effects of both turbulence level and scale into a CFD analysis. The model is derived from experimental data for cylindrical and elliptical leadng edges. Results using this model are compared to experimental data for both vane and rotor geometries. The comparisons are made to illustrate that using a model which includes the effects of turbulence length scale improves agreement with data, and to illustrate where improvements in the modeling are needed.

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

Bionic Design of Wind Turbine Blade Based on Long-Eared Owl's Airfoil

PubMed Central

Li, Ming

2017-01-01

The main purpose of this paper is to demonstrate a bionic design for the airfoil of wind turbines inspired by the morphology of Long-eared Owl's wings. Glauert Model was adopted to design the standard blade and the bionic blade, respectively. Numerical analysis method was utilized to study the aerodynamic characteristics of the airfoils as well as the blades. Results show that the bionic airfoil inspired by the airfoil at the 50% aspect ratio of the Long-eared Owl's wing gives rise to a superior lift coefficient and stalling performance and thus can be beneficial to improving the performance of the wind turbine blade. Also, the efficiency of the bionic blade in wind turbine blades tests increases by 12% or above (up to 44%) compared to that of the standard blade. The reason lies in the bigger pressure difference between the upper and lower surface which can provide stronger lift. PMID:28243053

Bionic Design of Wind Turbine Blade Based on Long-Eared Owl's Airfoil.

PubMed

Tian, Weijun; Yang, Zhen; Zhang, Qi; Wang, Jiyue; Li, Ming; Ma, Yi; Cong, Qian

2017-01-01

The main purpose of this paper is to demonstrate a bionic design for the airfoil of wind turbines inspired by the morphology of Long-eared Owl's wings. Glauert Model was adopted to design the standard blade and the bionic blade, respectively. Numerical analysis method was utilized to study the aerodynamic characteristics of the airfoils as well as the blades. Results show that the bionic airfoil inspired by the airfoil at the 50% aspect ratio of the Long-eared Owl's wing gives rise to a superior lift coefficient and stalling performance and thus can be beneficial to improving the performance of the wind turbine blade. Also, the efficiency of the bionic blade in wind turbine blades tests increases by 12% or above (up to 44%) compared to that of the standard blade. The reason lies in the bigger pressure difference between the upper and lower surface which can provide stronger lift.

Investigation of HP Turbine Blade Failure in a Military Turbofan Engine

NASA Astrophysics Data System (ADS)

Mishra, R. K.; Thomas, Johny; Srinivasan, K.; Nandi, Vaishakhi; Bhatt, R. Raghavendra

2017-04-01

Failure of a high pressure (HP) turbine blade in a military turbofan engine is investigated to determine the root cause of failure. Forensic and metallurgical investigations are carried out on the affected blades. The loss of coating and the presence of heavily oxidized intergranular fracture features including substrate material aging and airfoil curling in the trailing edge of a representative blade indicate that the coating is not providing adequate oxidation protection and the blade material substrate is not suitable for the application at hand. Coating spallation followed by substrate oxidation and aging leading to intergranular cracking and localized trailing edge curling is the root cause of the blade failure. The remaining portion of the blade fracture surface showed ductile overload features in the final failure. The damage observed in downstream components is due to secondary effects.

Aerodynamic losses calculation of a turbine blade with film cooling with forward and backward injection by numerical method

NASA Astrophysics Data System (ADS)

Prajapati, Anil

Thermal efficiency and power output of gas turbines can be increased by increasing the turbine blade inlet temperature. However, the main problem is the durability of the turbine blade due to the thermal stress on it at high temperature. This has led to the development of film cooling technology, in which coolant is injected from a series of cooling holes made on the blade surface to form an insulating blanket over the blade surface. However, it has to pay the aerodynamic penalties due to the injection of coolant, which are not fully understood. Pressure loss coefficient is one of the easy and widely used parameters to determine the aerodynamic loss occurred on a turbine blade. The losses occurred on the turbine blade with forward injection and backward injection cooling are studied at a different blowing ratios by a numerical simulation, which shows that the loss is higher in the case of backward injection than in forward injection. Fan-shaped cooling holes are also considered to compare with the cylindrical holes. It is observed that the loss is increased due to the fan-shaped holes in the forward injection whereas there is not a substantial difference due to the fan-shaped holes in the backward injection. The aerodynamic loss due to the location of coolant injection is studied by using injection from the leading edge, pressure side, suction side and trailing edge respectively. The study is performed to determine the effect of incidence angles and coolant injection angles on the aerodynamic loss.

Spectral analysis of unsteady surface pressure on a pusher propeller

NASA Technical Reports Server (NTRS)

Farokhi, Saeed

1992-01-01

A propeller of an advanced turboprop testbed aircraft in pusher configuration is instrumented with 22 miniature blade-mounted transducers (BMTs) at two radii. Upstream pylon wake interaction with the propeller is the source of a one-per-cycle excitation for the blades in flight. The time history of fluctuating pressure signals over 26 flight conditions is statistically analyzed in the frequency domain. The rms amplitude of fluctuating pressure signals measured by suction surface BMTs indicates a very strong presence of the fundamental frequency over most of the upper surface. The pylon wake pressure signature on the propeller trailing edge, i.e., x/c not less than 0.80, shows predominantly random turbulence; hence, the amplitude of the fundamental frequency wave is fairly small. The resurgence of a large amplitude fundamental harmonic with coherent pylon wake signature further downstream, say at 90 percent chord, is unexpected behavior. The appearance of a dominating second propeller shaft order in the spectra of the rms pressure in transonic flight conditions identifies the presence of a two-per-cycle excitation source in the azimuthal direction. This is due to the presence of a shock wave, as evidenced by the pressure-time history plots.

An experimental investigation of the helicopter rotor blade element airloads on a model rotor in the blade stall regime

NASA Technical Reports Server (NTRS)

Fisher, R. K., Jr.; Tompkins, J. E.; Bobo, C. J.; Child, R. F.

1971-01-01

A wind tunnel test program was conducted on an eight foot diameter model rotor system to determine blade element airloads characteristics in the unstalled and stalled flight regimes. The fully articulated model rotor system utilized three blades with a Vertol 23010-1.58 airfoil section, the blades being 1/7.5 scale models of the Ch-47C rotor blades. Instrumentation was incorporated at the blade 75% radial station to measure pressure and skin friction distributions, surface streamline directions and local angle of attack. The test program was conducted in three phases; non-rotating, hover and forward flight at advance ratios of 0.15, 0.35 and 0.60. Test data were analyzed with respect to providing insight to the mechanisms affecting blade stall, particularly retreating blade stall during forward flight conditions. From such data, an assessment was made as to the applicability of current theoretical analyses used for the prediction of blade element airloads in the stall regime.

The Role of Free Stream Turbulence on the Aerodynamic Performance of a Wind Turbine Blade

NASA Astrophysics Data System (ADS)

Maldonado, Victor; Thormann, Adrien; Meneveau, Charles; Castillo, Luciano

2014-11-01

Effects of free stream turbulence with large integral scale on the aerodynamic performance of an S809 airfoil-based wind turbine blade at low Reynolds number are studied using wind tunnel experiments. A constant chord (2-D) S809 airfoil wind turbine blade model with an operating Reynolds number of 208,000 based on chord length was tested for a range of angles of attack representative of fully attached and stalled flow as encountered in typical wind turbine operation. The smooth-surface blade was subjected to a quasi-laminar free stream with very low free-stream turbulence as well as to elevated free-stream turbulence generated by an active grid. This turbulence contained large-scale eddies with levels of free-stream turbulence intensity of up to 6.14% and an integral length scale of about 60% of chord-length. The pressure distribution was acquired using static pressure taps and the lift was subsequently computed by numerical integration. The wake velocity deficit was measured utilizing hot-wire anemometry to compute the drag coefficient also via integration. In addition, the mean flow was quantified using 2-D particle image velocimetry (PIV) over the suction surface of the blade. Results indicate that turbulence, even with very large-scale eddies comparable in size to the chord-length, significantly improves the aerodynamic performance of the blade by increasing the lift coefficient and overall lift-to-drag ratio, L/D for all angles tested except zero degrees.

Numerical investigation of mist/air impingement cooling on ribbed blade leading-edge surface.

PubMed

Bian, Qingfei; Wang, Jin; Chen, Yi-Tung; Wang, Qiuwang; Zeng, Min

2017-12-01

The working gas turbine blades are exposed to the environment of high temperature, especially in the leading-edge region. The mist/air two-phase impingement cooling has been adopted to enhance the heat transfer on blade surfaces and investigate the leading-edge cooling effectiveness. An Euler-Lagrange particle tracking method is used to simulate the two-phase impingement cooling on the blade leading-edge. The mesh dependency test has been carried out and the numerical method is validated based on the available experimental data of mist/air cooling with jet impingement on a concave surface. The cooling effectiveness on three target surfaces is investigated, including the smooth and the ribbed surface with convex/concave columnar ribs. The results show that the cooling effectiveness of the mist/air two-phase flow is better than that of the single-phase flow. When the ribbed surfaces are used, the heat transfer enhancement is significant, the surface cooling effectiveness becomes higher and the convex ribbed surface presents a better performance. With the enhancement of the surface heat transfer, the pressure drop in the impingement zone increases, but the incremental factor of the flow friction is smaller than that of the heat transfer enhancement. Copyright © 2017 Elsevier Ltd. All rights reserved.

Instrumentation for measuring the dynamic pressure on rotating compressor blades

NASA Technical Reports Server (NTRS)

Grant, H. P.; Lanati, G. A.

1978-01-01

To establish the capability for measurement of oscillatory pressure on rotating blades, miniature fast response semiconductor strain gage pressure transducers (2mm x 0.33mm) were mounted in several configurations on thin titanium and steel compressor blades and subjected to pressure cycles from 1 to 310 kPa during static tests and spin tests. Static test conditions included 20 C to 150 C, 0 to 3000 tensile microstrain, -1000 to +1000 bending microstrain and + or - 650G vibration. The spin test conditions included 20 C to 82 C at 0 to 90,000G. Durability was excellent. Pressure transducer sensitivity changed by only a few percent over this range of environmental conditions. Noise signal due to oscillatory acceleration normal to the diaphragm was acceptable (0.33Pa/G). Noise signal due to oscillatory strain was acceptable (0.5 Pa/microstrain) when the transducer was mounted on a 0.05mm rubber pad, with a total buildup of 0.38mm on the measure surface. Back mounting or partial recessing to eliminate buildup, increased the strain effect to 1.2 Pa/microstrain. Flush mounting within the blade to eliminate buildup reduced the strain effect, but required development of a special transducer shape. This transducer was not available in time for spin tests. Unpredictable zero drift + or - 14 kPa ruled out the use of these mounting arrangements for accurate steady-state (D.C.) measurements on rotating blades. The two best configurations fully developed and spin tested were then successfully applied in the NAS3-20606 rotating fan flutter program for quantitative measurement of oscillatory pressure amplitudes.

TACT1- TRANSIENT THERMAL ANALYSIS OF A COOLED TURBINE BLADE OR VANE EQUIPPED WITH A COOLANT INSERT

NASA Technical Reports Server (NTRS)

Gaugler, R. E.

1994-01-01

As turbine-engine core operating conditions become more severe, designers must develop more effective means of cooling blades and vanes. In order to design reliable, cooled turbine blades, advanced transient thermal calculation techniques are required. The TACT1 computer program was developed to perform transient and steady-state heat-transfer and coolant-flow analyses for cooled blades, given the outside hot-gas boundary condition, the coolant inlet conditions, the geometry of the blade shell, and the cooling configuration. TACT1 can analyze turbine blades, or vanes, equipped with a central coolant-plenum insert from which coolant-air impinges on the inner surface of the blade shell. Coolant-side heat-transfer coefficients are calculated with the heat transfer mode at each station being user specified as either impingement with crossflow, forced convection channel flow, or forced convection over pin fins. A limited capability to handle film cooling is also available in the program. The TACT1 program solves for the blade temperature distribution using a transient energy equation for each node. The nodal energy balances are linearized, one-dimensional, heat-conduction equations which are applied at the wall-outer-surface node, at the junction of the cladding and the metal node, and at the wall-inner-surface node. At the mid-metal node a linear, three-dimensional, heat-conduction equation is used. Similarly, the coolant pressure distribution is determined by solving the set of transfer momentum equations for the one-dimensional flow between adjacent fluid nodes. In the coolant channel, energy and momentum equations for one-dimensional compressible flow, including friction and heat transfer, are used for the elemental channel length between two coolant nodes. The TACT1 program first obtains a steady-state solution using iterative calculations to obtain convergence of stable temperatures, pressures, coolant-flow split, and overall coolant mass balance. Transient calculations are based on the steady-state solutions obtained. Input to the TACT1 program includes a geometrical description of the blade and insert, the nodal spacing to be used, and the boundary conditions describing the outside hot-gas and the coolant-inlet conditions. The program output includes the value of nodal temperatures and pressures at each iteration. The final solution output includes the temperature at each coolant node, and the coolant flow rates and Reynolds numbers. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 360 computer with a central memory requirement of approximately 480K of 8 bit bytes. The TACT1 program was developed in 1978.

Calculation of Multistage Turbomachinery Using Steady Characteristic Boundary Conditions

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.

1998-01-01

A multiblock Navier-Stokes analysis code for turbomachinery has been modified to allow analysis of multistage turbomachines. A steady averaging-plane approach was used to pass information between blade rows. Characteristic boundary conditions written in terms of perturbations about the mean flow from the neighboring blade row were used to allow close spacing between the blade rows without forcing the flow to be axisymmetric. In this report the multiblock code is described briefly and the characteristic boundary conditions and the averaging-plane implementation are described in detail. Two approaches for averaging the flow properties are also described. A two-dimensional turbine stator case was used to compare the characteristic boundary conditions with standard axisymmetric boundary conditions. Differences were apparent but small in this low-speed case. The two-stage fuel turbine used on the space shuttle main engines was then analyzed using a three-dimensional averaging-plane approach. Computed surface pressure distributions on the stator blades and endwalls and computed distributions of blade surface heat transfer coefficient on three blades showed very good agreement with experimental data from two tests.

Validation of helicopter noise prediction techniques

NASA Technical Reports Server (NTRS)

Succi, G. P.

1981-01-01

The current techniques of helicopter rotor noise prediction attempt to describe the details of the noise field precisely and remove the empiricisms and restrictions inherent in previous methods. These techniques require detailed inputs of the rotor geometry, operating conditions, and blade surface pressure distribution. The purpose of this paper is to review those techniques in general and the Farassat/Nystrom analysis in particular. The predictions of the Farassat/Nystrom noise computer program, using both measured and calculated blade surface pressure data, are compared to measured noise level data. This study is based on a contract from NASA to Bolt Beranek and Newman Inc. with measured data from the AH-1G Helicopter Operational Loads Survey flight test program supplied by Bell Helicopter Textron.

Estimating Blade Section Airloads from Blade Leading-Edge Pressure Measurements

NASA Technical Reports Server (NTRS)

vanAken, Johannes M.

2003-01-01

The Tilt-Rotor Aeroacoustic Model (TRAM) test in the Duitse-Nederlandse Wind (DNW) Tunnel acquired blade pressure data for forward flight test conditions of a tiltrotor in helicopter mode. Chordwise pressure data at seven radial locations were integrated to obtain the blade section normal force. The present investigation evaluates the use of linear regression analysis and of neural networks in estimating the blade section normal force coefficient from a limited number of blade leading-edge pressure measurements and representative operating conditions. These network models are subsequently used to estimate the airloads at intermediate radial locations where only blade pressure measurements at the 3.5% chordwise stations are available.

Study of stator-vane fluctuating pressures in a turbofan engine for static and flight tests

NASA Technical Reports Server (NTRS)

Mueller, A. W.

1984-01-01

As part of a program to study the fan noise generated from turbofan engines, fluctuating surface pressures induced by fan-rotor wakes were measured on core- and bypass-stator outlet guide vanes of a modified JT15D-1 engine. Tests were conducted with the engine operating on an outdoor test stand and in flight. The amplitudes of pressures measured at fan-rotor blade-passage fundamental frequencies were generally higher and appeared less stable for the static tests than for the flight tests. Fluctuating pressures measured at the blade-passage frequency of the high-speed core compressor were interpreted to be acoustic; however, disturbance trace velocities for either the convected rotor wakes or acoustic pressures were difficult to interpret because of the complex environment.

Using a shock control bump to improve the performance of an axial compressor blade section

NASA Astrophysics Data System (ADS)

Mazaheri, K.; Khatibirad, S.

2017-03-01

Here, we use numerical analysis to study the effects of a shock control bump (SCB) on the performance of a transonic axial compressor blade section and to optimize its shape and location to improve the compressor performance. A section of the NASA rotor 67 blade is used for this study. Two Bézier curves, each consisting of seven control points, are used to model the suction and pressure surfaces of the blade section. The SCB is modeled with the Hicks-Henne function and, using five design parameters, is added to the suction side. The total pressure loss through a cascade of blade sections is selected as the cost function. A continuous adjoint optimization method is used along with a RANS solver to find a new blade section shape. A grid independence study is performed, and all optimization and flow solver algorithms are validated. Two single-point optimizations are performed in the design condition and in an off-design condition. It is shown that both optimized shapes have overall better performance for both on-design and off-design conditions. An analysis is given regarding how the SCB has changed the wave structure between blade sections resulting in a more favorable flow pattern.

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.

Active control of wake/blade-row interaction noise through the use of blade surface actuators

NASA Technical Reports Server (NTRS)

Kousen, Kenneth A.; Verdon, Joseph M.

1993-01-01

A combined analytical/computational approach for controlling of the noise generated by wake/blade-row interaction through the use of anti-sound actuators on the blade surfaces is described. A representative two-dimensional section of a fan stage, composed of an upstream fan rotor and a downstream fan exit guide vane (FEGV), is examined. An existing model for the wakes generated by the rotor is analyzed to provide realistic magnitudes for the vortical excitations imposed at the inlet to the FEGV. The acoustic response of the FEGV is determined at multiples of the blade passing frequency (BPF) by using the linearized unsteady flow analysis, LINFLO. Acoustic field contours are presented at each multiple of BPF illustrating the generated acoustic response disturbances. Anti-sound is then provided by placing oscillating control surfaces, whose lengths and locations are specified arbitrarily, on the blades. An analysis is then conducted to determine the complex amplitudes required for the control surface motions to best reduce the noise. It is demonstrated that if the number of acoustic response modes to be controlled is equal to the number of available independent control surfaces, complete noise cancellation can be achieved. A weighted least squares minimization procedure for the control equations is given for cases in which the number of acoustic modes exceeds the number of available control surfaces. The effectiveness of the control is measured by the magnitude of a propagating acoustic response vector, which is related to the circumferentially averaged sound pressure level (SPL), and is minimized by a standard least-squares minimization procedure.

Active control of wake/blade-row interaction noise through the use of blade surface actuators

NASA Astrophysics Data System (ADS)

Kousen, Kenneth A.; Verdon, Joseph M.

1993-12-01

A combined analytical/computational approach for controlling of the noise generated by wake/blade-row interaction through the use of anti-sound actuators on the blade surfaces is described. A representative two-dimensional section of a fan stage, composed of an upstream fan rotor and a downstream fan exit guide vane (FEGV), is examined. An existing model for the wakes generated by the rotor is analyzed to provide realistic magnitudes for the vortical excitations imposed at the inlet to the FEGV. The acoustic response of the FEGV is determined at multiples of the blade passing frequency (BPF) by using the linearized unsteady flow analysis, LINFLO. Acoustic field contours are presented at each multiple of BPF illustrating the generated acoustic response disturbances. Anti-sound is then provided by placing oscillating control surfaces, whose lengths and locations are specified arbitrarily, on the blades. An analysis is then conducted to determine the complex amplitudes required for the control surface motions to best reduce the noise. It is demonstrated that if the number of acoustic response modes to be controlled is equal to the number of available independent control surfaces, complete noise cancellation can be achieved. A weighted least squares minimization procedure for the control equations is given for cases in which the number of acoustic modes exceeds the number of available control surfaces. The effectiveness of the control is measured by the magnitude of a propagating acoustic response vector, which is related to the circumferentially averaged sound pressure level (SPL), and is minimized by a standard least-squares minimization procedure.

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.

Aerodynamic design of a rotor blade for minimum noise radiation

NASA Technical Reports Server (NTRS)

Karamcheti, K.; Yu, Y. H.

1974-01-01

An analysis of the aerodynamic design of a hovering rotor blade for obtaining minimum aerodynamic rotor noise has been carried out. In this analysis, which is based on both acoustical and aerodynamic considerations, attention is given only to the rotational noise due to the pressure fluctuations on the blade surfaces. The lift distribution obtained in this analysis has different characteristics from those of the conventional distribution. The present distribution shows negative lift values over a quarter of the span from the blade tip, and a maximum lift at about the midspan. Results are presented to show that the noise field is considerably affected by the shape of the lift distribution along the blade and that noise reduction of about 5 dB may be obtained by designing the rotor blade to yield minimum noise.

The experimental verification of a streamline curvature numerical analysis method applied to the flow through an axial flow fan

NASA Technical Reports Server (NTRS)

Pierzga, M. J.

1981-01-01

The experimental verification of an inviscid, incompressible through-flow analysis method is presented. The primary component of this method is an axisymmetric streamline curvature technique which is used to compute the hub-to-tip flow field of a given turbomachine. To analyze the flow field in the blade-to-blade plane of the machine, the potential flow solution of an infinite cascade of airfoils is also computed using a source model technique. To verify the accuracy of such an analysis method an extensive experimental verification investigation was conducted using an axial flow research fan. Detailed surveys of the blade-free regions of the machine along with intra-blade surveys using rotating pressure sensing probes and blade surface static pressure taps provide a one-to-one relationship between measured and predicted data. The results of this investigation indicate the ability of this inviscid analysis method to predict the design flow field of the axial flow fan test rotor to within a few percent of the measured values.

Design Optimization of a Centrifugal Fan with Splitter Blades

NASA Astrophysics Data System (ADS)

Heo, Man-Woong; Kim, Jin-Hyuk; Kim, Kwang-Yong

2015-05-01

Multi-objective optimization of a centrifugal fan with additionally installed splitter blades was performed to simultaneously maximize the efficiency and pressure rise using three-dimensional Reynolds-averaged Navier-Stokes equations and hybrid multi-objective evolutionary algorithm. Two design variables defining the location of splitter, and the height ratio between inlet and outlet of impeller were selected for the optimization. In addition, the aerodynamic characteristics of the centrifugal fan were investigated with the variation of design variables in the design space. Latin hypercube sampling was used to select the training points, and response surface approximation models were constructed as surrogate models of the objective functions. With the optimization, both the efficiency and pressure rise of the centrifugal fan with splitter blades were improved considerably compared to the reference model.

Heat transfer in internal channel of a blade: Effects of rotation in a trailing edge cooling system

NASA Astrophysics Data System (ADS)

Andrei, Luca; Andreini, Antonio; Bonanni, Leonardo; Facchini, Bruno

2012-06-01

The aerothermal performance of a trailing edge (TE) internal cooling system of a high pressure gas turbine blade was evaluated under stationary and rotating conditions. The investigated geometry consists of a 30:1 scaled model reproducing a typical wedge shaped discharge duct with one row of enlarged pedestals. The airflow pattern inside the device simulates a highly loaded rotor blade cooling scheme with a 90 [deg] turning flow from the radial hub inlet to the tangential TE outlet. Two different tip configurations were tested, the first one with a completely closed section, the second one with a 5 holes outlet surfaces discharging at ambient pressure. In order to assess rotation effects, a rotating test rig, composed of a rotating arm holding both the PMMA TE model and the instrumentation, was purposely developed and manufactured. A thin Inconel heating foil and wide band Thermo-chromic Liquid Crystals are used to perform steady state heat transfer measurements on the blade pressure side. A rotary joint ensures the pneumatic connection between the blower and the rotating apparatus; moreover several slip rings are used for both instrumentation power supply and thermocouple connection. A parallel CFD analysis involving steady-state RANS modeling was conducted to allow an insight of the flow field inside the redirecting channel and the interpedestal ducts to better interpret the developing vortical structures. Low-Reynolds grid clustering permits to integrate up to the wall both the momentum and the thermal boundary layer. Calculations were performed by means of an in-house developed pressure based solver exploiting the k-ω SST turbulence model implemented in the framework of the open-source finite volume discretization toolbox OpenFOAM®. Analyzed flow conditions correspond to Reynolds number of 20000 in the hub inlet section and angular speed varies to obtain rotation numbers in the range from 0 to 0.3. The orientation of the rotation axis is orthogonal to the heated surface as to resemble a 90 [deg] blade metal angle. Results are reported in terms of detailed heat transfer coefficient 2D maps on the suction side surface as well as spanwise profiles inside the pedestal ducts.

Development of a three-dimensional multistage inverse design method for aerodynamic matching of axial compressor blading

NASA Astrophysics Data System (ADS)

van Rooij, Michael P. C.

Current turbomachinery design systems increasingly rely on multistage Computational Fluid Dynamics (CFD) as a means to assess performance of designs. However, design weaknesses attributed to improper stage matching are addressed using often ineffective strategies involving a costly iterative loop between blading modification, revision of design intent, and evaluation of aerodynamic performance. A design methodology is presented which greatly improves the process of achieving design-point aerodynamic matching. It is based on a three-dimensional viscous inverse design method which generates the blade camber surface based on prescribed pressure loading, thickness distribution and stacking line. This inverse design method has been extended to allow blading analysis and design in a multi-blade row environment. Blade row coupling was achieved through a mixing plane approximation. Parallel computing capability in the form of MPI has been implemented to reduce the computational time for multistage calculations. Improvements have been made to the flow solver to reach the level of accuracy required for multistage calculations. These include inclusion of heat flux, temperature-dependent treatment of viscosity, and improved calculation of stress components and artificial dissipation near solid walls. A validation study confirmed that the obtained accuracy is satisfactory at design point conditions. Improvements have also been made to the inverse method to increase robustness and design fidelity. These include the possibility to exclude spanwise sections of the blade near the endwalls from the design process, and a scheme that adjusts the specified loading area for changes resulting from the leading and trailing edge treatment. Furthermore, a pressure loading manager has been developed. Its function is to automatically adjust the pressure loading area distribution during the design calculation in order to achieve a specified design objective. Possible objectives are overall mass flow and compression ratio, and radial distribution of exit flow angle. To supplement the loading manager, mass flow inlet and exit boundary conditions have been implemented. Through appropriate combination of pressure or mass flow inflow/outflow boundary conditions and loading manager objectives, increased control over the design intent can be obtained. The three-dimensional multistage inverse design method with pressure loading manager was demonstrated to offer greatly enhanced blade row matching capabilities. Multistage design allows for simultaneous design of blade rows in a mutually interacting environment, which permits the redesigned blading to adapt to changing aerodynamic conditions resulting from the redesign. This ensures that the obtained blading geometry and performance implied by the prescribed pressure loading distribution are consistent with operation in the multi-blade row environment. The developed methodology offers high aerodynamic design quality and productivity, and constitutes a significant improvement over existing approaches used to address design-point aerodynamic matching.

Measurements of Heat Transfer, Flow, and Pressures in a Simulated Turbine Blade Internal Cooling Passage

NASA Technical Reports Server (NTRS)

Russell, Louis M.; Thurman, Douglas R.; Poinsatte, Philip E.; Hippensteele, Steven A.

1998-01-01

An experimental study was made to obtain quantitative information on heat transfer, flow, and pressure distribution in a branched duct test section that had several significant features of an internal cooling passage of a turbine blade. The objective of this study was to generate a set of experimental data that could be used for validation of computer codes that would be used to model internal cooling. Surface heat transfer coefficients and entrance flow conditions were measured at nominal entrance Reynolds numbers of 45,000, 335,000, and 726,000. Heat transfer data were obtained by using a steady-state technique in which an Inconel heater sheet is attached to the surface and coated with liquid crystals. Visual and quantitative flow-field data from particle image velocimetry measurements for a plane at midchannel height for a Reynolds number of 45,000 were also obtained. The flow was seeded with polystyrene particles and illuminated by a laser light sheet. Pressure distribution measurements were made both on the surface with discrete holes and in the flow field with a total pressure probe. The flow-field measurements yielded flow-field velocities at selected locations. A relatively new method, pressure sensitive paint, was also used to measure surface pressure distribution. The pressure paint data obtained at Reynolds numbers of 335,000 and 726,000 compared well with the more standard method of measuring pressures by using discrete holes.

Numerical modeling of wind turbine aerodynamic noise in the time domain.

PubMed

Lee, Seunghoon; Lee, Seungmin; Lee, Soogab

2013-02-01

Aerodynamic noise from a wind turbine is numerically modeled in the time domain. An analytic trailing edge noise model is used to determine the unsteady pressure on the blade surface. The far-field noise due to the unsteady pressure is calculated using the acoustic analogy theory. By using a strip theory approach, the two-dimensional noise model is applied to rotating wind turbine blades. The numerical results indicate that, although the operating and atmospheric conditions are identical, the acoustical characteristics of wind turbine noise can be quite different with respect to the distance and direction from the wind turbine.

Two-dimensional cold-air cascade study of a film-cooled turbine stator blade. 1: Experimental results of pressure-surface film cooling tests

NASA Technical Reports Server (NTRS)

Moffitt, T. P.; Prust, H. W., Jr.; Bartlett, W. M.

1974-01-01

The effect of film coolant ejection from the pressure side of a stator blade was determined in a two-dimensional cascade. Stator exit surveys were made for each of six rows of coolant holes. Successive multirow tests were made with two, three, four, five, and six rows of coolant holes open. The results of the multirow tests are compared with the predicted multirow performance obtained by adding the single-row data. Results are presented in terms of stator primary-air efficiency as a function of coolant fraction.

Analytical study on a two-dimensional plane of the off-design flow properties of tandem-bladed compressor stators

NASA Technical Reports Server (NTRS)

Sanger, N. L.

1973-01-01

The flow characteristics of several tandem bladed compressor stators were analytically evaluated over a range of inlet incidence angles. The ratios of rear-segment to front-segment chord and camber were varied. Results were also compared to the analytical performance of a reference solid blade section. All tandem blade sections exhibited lower calculated losses than the solid stator. But no one geometric configuration exhibited clearly superior characteristics. The front segment accepts the major effect of overall incidence angle change. Rear- to front-segment camber ratios of 4 and greater appeared to be limited by boundary-layer separation from the pressure surface of the rear segment.

Transonic flow analysis for rotors. Part 2: Three-dimensional, unsteady, full-potential calculation

NASA Technical Reports Server (NTRS)

Chang, I. C.

1985-01-01

A numerical method is presented for calculating the three-dimensional unsteady, transonic flow past a helicopter rotor blade of arbitrary geometry. The method solves the full-potential equations in a blade-fixed frame of reference by a time-marching implicit scheme. At the far-field, a set of first-order radiation conditions is imposed, thus minimizing the reflection of outgoing wavelets from computational boundaries. Computed results are presented to highlight radial flow effects in three dimensions, to compare surface pressure distributions to quasi-steady predictions, and to predict the flow field on a swept-tip blade. The results agree well with experimental data for both straight- and swept-tip blade geometries.

Energy efficient engine high-pressure turbine supersonic cascade technology report

NASA Technical Reports Server (NTRS)

Kopper, F. C.; Milano, R.; Davis, R. L.; Dring, R. P.; Stoeffler, R. C.

1981-01-01

The performance of two vane endwall geometries and three blade sections for the high-pressure turbine was evaluated in terms of the efficiency requirements of the Energy Efficient Engine high-pressure turbine component. The van endwall designs featured a straight wall and S-wall configuration. The blade designs included a base blade, straightback blade, and overcambered blade. Test results indicated that the S-wall vane configuration and the base blade configuration offered the most promising performance characteristics for the Energy Efficient Engine high-pressure turbine component.

Fretting Fatigue Analysis of Additively Manufactured Blade Root Made of Intermetallic Ti-48Al-2Cr-2Nb Alloy at High Temperature.

PubMed

Lavella, Mario; Botto, Daniele

2018-06-21

Slots in the disk of aircraft turbines restrain the centrifugal load of blades. Contact surfaces between the blade root and the disk slot undergo high contact pressure and relative displacement that is the typical condition in which fretting occurs. The load level ranges from zero to the maximum during take-off. This cycle is repeated for each mission. In this paper, a fretting fatigue analysis of additively manufactured blades is presented. Blades are made of an intermetallic alloy γTiAl. Fretting fatigue experiments were performed at a frequency of 0.5 Hz and at a temperature of 640 °C to match the operating condition of real blades. The minimum load was fixed at 0.5 KN and three maximum loads were applied, namely 16, 18 and 20 kN. Both an analytical and a two-dimensional finite element model were used to evaluate the state of stress at the contact interfaces. The results of the analytical model showed good agreement with the numerical model. Experiments showed that cracks nucleate where the analytical model predicts the maximum contact pressure and the numerical model predicts the maximum equivalent stress. A parametric analysis performed with the analytical model indicates that there exists an optimum geometry to minimize the contact pressure. Tests showed that the component life changed dramatically with the maximum load variation. Optical topography and scanning electron microscopy (SEM) analysis reveals information about the damage mechanism.

Department of Defense High Performance Computing Modernization Program. 2007 Annual Report

DTIC Science & Technology

2008-03-01

Directorate, Kirtland AFB, NM Applications of Time-Accurate CFD in Order to Account for Blade -Row Interactions and Distortion Transfer in the Design of...Patterson AFB, OH Direct Numerical Simulations of Active Control for Low- Pressure Turbine Blades Herman Fasel, University of Arizona, Tucson, AZ (Air Force...interactions with the rotor wake . These HI-ARMS computations compare favorably with available wind tunnel test measurements of surface and flowfield

14 CFR 29.653 - Pressure venting and drainage of rotor blades.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Pressure venting and drainage of rotor... § 29.653 Pressure venting and drainage of rotor blades. (a) For each rotor blade— (1) There must be means for venting the internal pressure of the blade; (2) Drainage holes must be provided for the blade...

14 CFR 27.653 - Pressure venting and drainage of rotor blades.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Pressure venting and drainage of rotor....653 Pressure venting and drainage of rotor blades. (a) For each rotor blade— (1) There must be means for venting the internal pressure of the blade; (2) Drainage holes must be provided for the blade; and...

An Investigation of Single- and Dual-Rotation Propellers at Positive and Negative Thrust, and in Combination with an NACA 1-series D-Type Cowling at Mach Numbers up to 0.84

NASA Technical Reports Server (NTRS)

Reynolds, Robert M; Samonds, Robert I; Walker, John H

1957-01-01

An investigation has been made to determine the aerodynamic characteristics of the NACA 4-(5)(05)-041 four-blade, single-relation propeller and the NACA 4-(5)(05)-037 six- and eight-blade, dual-rotation propellers in combination with various spinners and NACA d-type spinner-cowling combinations at Mach numbers up to 0.84. Propeller force characteristics, local velocity distributions in the propeller planes, inlet pressure recoveries, and static-pressure distributions on the cowling surfaces were measured for a wide range of blade angles, advance ratios, and inlet-velocity ratios. Included are data showing: (a) the effect of extended cylindrical spinners on the characteristics of the single-rotation propeller, (b) the effect of variation of the difference in blade angle setting between the front and rear components of the dual-rotation propellers, (c) the negative- and static-thrust characteristics of the propellers with 1 series spinners, and (d) the effects of ideal- and platform-type propeller-spinner junctures on the pressure-recovery characteristics of the single-rotation propeller-spinner-cowling combination.

An inverse inviscid method for the design of quasi-three dimensional rotating turbomachinery cascades

NASA Technical Reports Server (NTRS)

Bonataki, E.; Chaviaropoulos, P.; Papailiou, K. D.

1991-01-01

A new inverse inviscid method suitable for the design of rotating blade sections lying on an arbitrary axisymmetric stream-surface with varying streamtube width is presented. The geometry of the axisymmetric stream-surface and the streamtube width variation with meridional distance, the number of blades, the inlet flow conditions, the rotational speed and the suction and pressure side velocity distributions as functions of the normalized arc-length are given. The flow is considered irrotational in the absolute frame of reference and compressible. The output of the computation is the blade section that satisfies the above data. The method solves the flow equations on a (phi 1, psi) potential function-streamfunction plane for the velocity modulus, W and the flow angle beta; the blade section shape can then be obtained as part of the physical plane geometry by integrating the flow angle distribution along streamlines. The (phi 1, psi) plane is defined so that the monotonic behavior of the potential function is guaranteed, even in cases with high peripheral velocities. The method is validated on a rotating turbine case and used to design new blades. To obtain a closed blade, a set of closure conditions were developed and referred.

Analysis of surface scale on the Ni-based superalloy CMSX-10N and proposed mechanism of formation

NASA Astrophysics Data System (ADS)

Simmonds, S.; D'Souza, N.; Ryder, K. S.; Dong, H.

2012-01-01

There is a continuing demand to raise the operating temperature of jet engine turbine blades to meet the need for higher turbine entry temperatures (TET) in order to increase thermal efficiency and thrust. Modern, high-pressure turbine blades are made from Ni-based superalloys in single-crystal form via the investment casting process. One important post-cast surface defect, known as 'surface scale', has been investigated on the alloy CMSX-10N. This is an area of distinct discolouration of the aerofoil seen after casting. Auger electron and X-ray photoelectron spectroscopy analysis were carried out on both scaled and un-scaled areas. In the scaled region, a thin layer (~800nm) of Ni oxide is evident. In the un-scaled regions there is a thicker Al2O3 layer. It is shown that, as the blade cools during casting, differential thermal contraction of mould and alloy causes the solid blade to 'detach' from the mould in these scaled areas. The formation of Ni Oxides is facilitated by this separation.

A modal aeroelastic analysis scheme for turbomachinery blading. M.S. Thesis - Case Western Reserve Univ. Final Report

NASA Technical Reports Server (NTRS)

Smith, Todd E.

1991-01-01

An aeroelastic analysis is developed which has general application to all types of axial-flow turbomachinery blades. The approach is based on linear modal analysis, where the blade's dynamic response is represented as a linear combination of contributions from each of its in-vacuum free vibrational modes. A compressible linearized unsteady potential theory is used to model the flow over the oscillating blades. The two-dimensional unsteady flow is evaluated along several stacked axisymmetric strips along the span of the airfoil. The unsteady pressures at the blade surface are integrated to result in the generalized force acting on the blade due to simple harmonic motions. The unsteady aerodynamic forces are coupled to the blade normal modes in the frequency domain using modal analysis. An iterative eigenvalue problem is solved to determine the stability of the blade when the unsteady aerodynamic forces are included in the analysis. The approach is demonstrated by applying it to a high-energy subsonic turbine blade from a rocket engine turbopump power turbine. The results indicate that this turbine could undergo flutter in an edgewise mode of vibration.

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.

Design and analysis of axial aspirated compressor stages

NASA Astrophysics Data System (ADS)

Merchant, Ali A.

The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of two unique aspirated compressor stages: a low-speed stage with a design pressure ratio of 1.6 at a tip speed of 750 ft/s, and a high-speed stage with a design pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated compressor stages were designed using a new procedure which is a synthesis of low speed and high speed blade design techniques combined with a flexible inverse design method which enabled precise independent control over the shape of the blade suction and pressure surfaces. Integration of the boundary layer suction calculation into the overall design process is an essential ingredient of the new procedure. The blade design system consists of two axisymmetric through-flow codes coupled with a quasi three-dimensional viscous cascade plane code with inverse design capability. Validation of the completed designs were carried out with three-dimensional Euler and Navier-Stokes calculations. A single spanwise slot on the blade suction surface is used to bleed the boundary layer. The suction mass flow requirement for the low-speed and high-speed stages are 1% and 4% of the inlet mass flow, respectively. Additional suction between 1-2% is also required on the compressor endwalls near shock impingement locations. The rotor is modeled with a tip shroud to eliminate tip clearance effects and to discharge the suction flow radially from the flowpath. Three-dimensional viscous evaluation of the designs showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The suction requirements predicted by the quasi three-dimensional calculation were confirmed by the three-dimensional viscous calculations. The three-dimensional viscous analysis predicted a peak pressure ratio of 1.59 at an isentropic efficiency of 89% for the low-speed stage, and a peak pressure ratio of 3.68 at an isentropic efficiency of 94% for the high-speed rotor. (Copies available exclusively from MIT Libraries, Rm. 14-0551, Cambridge, MA 02139-4307. Ph. 617-253-5668; Fax 617-253-1690.)

Intermittent Behavior of the Separated Boundary Layer along the Suction Surface of a Low Pressure Turbine Blade under Periodic Unsteady Flow Conditions

NASA Technical Reports Server (NTRS)

Oeztuerk, B; Schobeiri, M. T.; Ashpis, David E.

2005-01-01

The paper experimentally and theoretically studies the effects of periodic unsteady wake flow and aerodynamic characteristics on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experiments were carried out at Reynolds number of 110,000 (based on suction surface length and exit velocity). For one steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, intermittency behaviors were experimentally and theoretically investigated. The current investigation attempts to extend the intermittency unsteady boundary layer transition model developed in previously to the LPT cases, where separation occurs on the suction surface at a low Reynolds number. The results of the unsteady boundary layer measurements and the intermittency analysis were presented in the ensemble-averaged and contour plot forms. The analysis of the boundary layer experimental data with the flow separation, confirms the universal character of the relative intermittency function which is described by a Gausssian function.

Analysis of cavitation damage on the Space Shuttle main engine high pressure oxidizer turbopump

NASA Technical Reports Server (NTRS)

Stinebring, D. R.

1985-01-01

The performance of the Space Shuttle Main Engines (SSME) has met or exceeded specifications. However, the durability for selected components has not met the desired lifetime criteria. Thus, the High-Pressure Oxidizer Turbopump (HPOTP) has experienced cavitation erosion problems in a number of locations in the pump. An investigation was conducted, taking into account an analysis of the cavitation damage, the development of a flow model for the pump, and the recommendation of design changes which would increase the life expectancy of the unit. The present paper is concerned with the cavitation damage analysis. A model is presented which relates the heavy damage on the housing and over the inducer blades to unsteady blade surface cavitation. This cavitation occurs on the inducer blades in the wakes downstream of the pump inlet housing vanes.

Streamwise vorticity in a turbine rotor with conical endwalls

NASA Astrophysics Data System (ADS)

Kost, Friedrich

1993-04-01

To investigate the spatial flow structure caused by sweep and dihedral effects in turbomachinery blade rows, detailed measurements were conducted in a windtunnel for rotating annular cascades. The special configuration consisted of a turbine rotor equipped with straight blades, a conical hub, and a conical casing with a cone half angle of 30 deg. Numerous flow data were obtained from surface pressure distributions at seven radial blade sections and from laser velocimetry upstream, downstream, and inside the rotor. It is shown that large deviations from an axisymmetric surface exist in conical flow. The conical flow gives rise to the production of streamwise vorticity which results in increased flow losses. It is furthermore shown that the secondary flow structure is mainly determined by the rotation of the turbine.

Numerical Investigation of Cavitation Improvement for a Francis Turbine

NASA Astrophysics Data System (ADS)

Yao, Zhifeng; Xiao, Ruofu; Wang, Fujun; Yang, Wei

2015-12-01

Cavitation in hydraulic machine is undesired due to its negative effects on performances. To improve cavitation performance of a Francis turbine without the change of the best efficiency point, a model runner geometry optimization was carried out. Firstly, the runner outlet diameter was appropriately increased to reduce the flow velocity at runner outlet region. Then, to avoid the change of the flow rate at the best efficiency point, the blade shapes were carefully adjusted by decreasing the blade outlet angles and increasing the blade wrap angles. A large number of the modified runners were tested by computational fluid dynamic (CFD) method. Finally the most appropriate one was selected, which has the runner outlet diameter 10% larger, the blade outlet angles 3 degrees smaller and the blade wrap angles 5 degrees larger. The results showed that the critical cavitation coefficient of the model runner decreased at every unit rotational speed after the optimization, and the effect was much remarkable at relative high flow rate. Besides, by analysing the internal flow field, it was found that the zone of the low pressure on pressure surface of the optimized turbine blades was reduced, the backflow and vortex rope in draft tube were reduced, and the cavitation zone was reduced obviously.

Methodology of Blade Unsteady Pressure Measurement in the NASA Transonic Flutter Cascade

NASA Technical Reports Server (NTRS)

Lepicovsky, J.; McFarland, E. R.; Capece, V. R.; Jett, T. A.; Senyitko, R. G.

2002-01-01

In this report the methodology adopted to measure unsteady pressures on blade surfaces in the NASA Transonic Flutter Cascade under conditions of simulated blade flutter is described. The previous work done in this cascade reported that the oscillating cascade produced waves, which for some interblade phase angles reflected off the wind tunnel walls back into the cascade, interfered with the cascade unsteady aerodynamics, and contaminated the acquired data. To alleviate the problems with data contamination due to the back wall interference, a method of influence coefficients was selected for the future unsteady work in this cascade. In this approach only one blade in the cascade is oscillated at a time. The majority of the report is concerned with the experimental technique used and the experimental data generated in the facility. The report presents a list of all test conditions for the small amplitude of blade oscillations, and shows examples of some of the results achieved. The report does not discuss data analysis procedures like ensemble averaging, frequency analysis, and unsteady blade loading diagrams reconstructed using the influence coefficient method. Finally, the report presents the lessons learned from this phase of the experimental effort, and suggests the improvements and directions of the experimental work for tests to be carried out for large oscillation amplitudes.

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).

Non-destructive measurement and role of surface residual stress monitoring in residual life assessment of a steam turbine blading material

NASA Astrophysics Data System (ADS)

Prabhu-Gaunkar, Gajanana; Rawat, M. S.; Prasad, C. R.

2014-02-01

Steam turbine blades in power generation equipment are made from martensitic stainless steels having high strength, good toughness and corrosion resistance. However, these steels are susceptible to pitting which can promote early failures of blades in the turbines, particularly in the low pressure dry/wet areas by stress corrosion and corrosion fatigue. Presence of tensile residual stresses is known to accelerate failures whereas compressive stresses can help in delaying failures. Shot peening has been employed as an effective tool to induce compressive residual stresses which offset a part of local surface tensile stresses in the surface layers of components. Maintaining local stresses at stress raisers, such as pits formed during service, below a threshold level can help in preventing the initiation microcracks and failures. The thickness of the layer in compression will, however, depend of the shot peening parameters and should extend below the bottom of corrosion pits. The magnitude of surface compressive drops progressively during service exposure and over time the effectiveness of shot peening is lost making the material susceptible to micro-crack initiation once again. Measurement and monitoring of surface residual stress therefore becomes important for assessing residual life of components in service. This paper shows the applicability of surface stress monitoring to life assessment of steam turbine blade material based on data generated in laboratory on residual surface stress measurements in relation to fatigue exposure. An empirical model is proposed to calculate the remaining life of shot peened steam turbine blades in service.

Helicopter Rotor Blade Computation in Unsteady Flows Using Moving Overset Grids

NASA Technical Reports Server (NTRS)

Ahmad, Jasim; Duque, Earl P. N.

1996-01-01

An overset grid thin-layer Navier-Stokes code has been extended to include dynamic motion of helicopter rotor blades through relative grid motion. The unsteady flowfield and airloads on an AH-IG rotor in forward flight were computed to verify the methodology and to demonstrate the method's potential usefulness towards comprehensive helicopter codes. In addition, the method uses the blade's first harmonics measured in the flight test to prescribe the blade motion. The solution was impulsively started and became periodic in less than three rotor revolutions. Detailed unsteady numerical flow visualization techniques were applied to the entire unsteady data set of five rotor revolutions and exhibited flowfield features such as blade vortex interaction and wake roll-up. The unsteady blade loads and surface pressures compare well against those from flight measurements. Details of the method, a discussion of the resulting predicted flowfield, and requirements for future work are presented. Overall, given the proper blade dynamics, this method can compute the unsteady flowfield of a general helicopter rotor in forward flight.

Heat transfer rate and film cooling effectiveness measurements in a transient cascade

NASA Astrophysics Data System (ADS)

Schultz, D. L.; Oldfield, M. L. G.; Jones, T. V.

1980-09-01

A transient cascade useful for heat transfer rate measurements is briefly described. The facility employs a free piston which compresses the test gas to temperatures around 450 K and pressures of about 3.5 to 7.5 Atm. The model is initially at room temperature and it is necessary to attain the correct gas to wall temperature ratio. The exit Mach number is set by the inlet total pressure and the pressure in the exit dump tank. Thin film heat transfer gauges are used for the measurement of heat transfer rate, deposited on machineable glass ceramic blades. The inherently fast response of these transducers makes them useful for the investigation of boundary layer transition on blade surfaces and some typical results are included.

Effects of splitter blades on the flows and characteristics in centrifugal impellers

NASA Astrophysics Data System (ADS)

Miyamoto, Hiroyuki; Nakashima, Yukitoshi; Ohba, Hideki

1992-05-01

A five-flow pressure probe was used to perform flow measurements in an unshrouded impeller and in a shrouded impeller with splitter blades. The effect of the splitter blade on passage flow and impeller performance was analyzed by comparing the results with those of impellers without splitter blades. It is found that, in impellers with splitter blades, the blade loadings tend to become smaller, and the absolute circumferential velocities and total pressures become considerably larger than those in impellers without splitter blades. It is noted that the splitter blade effect on static pressure differs between the unshrouded and shrouded impellers.

Design and Test of Mixed-flow Impellers III : Design and Experimental Results for Impeller Model MFI-2A and Comparison with Impeller Model MFI-1A

NASA Technical Reports Server (NTRS)

Hamrick, Joseph T; Osborn, Walter M; Beede, William L

1953-01-01

A mixed-flow impeller was designed to give a prescribed blade-surface velocity distribution at mean blade height for a given hub-shroud profile. The blade shape at mean blade height, which was produced by the prescribed velocity distribution, was extended by means of radial lines to form the composite blade shape from hub to shroud. The resulting blade was relatively thick; therefore, it was necessary to retain the inverse blade taper which resulted from extension of the radial lines in order to prevent merging or near merging of the separate blades near the hub. For the first test version of the impeller, designated the MFI-2A, the blade height was arbitrarily made greater than that for the basic impeller (the MFI-2) to allow for viscous effects. At design equivalent speed of 1400 feet per second the peak pressure ratio and maximum adiabatic efficiency were 3.95 and 79 percent, respectively. The adiabatic efficiency of the MFI-2A is four points lower than that for impeller model MFI-1A, but because of the higher slip factor for the MFI-2A, the pressure ratios are approximately equal. The procedures followed in the design of the MFI-1A and MFI-2A were, in general, the same; and, although the prescribed initial condition resulted in geometrical configurations that were quite dissimilar, the resulting performance characteristics compare favorably with designs for which considerable development work has been necessary.

Mechanism study on pressure fluctuation of pump-turbine runner with large blade lean angle

NASA Astrophysics Data System (ADS)

Yulin, Fan; Xuhe, Wang; Baoshan, Zhu; Dongyue, Zhou; Xijun, Zhou

2016-11-01

Excessive pressure fluctuations in the vaneless space can cause mechanical vibration and even mechanical failures in pump-turbine operation. Mechanism studies on the pressure fluctuations and optimization design of blade geometry to reduce the pressure fluctuations have important significance in industrial production. In the present paper, two pump-turbine runners with big positive and negative blade lean angle were designed by using a multiobjective design strategy. Model test showed that the runner with negative blade lean angle not only had better power performance, but also had lower pressure fluctuation than the runner with positive blade lean angle. In order to figure out the mechanism of pressure fluctuation reduction in the vaneless;jik8space, full passage model for both runners were built and transient CFD computations were conducted to simulate the flow states inside the channel. Detailed flow field analyses indicated that the difference of low-pressure area in the trailing edge of blade pressure side were the main causes of pressure fluctuation reduction in the vaneless space.

Blade Tip Pressure Measurements Using Pressure Sensitive Paint

NASA Technical Reports Server (NTRS)

Wong, Oliver D.; Watkins, Anthony Neal; Goodman, Kyle Z.; Crafton, James; Forlines, Alan; Goss, Larry; Gregory, James W.; Juliano, Thomas J.

2012-01-01

This paper discusses the application of pressure sensitive paint using laser-based excitation for measurement of the upper surface pressure distribution on the tips of rotor blades in hover and simulated forward flight. The testing was conducted in the Rotor Test Cell and the 14- by 22-ft Subsonic Tunnel at the NASA Langley Research Center on the General Rotor Model System (GRMS) test stand. The Mach-scaled rotor contained three chordwise rows of dynamic pressure transducers for comparison with PSP measurements. The rotor had an 11 ft 1 in. diameter, 5.45 in. main chord and a swept, tapered tip. Three thrust conditions were examined in hover, C(sub T) = 0.004, 0.006 and 0.008. In forward flight, an additional thrust condition, C(sub T) = 0.010 was also examined. All four thrust conditions in forward flight were conducted at an advance ratio of 0.35.

Performance of Single-Stage Turbine of Mark 25 Torpedo Power Plant with Two Special Nozzles. II; Efficiency with 20 Degrees-Inlet-Angle Rotor Blades

NASA Technical Reports Server (NTRS)

Schum, Harold J.; Whitney, Warren J.

1949-01-01

A single-stage modification of the turbine from a Mark 25 torpedo power plant was investigated to determine the performance with two nozzle designs in combination with special rotor blades having a 20 inlet angle. The performance is presented in terms of blade, rotor, and brake efficiency as a function of blade-jet speed ratio for pressure ratios of 8, 15 (design), and 20. The blade efficiency with the nozzle having circular pas- sages (K) was equal to or higher than that with the nozzle having rectangular passages (J) for all pressure ratios and speeds investigated. The maximum blade efficiency of 0.571 was obtained with nozzle K at a pressure ratio of 8 and a blade-jet speed ratio of 0.296. The difference in blade efficiency was negligible at a pressure ratio of 8 at the low speeds; the maxim difference was 0.040 at a pressure ratio of 20 and a blade-jet speed ratio of 0.260.

Aerodynamic and aeroacoustic for wind turbine

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

Mohamed, Maizi; Rabah, Dizene

2015-03-10

This paper describes a hybrid approach forpredicting noise radiated from the rotating Wind Turbine (HAWT) blades, where the sources are extracted from an unsteady Reynolds-Averaged-Navier Stocks (URANS) simulation, ANSYS CFX 11.0, was used to calculate The near-field flow parameters around the blade surface that are necessary for FW-H codes. Comparisons with NREL Phase II experimental results are presented with respect to the pressure distributions for validating a capacity of the solver to calculate the near-field flow on and around the wind turbine blades, The results show that numerical data have a good agreement with experimental. The acoustic pressure, presented asmore » a sum of thickness and loading noise components, is analyzed by means of a discrete fast Fourier transformation for the presentation of the time acoustic time histories in the frequency domain. The results convincingly show that dipole source noise is the dominant noise source for this wind turbine.« less

Acoustic pressures emanating from a turbomachine stage

NASA Technical Reports Server (NTRS)

Ramachandra, S. M.

1984-01-01

A knowledge of the acoustic energy emission of each blade row of a turbomachine is useful for estimating the overall noise level of the machine and for determining its discrete frequency noise content. Because of the close spacing between the rotor and stator of a compressor stage, the strong aerodynamic interactions between them have to be included in obtaining the resultant flow field. A three dimensional theory for determining the discrete frequency noise content of an axial compressor consisting of a rotor and a stator each with a finite number of blades are outlined. The lifting surface theory and the linearized equation of an ideal, nonsteady compressible fluid motion are used for thin blades of arbitrary cross section. The combined pressure field at a point of the fluid is constructed by linear addition of the rotor and stator solutions together with an interference factor obtained by matching them for net zero vorticity behind the stage.

Removably attachable snubber assembly

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

Martin, Jr., Nicholas F.; Wiebe, David J.

A removably attachable snubber assembly for turbine blades includes a turbine blade airfoil including a trailing edge and a leading edge joined by a pressure side and a suction side to provide an outer surface extending in a radial direction to a tip. At least one snubber attachment platform is integrally formed onto the outer surface of the turbine blade airfoil. The at least one snubber attachment platform includes an interlocking mechanism. A snubber is removably attachable to the at least one snubber attachment platform, the snubber including a first end, a second end, a trailing edge, a leading edge,more » a snubber length, and a snubber width. The snubber also includes a removable attachment mechanism on at least one of the first end and the second end that connects with the interlocking mechanism on the at least one snubber attachment platform.« less

A study of pump cavitation damage

NASA Astrophysics Data System (ADS)

Brophy, M. C.; Stinebring, D. R.; Billet, M. L.

1983-11-01

The cavitation assessment for the space shuttle main engine high pressure oxidizer turbopump is documented. A model of the flow through the pump was developed. Initially, a computational procedure was used to analyze the flow through the inlet casing including the prediction of wakes downstream of the casing vanes. From these flow calculations, cavitation patterns on the inducer blades were approximated and the damage rate estimated. The model correlates the heavy damage on the housing and over the inducer with unsteady blade surface cavitation. The unsteady blade surface cavitation is due to the large incidence changes caused by the wakes of the upstream vanes. Very high cavitation damage rates are associated with this type of cavitation. Design recommendations for reducing the unsteady cavitation include removing the set of vanes closest to the inducer and modifying the remaining vanes.

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.

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.

Blade Pressure Distribution for a Moderately Loaded Propeller.

DTIC Science & Technology

1980-09-01

lifting surface, ft 2 s chordwise location as fraction of chord length t time , sec t maximum thickness of blade, ft0 U free stream velocity, ft/sec (design...developed in Reference 1, it takes into account the quadratic form of the Bernoulli equation, since the pertubation velocities are some- times of the...normal derivatives at the loading and control point, respectively. It should be noted that the time factor has been eliminated from both sides of Eq. (3

Interaction of a turbulent vortex with a lifting surface

NASA Technical Reports Server (NTRS)

Lee, D. J.; Roberts, L.

1985-01-01

The impulsive noise due to blade-vortex-interaction is analyzing in the time domain for the extreme case when the blade cuts through the center of the vortex core with the assumptions of no distortion of the vortex path or of the vortex core. An analytical turbulent vortex core model, described in terms of the tip aerodynamic parameters, is used and its effects on the unsteady loading and maximum acoustic pressure during the interaction are determined.

A comparison of the calculated and experimental off-design performance of a radial flow turbine

NASA Technical Reports Server (NTRS)

Tirres, Lizet

1992-01-01

Off design aerodynamic performance of the solid version of a cooled radial inflow turbine is analyzed. Rotor surface static pressure data and other performance parameters were obtained experimentally. Overall stage performance and turbine blade surface static to inlet total pressure ratios were calculated by using a quasi-three dimensional inviscid code. The off design prediction capability of this code for radial inflow turbines shows accurate static pressure prediction. Solutions show a difference of 3 to 5 points between the experimentally obtained efficiencies and the calculated values.

A comparison of the calculated and experimental off-design performance of a radial flow turbine

NASA Technical Reports Server (NTRS)

Tirres, Lizet

1991-01-01

Off design aerodynamic performance of the solid version of a cooled radial inflow turbine is analyzed. Rotor surface static pressure data and other performance parameters were obtained experimentally. Overall stage performance and turbine blade surface static to inlet total pressure ratios were calculated by using a quasi-three dimensional inviscid code. The off design prediction capability of this code for radial inflow turbines shows accurate static pressure prediction. Solutions show a difference of 3 to 5 points between the experimentally obtained efficiencies and the calculated values.

Structural Evaluation of Exo-Skeletal Engine Fan Blades

NASA Technical Reports Server (NTRS)

Kuguoglu, Latife; Abumeri, Galib; Chamis, Christos C.

2003-01-01

The available computational simulation capability is used to demonstrate the structural viability of composite fan blades of innovative Exo-Skeletal Engine (ESE) developed at NASA Glenn Research Center for a subsonic mission. Full structural analysis and progressive damage evaluation of ESE composite fan blade is conducted through the NASA in-house computational simulation software system EST/BEST. The results of structural assessment indicate that longitudinal stresses acting on the blade are in compression. At a design speed of 2000 rpm, pressure and suction surface outer most ply stresses in longitudinal, transverse and shear direction are much lower than the corresponding composite ply strengths. Damage is initiated at 4870 rpm and blade fracture takes place at rotor speed of 7735 rpm. Damage volume is 51 percent. The progressive damage, buckling, stress and strength results indicate that the design at hand is very sound because of the factor of safety, damage tolerance, and buckling load of 6811 rpm.

Deep Part Load Flow Analysis in a Francis Model turbine by means of two-phase unsteady flow simulations

NASA Astrophysics Data System (ADS)

Conrad, Philipp; Weber, Wilhelm; Jung, Alexander

2017-04-01

Hydropower plants are indispensable to stabilize the grid by reacting quickly to changes of the energy demand. However, an extension of the operating range towards high and deep part load conditions without fatigue of the hydraulic components is desirable to increase their flexibility. In this paper a model sized Francis turbine at low discharge operating conditions (Q/QBEP = 0.27) is analyzed by means of computational fluid dynamics (CFD). Unsteady two-phase simulations for two Thoma-number conditions are conducted. Stochastic pressure oscillations, observed on the test rig at low discharge, require sophisticated numerical models together with small time steps, large grid sizes and long simulation times to cope with these fluctuations. In this paper the BSL-EARSM model (Explicit Algebraic Reynolds Stress) was applied as a compromise between scale resolving and two-equation turbulence models with respect to computational effort and accuracy. Simulation results are compared to pressure measurements showing reasonable agreement in resolving the frequency spectra and amplitude. Inner blade vortices were predicted successfully in shape and size. Surface streamlines in blade-to-blade view are presented, giving insights to the formation of the inner blade vortices. The acquired time dependent pressure fields can be used for quasi-static structural analysis (FEA) for fatigue calculations in the future.

Heat transfer to throat tubes in a square-chambered rocket engine at the NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Nesbitt, James A.; Brindley, William J.

1989-01-01

A gaseous H2/O2 rocket engine was constructed at the NASA-Lewis to provide a high heat flux source representative of the heat flux to the blades in the high pressure fuel turbopump (HPFTP) during startup of the space shuttle main engines. The high heat flux source was required to evaluate the durability of thermal barrier coatings being investigated for use on these blades. The heat transfer, and specifically, the heat flux to tubes located at the throat of the test rocket engine was evaluated and compared to the heat flux to the blades in the HPFTP during engine startup. Gas temperatures, pressures and heat transfer coefficients in the test rocket engine were measured. Near surface metal temperatures below thin thermal barrier coatings were also measured at various angular orientations around the throat tube to indicate the angular dependence of the heat transfer coefficients. A finite difference model for a throat tube was developed and a thermal analysis was performed using the measured gas temperatures and the derived heat transfer coefficients to predict metal temperatures in the tube. Near surface metal temperatures of an uncoated throat tube were measured at the stagnation point and showed good agreement with temperatures predicted by the thermal model. The maximum heat flux to the throat tube was calculated and compared to that predicted for the leading edge of an HPFTP blade. It is shown that the heat flux to an uncooled throat tube is slightly greater than the heat flux to an HPFTP blade during engine startup.

A computer program to predict rotor rotational noise of a stationary rotor from blade loading coefficient

NASA Technical Reports Server (NTRS)

Ramakrishnan, R.; Randall, D.; Hosier, R. N.

1976-01-01

The programing language used is FORTRAN IV. A description of all main and subprograms is provided so that any user possessing a FORTRAN compiler and random access capability can adapt the program to his facility. Rotor blade surface-pressure spectra can be used by the program to calculate: (1) blade station loading spectra, (2) chordwise and/or spanwise integrated blade-loading spectra, and (3) far-field rotational noise spectra. Any of five standard inline functions describing the chordwise distribution of the blade loading can be chosen in order to study parametrically the acoustic predictions. The program output consists of both printed and graphic descriptions of the blade-loading coefficient spectra and far-field acoustic spectrum. The results may also be written on binary file for future processing. Examples of the application of the program along with a description of the rotational noise prediction theory on which the program is based are also provided.

Three-dimensional inviscid analysis of radial-turbine flow and a limited comparison with experimental data

NASA Technical Reports Server (NTRS)

Choo, Y. K.; Civinskas, K. C.

1985-01-01

The three-dimensional inviscid DENTON code is used to analyze flow through a radial-inflow turbine rotor. Experimental data from the rotor are compared with analytical results obtained by using the code. The experimental data available for comparison are the radial distributions of circumferentially averaged values of absolute flow angle and total pressure downstream of the rotor exit. The computed rotor-exit flow angles are generally underturned relative to the experimental values, which reflect the boundary-layer separation at the trailing edge and the development of wakes downstream of the rotor. The experimental rotor is designed for a higher-than-optimum work factor of 1.126 resulting in a nonoptimum positive incidence and causing a region of rapid flow adjustment and large velocity gradients. For this experimental rotor, the computed radial distribution of rotor-exit to turbine-inlet total pressure ratios are underpredicted due to the errors in the finite-difference approximations in the regions of rapid flow adjustment, and due to using the relatively coarser grids in the middle of the blade region where the flow passage is highly three-dimensional. Additional results obtained from the three-dimensional inviscid computation are also presented, but without comparison due to the lack of experimental data. These include quasi-secondary velocity vectors on cross-channel surfaces, velocity components on the meridional and blade-to-blade surfaces, and blade surface loading diagrams. Computed results show the evolution of a passage vortex and large streamline deviations from the computational streamwise grid lines. Experience gained from applying the code to a radial turbine geometry is also discussed.

Three-dimensional inviscid analysis of radial turbine flow and a limited comparison with experimental data

NASA Technical Reports Server (NTRS)

Choo, Y. K.; Civinskas, K. C.

1985-01-01

The three-dimensional inviscid DENTON code is used to analyze flow through a radial-inflow turbine rotor. Experimental data from the rotor are compared with analytical results obtained by using the code. The experimental data available for comparison are the radial distributions of circumferentially averaged values of absolute flow angle and total pressure downstream of the rotor exit. The computed rotor-exit flow angles are generally underturned relative to the experimental values, which reflect the boundary-layer separation at the trailing edge and the development of wakes downstream of the rotor. The experimental rotor is designed for a higher-than-optimum work factor of 1.126 resulting in a nonoptimum positive incidence and causing a region of rapid flow adjustment and large velocity gradients. For this experimental rotor, the computed radial distribution of rotor-exit to turbine-inlet total pressure ratios are underpredicted due to the errors in the finite-difference approximations in the regions of rapid flow adjustment, and due to using the relatively coarser grids in the middle of the blade region where the flow passage is highly three-dimensional. Additional results obtained from the three-dimensional inviscid computation are also presented, but without comparison due to the lack of experimental data. These include quasi-secondary velocity vectors on cross-channel surfaces, velocity components on the meridional and blade-to-blade surfaces, and blade surface loading diagrams. Computed results show the evolution of a passage vortex and large streamline deviations from the computational streamwise grid lines. Experience gained from applying the code to a radial turbine geometry is also discussed.

Measurements of Free-Space Oscillating Pressures Near Propellers at Flight Mach Numbers to 0.72

NASA Technical Reports Server (NTRS)

Kurbjun, Max C; Vogeley, Arthur W

1958-01-01

In the course of a short flight program initiated to check the theory of Garrick and Watkins (NACA rep. 1198), a series of measurements at three stations were made of the oscillating pressures near a tapered-blade plan-form propeller and rectangular-blade plan form propeller at flight Mach numbers up to 0.72. In contradiction to the results for the propeller studied in NACA rep. 1198, the oscillating pressures in the plane ahead of the propeller were found to be higher than those immediately behind the propeller. Factors such as variation in torque and thrust distribution, since the blades of the present investigation were operating above their design forward speed, may account for this contradiction. The effect of blade plan form shows that a tapered-blade plan-form propeller will produce lower sound-pressure levels than a rectangular-blade plan-form propeller for the low blade-passage harmonics (the frequencies where structural considerations are important) and produce higher sound-pressure levels for the higher blade-passage harmonics (frequencies where passenger comfort is important).

Flow and Performance Calculations of Axial Compressor near Stall Margin

NASA Astrophysics Data System (ADS)

Hwang, Yoojun; Kang, Shin-Hyoung

2010-06-01

Three-dimensional flows through a Low Speed Research Axial Compressor were numerically conducted in order to estimate the performance through unsteady and steady-state simulations. The first stage with the inlet guide vane was investigated at the design point to confirm that the rotor blade induced periodicity exists. Special attention was paid to the flow near the stall condition to inspect the flow behavior in the vicinity of the stall margin. The performance predicted under the steady-state assumption is in good agreement with the measured data. However, the steady-state calculations induce more blockage through the blade passage. Flow separations on the blade surface and end-walls are reduced when unsteady simulation is conducted. The negative jet due to the wake of the rotor blade periodically distorts the boundary layer on the surface of the stator blade and improves the performance of the compressor in terms of the pressure rise. The advantage of the unsteadiness increases as the flow rate reduces. In addition, the rotor tip leakage flow is forced downstream by the unsteadiness. Consequently, the behavior contributes to extending the range of operation by preventing the leakage flow from proceeding upstream near the stall margin.

Turbine blades and systems with forward blowing slots

DOEpatents

Zuteck, Michael D.; Zalusky, Leigh; Lees, Paul

2015-09-15

A blade for use in a wind turbine comprises a pressure side and suction side meeting at a trailing edge and leading edge. The pressure side and suction side provide lift to the turbine blade upon the flow of air from the leading edge to the trailing edge and over the pressure side and suction side. The blade includes one or more openings at the suction side, in some cases between the leading edge and the trailing edge. The one or more openings are configured to provide a pressurized fluid towards the leading edge of the blade, in some cases at an angle between about 0.degree. and 70.degree. with respect to an axis oriented from a centerline of the blade toward the leading edge.

Fiber reinforced superalloys for rocket engines

NASA Technical Reports Server (NTRS)

Petrasek, Donald W.; Stephens, Joseph R.

1989-01-01

High pressure turbopumps for advanced reusable liquid propellant rocket engines such as that for the Space Shuttle Main Engine (SSME) require turbine blade materials that operate under extreme conditions of temperature, hydrogen environment, high-cycle fatigue loading, thermal fatigue and thermal shock. Such requirements tax the capabilities of current blade materials. Based on projections of properties for tungsten fiber reinforced superalloy (FRS) composites, it was concluded that FRS turbine blades offer the potential of a several fold increase in life and over a 200 C increase in temperature capability over the current SSME blade material. FRS composites were evaluated with respect to mechanical property requirements for SSME blade applications. Compared to the current blade material, the thermal shock resistance of FRS materials is excellent, two to nine times better, and their thermal fatigue resistance is equal to or higher than the current blade material. FRS materials had excellent low and high-cycle fatigue strengths, and thermal shock-induced surface microcracks had no influence on their fatigue strength. The material also exhibited negligible embrittlement when exposed to a hydrogen environment.

Fiber reinforced superalloys for rocket engines

NASA Technical Reports Server (NTRS)

Petrasek, Donald W.; Stephens, Joseph R.

1988-01-01

High-pressure turbopumps for advanced reusable liquid-propellant rocket engines such as that for the Space Shuttle Main Engine (SSME) require turbine blade materials that operate under extreme conditions of temperature, hydrogen environment, high-cycle fatigue loading, thermal fatigue and thermal shock. Such requirements tax the capabilities of current blade materials. Based on projections of properties for tungsten fiber reinforced superalloy (FRS) composites, it was concluded that FRS turbine blades offer the potential of a several-fold increase in life and over a 200C increase in temperature capability over current SSME blade material. FRS composites were evaluated with respect to mechanical property requirements for SSME blade applications. Compared to the current blade material, the thermal shock resistance of FRS materials is excellent, two to nine times better, and their thermal fatigue resistance is equal to or higher than the current blade material. FRS materials had excellent low and high-cycle fatigue strengths, and thermal shock-induced surface microcracks had no influence on their fatigue strength. The material also exhibited negligible embrittlement when exposed to a hydrogen environment.

Unsteady Loss in the Stator Due to the Incoming Rotor Wake in a Highly-Loaded Transonic Compressor

NASA Technical Reports Server (NTRS)

Hah, Chunill

2015-01-01

The present paper reports an investigation of unsteady loss generation in the stator due to the incoming rotor wake in an advanced GE transonic compressor design with a high-fidelity numerical method. This advanced compressor with high reaction and high stage loading has been investigated both experimentally and analytically in the past. The measured efficiency in this advanced compressor is significantly lower than the design intention/goal. The general understanding is that the current generation of compressor design/analysis tools miss some important flow physics in this modern compressor design. To pinpoint the source of the efficiency miss, an advanced test with a detailed flow traverse was performed for the front one and a half stage at the NASA Glenn Research Center. Detailed data-match analysis by GE identified an unexpected high loss generation in the pressure side of the stator passage. Higher total temperature and lower total pressure are measured near the pressure side of the stator. Various analyses based on the RANS and URANS of the compressor stage do not calculate the measured higher total temperature and lower total pressure on the pressure side of the stator. In the present paper, a Large Eddy Simulation (LES) is applied to find the fundamental mechanism of this unsteady loss generation in the stator due to the incoming rotor wake. The results from the LES were first compared with the NASA test results and the GE interpretation of the test data. LES calculates lower total pressure and higher total temperature on the pressure side of the stator, as the measured data showed, resulting in large loss generation on the pressure side of the stator. Detailed examination of the unsteady flow field from LES shows that the rotor wake, which has higher total temperature and higher total pressure relative to the free stream, interacts quite differently with the pressure side of the blade compared to the suction side of the blade. The higher temperature in the wake remains high as the wake passes through the pressure side of the blade. On the other hand, the total temperature diffuses as it passes through near the suction surface. For the presently investigated compressor, the classical intra-stator wake transport to the pressure side of the blade by the slip velocity in the wake seems to be minor. The main causes of this phenomenon are three-dimensional unsteady vortex interactions near the blade surface. The stabilizing effect of the concave curvature on the suction side keeps the rotor wake thin. On the other hand, the destabilizing effect of the convex curvature of the pressure side makes the rotor wake thicker, which results in a higher total temperature measurement at the stator exit. Additionally, wake stretching through the stator seems to contribute to the redistribution of the total temperature and the loss generation.

Unsteady Pressures in a Transonic Fan Cascade Due to a Single Oscillating Airfoil

NASA Technical Reports Server (NTRS)

Lepicovsky, J.; McFarland, E. R.; Capece, V. R.; Hayden, J.

2002-01-01

An extensive set of unsteady pressure data was acquired along the midspan of a modern transonic fan blade for simulated flutter conditions. The data set was acquired in a nine-blade linear cascade with an oscillating middle blade to provide a database for the influence coefficient method to calculate instantaneous blade loadings. The cascade was set for an incidence of 10 dg. The data were acquired on three stationary blades on each side of the middle blade that was oscillated at an amplitude of 0.6 dg. The matrix of test conditions covered inlet Mach numbers of 0.5, 0.8, and 1.1 and the oscillation frequencies of 200, 300, 400, and 500 Hz. A simple quasiunsteady two-dimensional computer simulation was developed to aid in the running of the experimental program. For high Mach number subsonic inlet flows the blade pressures exhibit very strong, low-frequency, self-induced oscillations even without forced blade oscillations, while for low subsonic and supersonic inlet Mach numbers the blade pressure unsteadiness is quite low. The amplitude of forced pressure fluctuations on neighboring stationary blades strongly depends on the inlet Mach number and forcing frequency. The flowfield behavior is believed to be governed by strong nonlinear effects due to a combination of viscosity, compressibility, and unsteadiness. Therefore, the validity of the quasi-unsteady simplified computer simulation is limited to conditions when the flowfield is behaving in a linear, steady manner. Finally, an extensive set of unsteady pressure data was acquired to help development and verification of computer codes for blade flutter effects.

Experimental evidence of inter-blade cavitation vortex development in Francis turbines at deep part load condition

NASA Astrophysics Data System (ADS)

Yamamoto, K.; Müller, A.; Favrel, A.; Avellan, F.

2017-10-01

Francis turbines are subject to various types of cavitation flow depending on the operating condition. To enable a smooth integration of the renewable energy sources, hydraulic machines are now increasingly required to extend their operating range, especially down to extremely low discharge conditions called deep part load operation. The inter-blade cavitation vortex is a typical cavitation phenomenon observed at deep part load operation. However, its dynamic characteristics are insufficiently understood today. In an objective of revealing its characteristics, the present study introduces a novel visualization technique with instrumented guide vanes embedding the visualization devices, providing unprecedented views on the inter-blade cavitation vortex. The binary image processing technique enables the successful evaluation of the inter-blade cavitation vortex in the images. As a result, it is shown that the probability of the inter-blade cavitation development is significantly high close to the runner hub. Furthermore, the mean vortex line is calculated and the vortex region is estimated in the three-dimensional domain for the comparison with numerical simulation results. In addition, the on-board pressure measurements on a runner blade is conducted, and the influence of the inter-blade vortex on the pressure field is investigated. The analysis suggests that the presence of the inter-blade vortex can magnify the amplitude of pressure fluctuations especially on the blade suction side. Furthermore, the wall pressure difference between pressure and suction sides of the blade features partially low or negative values near the hub at the discharge region where the inter-blade vortex develops. This negative pressure difference on the blade wall suggests the development of a backflow region caused by the flow separation near the hub, which is closely related to the development of the inter-blade vortex. The development of the backflow region is confirmed by the numerical simulation, and the physical mechanisms of the inter-blade vortex development is, furthermore, discussed.

On the Physics of Flow Separation Along a Low Pressure Turbine Blade Under Unsteady Flow Conditions

NASA Technical Reports Server (NTRS)

Schobeiri, Meinhard T.; Ozturk, Burak; Ashpis, David E.

2005-01-01

The present study, which is the first of a series of investigations dealing with specific issues of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed at Texas A&M Turbomachinery Performance and Flow Research Laboratory using a large-scale unsteady turbine cascade research facility with an integrated wake generator and test section unit. To account for a high flow deflection of LPT-cascades at design and off-design operating points, the entire wake generator and test section unit including the traversing system is designed to allow a precise angle adjustment of the cascade relative to the incoming flow. This is done by a hydraulic platform, which simultaneously lifts and rotates the wake generator and test section unit. The unit is then attached to the tunnel exit nozzle with an angular accuracy of better than 0.05 , which is measured electronically. Utilizing a Reynolds number of 110,000 based on the blade suction surface length and the exit velocity, one steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities and turbulence intensities are investigated using hot-wire anemometry. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re=50,000, 75,000, 100,000, and 125,000 at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extent of the separation zone as well as its behavior under unsteady wake flow. The results presented in ensemble-averaged and contour plot forms contribute to understanding the physics of the separation phenomenon under periodic unsteady wake flow. Several physical mechanisms are discussed.

On the Physics of Flow Separation Along a Low Pressure Turbine Blade Under Unsteady Flow Conditions

NASA Technical Reports Server (NTRS)

Schobeiri, Meinhard T.; Ozturk, Burak; Ashpis, David E.

2003-01-01

The present study, which is the first of a series of investigations dealing with specific issues of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed at Texas A&M Turbomachinery Performance and Flow Research Laboratory using a large-scale unsteady turbine cascade research facility with an integrated wake generator and test section unit. To account for a high flow deflection of LPT-cascades at design and off-design operating points, the entire wake generator and test section unit including the traversing system is designed to allow a precise angle adjustment of the cascade relative to the incoming flow. This is done by a hydraulic platform, which simultaneously lifts and rotates the wake generator and test section unit. The unit is then attached to the tunnel exit nozzle with an angular accuracy of better than 0.05 , which is measured electronically. Utilizing a Reynolds number of 110,000 based on the blade suction surface length and the exit velocity, one steady and two different unsteady inlet flowconditions with the corresponding passing frequencies, wake velocities and turbulence intensities are investigated using hot-wire anemometry. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re=50,000, 75,000, 100,000, and 125,000 at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extent of the separation zone as well as its behavior under unsteady wake flow. The results presented in ensemble-averaged and contour plot forms contribute to understanding the physics of the separation phenomenon under periodic unsteady wake flow. Several physical mechanisms are discussed.

Experimental and analytical studies of a model helicopter rotor in hover

NASA Technical Reports Server (NTRS)

Caradonna, F. X.; Tung, C.

1981-01-01

A benchmark test to aid the development of various rotor performance codes was conducted. Simultaneous blade pressure measurements and tip vortex surveys were made for a wide range of tip Mach numbers including the transonic flow regime. The measured tip vortex strength and geometry permit effective blade loading predictions when used as input to a prescribed wake lifting surface code. It is also shown that with proper inflow and boundary layer modeling, the supercritical flow regime can be accurately predicted.

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.

Snubber assembly for turbine blades

DOEpatents

Marra, John J

2013-09-03

A snubber associated with a rotatable turbine blade in a turbine engine, the turbine blade including a pressure sidewall and a suction sidewall opposed from the pressure wall. The snubber assembly includes a first snubber structure associated with the pressure sidewall of the turbine blade, a second snubber structure associated with the suction sidewall of the turbine blade, and a support structure. The support structure extends through the blade and is rigidly coupled at a first end portion thereof to the first snubber structure and at a second end portion thereof to the second snubber structure. Centrifugal loads exerted by the first and second snubber structures caused by rotation thereof during operation of the engine are at least partially transferred to the support structure, such that centrifugal loads exerted on the pressure and suctions sidewalls of the turbine blade by the first and second snubber structures are reduced.

A review of various nozzle range of wire arc spray on FeCrBMnSi metal coating

NASA Astrophysics Data System (ADS)

Purwaningsih, Hariyati; Rochiem, Rochman; Suchaimi, Muhammad; Jatimurti, Wikan; Wibisono, Alvian Toto; Kurniawan, Budi Agung

2018-04-01

Low Temperature Hot Corrosion (LTHC) is type of hot corrosion which occurred on 700-800°C and usually on turbine blades. So, as a result the material of turbine blades is crack and degredation of rotation efficiency. Hot corrosion protection with the use of barrier that separate substrate and environment is one of using metal surface coating, wire arc spray method. This study has a purpose to analyze the effect of nozzle distance and gas pressure on FeCrBMnSi coating process using wire arc spray method on thermal resistance. The parameter of nozzle distance and gas pressure are used, resulted the best parameter on distance 400 mm and gas pressure 3 bar which has the bond strength of 12,58 MPa with porosity percentage of 5,93% and roughness values of 16,36 µm. While the examination of thermal cycle which by heating and cooling continuously, on the coating surface is formed oxide compound (Fe3O4) which cause formed crack propagation and delamination. Beside that hardness of coating surface is increase which caused by precipitate boride (Fe9B)0,2

Effects of Periodic Unsteady Wake Flow and Pressure Gradient on Boundary Layer Transition Along the Concave Surface of a Curved Plate. Part 3

NASA Technical Reports Server (NTRS)

Schobeiri, M. T.; Radke, R. E.

1996-01-01

Boundary layer transition and development on a turbomachinery blade is subjected to highly periodic unsteady turbulent flow, pressure gradient in longitudinal as well as lateral direction, and surface curvature. To study the effects of periodic unsteady wakes on the concave surface of a turbine blade, a curved plate was utilized. On the concave surface of this plate, detailed experimental investigations were carried out under zero and negative pressure gradient. The measurements were performed in an unsteady flow research facility using a rotating cascade of rods positioned upstream of the curved plate. Boundary layer measurements using a hot-wire probe were analyzed by the ensemble-averaging technique. The results presented in the temporal-spatial domain display the transition and further development of the boundary layer, specifically the ensemble-averaged velocity and turbulence intensity. As the results show, the turbulent patches generated by the wakes have different leading and trailing edge velocities and merge with the boundary layer resulting in a strong deformation and generation of a high turbulence intensity core. After the turbulent patch has totally penetrated into the boundary layer, pronounced becalmed regions were formed behind the turbulent patch and were extended far beyond the point they would occur in the corresponding undisturbed steady boundary layer.

User's manual for three dimensional boundary layer (BL3-D) code

NASA Technical Reports Server (NTRS)

Anderson, O. L.; Caplin, B.

1985-01-01

An assessment has been made of the applicability of a 3-D boundary layer analysis to the calculation of heat transfer, total pressure losses, and streamline flow patterns on the surface of both stationary and rotating turbine passages. In support of this effort, an analysis has been developed to calculate a general nonorthogonal surface coordinate system for arbitrary 3-D surfaces and also to calculate the boundary layer edge conditions for compressible flow using the surface Euler equations and experimental data to calibrate the method, calculations are presented for the pressure endwall, and suction surfaces of a stationary cascade and for the pressure surface of a rotating turbine blade. The results strongly indicate that the 3-D boundary layer analysis can give good predictions of the flow field, loss, and heat transfer on the pressure, suction, and endwall surface of a gas turbine passage.

Measurement of Unsteady Aerodynamics Load on the Blade of Field Horizontal Axis Wind Turbine

NASA Astrophysics Data System (ADS)

Kamada, Yasunari; Maeda, Takao; Naito, Keita; Ouchi, Yuu; Kozawa, Masayoshi

This paper describes an experimental field study of the rotor aerodynamics of wind turbines. The test wind turbine is a horizontal axis wind turbine, or: HAWT with a diameter of 10m. The pressure distributions on the rotating blade are measured with multi point pressure transducers. Sectional aerodynamic forces are analyzed from pressure distribution. Blade root moments are measured simultaneously by a pair of strain gauges. The inflow wind is measured by a three component sonic anemometer, the local inflow of the blade section are measured by a pair of 7 hole Pitot tubes. The relation between the aerodynamic moments on the blade root from pressure distribution and the mechanical moment from strain gauges is discussed. The aerodynamic moments are estimated from the sectional aerodynamic forces and show oscillation caused by local wind speed and direction change. The mechanical moment shows similar oscillation to the aerodynamic excepting the short period oscillation of the blade first mode frequency. The fluctuation of the sectional aerodynamic force triggers resonant blade oscillations. Where stall is present along the blade section, the blade's first mode frequency is dominant. Without stall, the rotating frequency is dominant in the blade root moment.

The influence of the radial pressure gradient on the blade root loss in an annular subsonic nozzle cascade

NASA Astrophysics Data System (ADS)

Meng, D.; Weng, Z.; Xiang, Y.

1985-09-01

This paper presents a method for predicting the blade root loss in an annular nozzle cascade in which consideration is given to the influence of the radial pressure gradient (RPG) on it. The variation of blade root losses under different RPG is obtained experimentally, and finite element method is used to calculate the pressure distribution in the blade passage.

Dynamic analysis method for prevention of failure in the first-stage low-pressure turbine blade with two-finger root

NASA Astrophysics Data System (ADS)

Park, Jung-Yong; Jung, Yong-Keun; Park, Jong-Jin; Kang, Yong-Ho

2002-05-01

Failures of turbine blades are identified as the leading causes of unplanned outages for steam turbine. Accidents of low-pressure turbine blade occupied more than 70 percent in turbine components. Therefore, the prevention of failures for low pressure turbine blades is certainly needed. The procedure is illustrated by the case study. This procedure is used to guide, and support the plant manager's decisions to avoid a costly, unplanned outage. In this study, we are trying to find factors of failures in LP turbine blade and to make three steps to approach the solution of blade failure. First step is to measure natural frequency in mockup test and to compare it with nozzle passing frequency. Second step is to use FEM and to calculate the natural frequencies of 7 blades and 10 blades per group in BLADE code. Third step is to find natural frequencies of grouped blade off the nozzle passing frequency.

Complementary Aerodynamic Performance Datasets for Variable Speed Power Turbine Blade Section from Two Independent Transonic Turbine Cascades

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.; Welch, Gerard E.; Giel, Paul W.; Ames, Forrest E.; Long, Jonathon A.

2015-01-01

Two independent experimental studies were conducted in linear cascades on a scaled, two-dimensional mid-span section of a representative Variable Speed Power Turbine (VSPT) blade. The purpose of these studies was to assess the aerodynamic performance of the VSPT blade over large Reynolds number and incidence angle ranges. The influence of inlet turbulence intensity was also investigated. The tests were carried out in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility and at the University of North Dakota (UND) High Speed Compressible Flow Wind Tunnel Facility. A large database was developed by acquiring total pressure and exit angle surveys and blade loading data for ten incidence angles ranging from +15.8deg to -51.0deg. Data were acquired over six flow conditions with exit isentropic Reynolds number ranging from 0.05×106 to 2.12×106 and at exit Mach numbers of 0.72 (design) and 0.35. Flow conditions were examined within the respective facility constraints. The survey data were integrated to determine average exit total-pressure and flow angle. UND also acquired blade surface heat transfer data at two flow conditions across the entire incidence angle range aimed at quantifying transitional flow behavior on the blade. Comparisons of the aerodynamic datasets were made for three "match point" conditions. The blade loading data at the match point conditions show good agreement between the facilities. This report shows comparisons of other data and highlights the unique contributions of the two facilities. The datasets are being used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range.

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.

Aerodynamic performances of three fan stator designs operating with rotor having tip speed of 337 meters per second and pressure ratio of 1.54. Relation of analytical code calculations to experimental performance

NASA Technical Reports Server (NTRS)

Gelder, T. F.; Schmidt, J. F.; Esgar, G. M.

1980-01-01

A hub-to-shroud and a blade-to-blade internal-flow analysis code, both inviscid and basically subsonic, were used to calculate the flow parameters within four stator-blade rows. The produced ratios of maximum suction-surface velocity to trailing-edge velocity correlated well in the midspan region, with the measured total-parameters over the minimum-loss to near stall operating range for all stators and speeds studied. The potential benefits of a blade designed with the aid of these flow analysis codes are illustrated by a proposed redesign of one of the four stators studied. An overall efficiency improvement of 1.6 points above the peak measured for that stator is predicted for the redesign.

Surface Integrity of Inconel 718 by Ball Burnishing

NASA Astrophysics Data System (ADS)

Sequera, A.; Fu, C. H.; Guo, Y. B.; Wei, X. T.

2014-09-01

Inconel 718 has wide applications in manufacturing mechanical components such as turbine blades, turbocharger rotors, and nuclear reactors. Since these components are subject to harsh environments such as high temperature, pressure, and corrosion, it is critical to improve the functionality to prevent catastrophic failure due to fatigue or corrosion. Ball burnishing as a low plastic deformation process is a promising technique to enhance surface integrity for increasing component fatigue and corrosion resistance in service. This study focuses on the experimental study on surface integrity of burnished Inconel 718. The effects of burnishing ball size and pressure on surface integrity factors such as surface topography, roughness, and hardness are investigated. The burnished surfaces are smoother than the as-machined ones. Surface hardness after burnishing is higher than the as-machined surfaces, but become stable over a certain burnishing pressure. There exists an optimal process space of ball sized and burnishing pressure for surface finish. In addition, surface hardness after burnishing is higher than the as-machined surfaces, which is confirmed by statistical analysis.

Two-dimensional Cascade Investigation of the Maximum Exit Tangential Velocity Component and Other Flow Conditions at the Exit of Several Turbine Blade Designs at Supercritical Pressure Ratios

NASA Technical Reports Server (NTRS)

Hauser, Cavour H; Plohr, Henry W

1951-01-01

The nature of the flow at the exit of a row of turbine blades for the range of conditions represented by four different blade configurations was evaluated by the conservation-of-momentum principle using static-pressure surveys and by analysis of Schlieren photographs of the flow. It was found that for blades of the type investigated, the maximum exit tangential-velocity component is a function of the blade geometry only and can be accurately predicted by the method of characteristics. A maximum value of exit velocity coefficient is obtained at a pressure ratio immediately below that required for maximum blade loading followed by a sharp drop after maximum blade loading occurs.

An unsteady helicopter rotor: Fuselage interaction analysis

NASA Technical Reports Server (NTRS)

Lorber, Peter F.; Egolf, T. Alan

1988-01-01

A computational method was developed to treat unsteady aerodynamic interactions between a helicopter rotor, wake, and fuselage and between the main and tail rotors. An existing lifting line prescribed wake rotor analysis and a source panel fuselage analysis were coupled and modified to predict unsteady fuselage surface pressures and airloads. A prescribed displacement technique is used to position the rotor wake about the fuselage. Either a rigid blade or an aeroelastic blade analysis may be used to establish rotor operating conditions. Sensitivity studies were performed to determine the influence of the wake fuselage geometry on the computation. Results are presented that describe the induced velocities, pressures, and airloads on the fuselage and on the rotor. The ability to treat arbitrary geometries is demonstrated using a simulated helicopter fuselage. The computational results are compared with fuselage surface pressure measurements at several locations. No experimental data was available to validate the primary product of the analysis: the vibratory airloads on the entire fuselage. A main rotor-tail rotor interaction analysis is also described, along with some hover and forward flight.

Machining the Integral Impeller and Blisk of Aero-Engines: A Review of Surface Finishing and Strengthening Technologies

NASA Astrophysics Data System (ADS)

Fu, Youzhi; Gao, Hang; Wang, Xuanping; Guo, Dongming

2017-05-01

The integral impeller and blisk of an aero-engine are high performance parts with complex structure and made of difficult-to-cut materials. The blade surfaces of the integral impeller and blisk are functional surfaces for power transmission, and their surface integrity has significant effects on the aerodynamic efficiency and service life of an aero-engine. Thus, it is indispensable to finish and strengthen the blades before use. This paper presents a comprehensive literature review of studies on finishing and strengthening technologies for the impeller and blisk of aero-engines. The review includes independent and integrated finishing and strengthening technologies and discusses advanced rotational abrasive flow machining with back-pressure used for finishing the integral impeller and blisk. A brief assessment of future research problems and directions is also presented.

Heat transfer and flow characteristics on a gas turbine shroud.

PubMed

Obata, M; Kumada, M; Ijichi, N

2001-05-01

The work described in this paper is an experimental investigation of the heat transfer from the main flow to a turbine shroud surface, which may be applicable to ceramic gas turbines. Three kinds of turbine shrouds are considered with a flat surface, a taper surface and a spiral groove surface opposite to the blades in an axial flow turbine of actual turbo-charger. Heat transfer measurements were performed for the experimental conditions of a uniform heat flux or a uniform wall temperature. The effects of the inlet flow angle, rotational speed, and tip clearance on the heat transfer coefficient were clarified under on- and off-design flow conditions. The mean heat transfer coefficient was correlated to the blade Reynolds number and tip clearance, and compared with an experimental correlation and measurements of a flat surface. A comparison was also made for the measurement of static pressure distributions.

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.

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

77 FR 57001 - Airworthiness Directives; Pratt & Whitney Canada, Auxiliary Power Units

Federal Register 2010, 2011, 2012, 2013, 2014

2012-09-17

... exhaust duct flanges. This AD was prompted by several events of high-pressure turbine blade fracture... have experienced several events of High Pressure Turbine (HPT) blade fracture, some of which have..., 2011, and our AD will not prevent high pressure turbine blades from failing. We do not agree. The root...

Passive Acoustic Detection of Wind Turbine In-Flow Conditions for Active Control and Optimization

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

Murray, Nathan E.

2012-03-12

Wind is a significant source of energy; however, the human capability to produce electrical energy still has many hurdles to overcome. One of these is the unpredictability of the winds in the atmospheric boundary layer (ABL). The ABL is highly turbulent in both stable and unstable conditions (based on the vertical temperature profile) and the resulting fluctuations can have a dramatic impact on wind turbine operation. Any method by which these fluctuations could be observed, estimated, or predicted could provide a benefit to the wind energy industry as a whole. Based on the fundamental coupling of velocity fluctuations to pressuremore » fluctuations in the nearly incompressible flow in the ABL, This work hypothesizes that a ground-based array of infrasonic pressure transducers could be employed to estimate the vertical wind profile over a height relevant for wind turbines. To analyze this hypothesis, experiments and field deployments were conducted. Wind tunnel experiments were performed for a thick turbulent boundary layer over a neutral or heated surface. Surface pressure and velocity probe measurements were acquired simultaneously. Two field deployments yielded surface pressure data from a 49 element array. The second deployment at the Reese Technology Center in Lubbock, TX, also included data from a smaller aperture, 96-element array and a 200-meter tall meteorological tower. Analysis of the data successfully demonstrated the ability to estimate the vertical velocity profile using coherence data from the pressure array. Also, dynamical systems analysis methods were successful in identifying and tracking a gust type event. In addition to the passive acoustic profiling method, this program also investigated a rapid response Doppler SODAR system, the optimization of wind turbine blades for enhanced power with reduced aeroacoustic noise production, and the implementation of a wireless health monitoring system for the wind turbine blades. Each of these other objectives was met successfully. The use of phase unwrapping applied to SODAR data was found to yield reasonable results for per-pulse measurements. A health monitoring system design analysis was able to demonstrate the ability to use a very small number of sensors to monitor blade health based on the blade's overall structural modes. Most notable was the development of a multi-objective optimization methodology that successfully yielded an aerodynamic blade design that produces greater power output with reduced aerodynamic loading noise. This optimization method could be significant for future design work.« less

A new aerodynamic integral equation based on an acoustic formula in the time domain

NASA Technical Reports Server (NTRS)

Farassat, F.

1984-01-01

An aerodynamic integral equation for bodies moving at transonic and supersonic speeds is presented. Based on a time-dependent acoustic formula for calculating the noise emanating from the outer portion of a propeller blade travelling at high speed (the Ffowcs Williams-Hawking formulation), the loading terms and a conventional thickness source terms are retained. Two surface and three line integrals are employed to solve an equation for the loading noise. The near-field term is regularized using the collapsing sphere approach to obtain semiconvergence on the blade surface. A singular integral equation is thereby derived for the unknown surface pressure, and is amenable to numerical solutions using Galerkin or collocation methods. The technique is useful for studying the nonuniform inflow to the propeller.

Effect of Reynolds Number and Periodic Unsteady Wake Flow Condition on Boundary Layer Development, Separation, and Intermittency Behavior Along the Suction Surface of a Low Pressure Turbine Blade

NASA Technical Reports Server (NTRS)

Schobeiri, M. T.; Ozturk, B.; Ashpis, David E.

2007-01-01

The paper experimentally studies the effects of periodic unsteady wake flow and different Reynolds numbers on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experimental investigations were performed on a large scale, subsonic unsteady turbine cascade research facility at Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. The experiments were carried out at Reynolds numbers of 110,000 and 150,000 (based on suction surface length and exit velocity). One steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities, and turbulence intensities were investigated. The reduced frequencies chosen cover the operating range of LP turbines. In addition to the unsteady boundary layer measurements, surface pressure measurements were performed. The inception, onset, and the extent of the separation bubble information collected from the pressure measurements were compared with the hot wire measurements. The results presented in ensemble-averaged, and the contour plot forms help to understand the physics of the separation phenomenon under periodic unsteady wake flow and different Reynolds number. It was found that the suction surface displayed a strong separation bubble for these three different reduced frequencies. For each condition, the locations defining the separation bubble were determined carefully analyzing and examining the pressure and mean velocity profile data. The location of the boundary layer separation was dependent of the Reynolds number. It is observed that starting point of the separation bubble and the re-attachment point move further downstream by increasing Reynolds number from 110,000 to 150,000. Also, the size of the separation bubble is smaller when compared to that for Re=110,000.

Effect of Reynolds Number and Periodic Unsteady Wake Flow Condition on Boundary Layer Development, Separation, and Re-attachment along the Suction Surface of a Low Pressure Turbine Blade

NASA Technical Reports Server (NTRS)

Ozturk, B.; Schobeiri, M. T.; Ashpis, David E.

2005-01-01

The paper experimentally studies the effects of periodic unsteady wake flow and different Reynolds numbers on boundary layer development, separation and re-attachment along the suction surface of a low pressure turbine blade. The experimental investigations were performed on a large scale, subsonic unsteady turbine cascade research facility at Turbomachinery Performance and Flow Research Laboratory (TPFL) of Texas A&M University. The experiments were carried out at Reynolds numbers of 110,000 and 150,000 (based on suction surface length and exit velocity). One steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities, and turbulence intensities were investigated. The reduced frequencies chosen cover the operating range of LP turbines. In addition to the unsteady boundary layer measurements, surface pressure measurements were performed. The inception, onset, and the extent of the separation bubble information collected from the pressure measurements were compared with the hot wire measurements. The results presented in ensemble-averaged, and the contour plot forms help to understand the physics of the separation phenomenon under periodic unsteady wake flow and different Reynolds number. It was found that the suction surface displayed a strong separation bubble for these three different reduced frequencies. For each condition, the locations defining the separation bubble were determined carefully analyzing and examining the pressure and mean velocity profile data. The location of the boundary layer separation was dependent of the Reynolds number. It is observed that starting point of the separation bubble and the re-attachment point move further downstream by increasing Reynolds number from 110,000 to 150,000. Also, the size of the separation bubble is smaller when compared to that for Re=110,000.

Fluctuating pressures on fan blades of a turbofan engine: Static and wind-tunnel investigations

NASA Technical Reports Server (NTRS)

Schoenster, J. A.

1982-01-01

To investigate the fan noise generated from turbofan engines, miniature pressure transducers were used to measure the fluctuating pressure on the fan blades of a JT15D engine. Tests were conducted with the engine operating on an outdoor test stand and in a wind tunnel. It was found that a potential flow interaction between the fan blades and six, large support struts in the bypass duct is a dominant noise source in the JT15D engine. Effects of varying fan speed and the forward speed on the blade pressure are also presented.

Leading edge film cooling effects on turbine blade heat transfer

NASA Technical Reports Server (NTRS)

Garg, Vijay K.; Gaugler, Raymond E.

1995-01-01

An existing three dimensional Navier-Stokes code, modified to include film cooling considerations, has been used to study the effect of spanwise pitch of shower-head holes and coolant to mainstream mass flow ratio on the adiabatic effectiveness and heat transfer coefficient on a film-cooled turbine vane. The mainstream is akin to that under real engine conditions with stagnation temperature = 1900 K and stagnation pressure = 3 MPa. It is found that with the coolant to mainstream mass flow ratio fixed, reducing P, the spanwise pitch for shower-head holes, from 7.5 d to 3.0 d, where d is the hole diameter, increases the average effectiveness considerably over the blade surface. However, when P/d= 7.5, increasing the coolant mass flow increases the effectiveness on the pressure surface but reduces it on the suction surface due to coolant jet lift-off. For P/d = 4.5 or 3.0, such an anomaly does not occur within the range of coolant to mainstream mass flow ratios analyzed. In all cases, adiabatic effectiveness and heat transfer coefficient are highly three-dimensional.

Verification of a three-dimensional viscous flow analysis for a single stage compressor

NASA Astrophysics Data System (ADS)

Matsuoka, Akinori; Hashimoto, Keisuke; Nozaki, Osamu; Kikuchi, Kazuo; Fukuda, Masahiro; Tamura, Atsuhiro

1992-12-01

A transonic flowfield around rotor blades of a highly loaded single stage axial compressor was numerically analyzed by a three dimensional compressible Navier-Stokes equation code using Chakravarthy and Osher type total variation diminishing (TVD) scheme. A stage analysis which calculates both flowfields around inlet guide vane (IGV) and rotor blades simultaneously was carried out. Comparing with design values and experimental data, computed results show slight difference quantitatively. But the numerical calculation simulates well the pressure rise characteristics of the compressor and its flow pattern including strong shock surface.

Introduction to boundary-layer theory. [viscous friction loss calculation for turbine blade design

NASA Technical Reports Server (NTRS)

Mcnally, W. D.

1973-01-01

The pressure ratio across a turbine provides a certain amount of ideal energy that is available to the turbine for producing work. The portion of the ideal energy that is not converted to work is considered to be a loss. One of the more important and difficult aspects of turbine design is the prediction of the losses. The primary cause of losses is the boundary layer that develops on the blade and end wall surfaces. Boundary-layer theory is used to calculate the parameters needed to estimate viscous (friction) losses.

The Effect of Wake Passing on Turbine Blade Film Cooling

NASA Technical Reports Server (NTRS)

Heidmann, James David

1996-01-01

The effect of upstream blade row wake passing on the showerhead film cooling performance of a downstream turbine blade has been investigated through a combination of experimental and computational studies. The experiments were performed in a steady-flow annular turbine cascade facility equipped with an upstream rotating row of cylindrical rods to produce a periodic wake field similar to that found in an actual turbine. Spanwise, chordwise, and temporal resolution of the blade surface temperature were achieved through the use of an array of nickel thin-film surface gauges covering one unit cell of showerhead film hole pattern. Film effectiveness and Nusselt number values were determined for a test matrix of various injectants, injectant blowing ratios, and wake Strouhal numbers. Results indicated a demonstratable reduction in film effectiveness with increasing Strouhal number, as well as the expected increase in film effectiveness with blowing ratio. An equation was developed to correlate the span-average film effectiveness data. The primary effect of wake unsteadiness was found to be correlated well by a chordwise-constant decrement of 0.094-St. Measurable spanwise film effectiveness variations were found near the showerhead region, but meaningful unsteady variations and downstream spanwise variations were not found. Nusselt numbers were less sensitive to wake and injection changes. Computations were performed using a three-dimensional turbulent Navier-Stokes code which was modified to model wake passing and film cooling. Unsteady computations were found to agree well with steady computations provided the proper time-average blowing ratio and pressure/suction surface flow split are matched. The remaining differences were isolated to be due to the enhanced mixing in the unsteady solution caused by the wake sweeping normally on the pressure surface. Steady computations were found to be in excellent agreement with experimental Nusselt numbers, but to overpredict experimental film effectiveness values. This is likely due to the inability to match actual hole exit velocity profiles and the absence of a credible turbulence model for film cooling.

Performance of Single-Stage Turbine of Mark 25 Torpedo Power Plant with Two Special Nozzles. III; Efficiency with Standard Rotor Blades

NASA Technical Reports Server (NTRS)

Schum, Harold J.; Whitney, Warren J.

1949-01-01

A Mark 25 torpedo power plant modified to operate as a single-stage turbine was investigated to determine the performance with two nozzle designs and a standard first-stage rotor having 0.40-inch blades with a 17O met-air angle. Both nozzles had smaller port cross-sectional areas than those nozzles of similar design, which were previously investigated. The performance of the two nozzles was compared on the basis of blade, rotor, and brake efficiencies as a function of blade-jet speed ratio for pressure ratios of 8, 15 (design), and 20. At pressure ratios of 15 and 20, the blade efficiency obtained with the nozzle having circular passages (K) was higher than that obtained with the nozzle having rectangular passages (J). At a pressure ratio of 8, the efficiencies obtained with the two nozzles were comparable for blade-jet speed ratios of less than 0.260. For blade-jet speed ratios exceeding this value, nozzle K yielded slightly higher efficiencies. The maximum blade efficiency of 0.569 was obtained with nozzle K at a pressure ratio of 8 and a blade-jet speed ratio of 0.295. At design speed and pressure ratio, nozzle K yielded a maximum blade efficiency of 0.534, an increase of 0.031 over that obtained with nozzle J. When the blade efficiencies of the two nozzles were compared with those of four other nozzles previously investigated, the maximum difference for the six nozzles with this rotor was 0.050. From, this comparison, no specific effect of nozzles size or shape on over-all performance was discernible.

Theoretical Calculation of Viscous-Inviscid Transonic Flows.

DTIC Science & Technology

1980-08-01

Taylor Naval Ship Research and Development Center Aviation and Surface Effects Department / (See reverse side) Bethesda, Maryland 20084 ! CONTROLLING...Interactions ... .......... ... 46 18 -ffect of Boundary Layer on Blade Surface Pressures in a Transonic Fan Rotor Tip Section Cascade...complicated by the viscous effect . The strong viscous-inviscid interaction caused by the shock wave thickens the boundary layer rapidly, and the flow eventually

Part 1 - Experimental study of the pressure fluctuations on propeller turbine runner blades during steady-state operation

NASA Astrophysics Data System (ADS)

Houde, S.; Fraser, R.; Ciocan, G. D.; Deschênes, C.

2012-11-01

A good evaluation of the unsteady pressure field on hydraulic turbine blades is critical in evaluating the turbine lifespan and its maintenance schedule. Low-head turbines such as Kaplan and Propeller, using a relatively low number of blades supported only at the hub, may also undergo significant deflections at the blade tips which will lead to higher amplitude vibration compared to Francis turbines. Furthermore, the precise evaluation of the unsteady pressure distribution on low-head turbines is still a challenge for computational fluid dynamics (CFD). Within the framework of an international research consortium on low-head turbines, a research project was instigated at the Hydraulic Machines Laboratory in Laval University (LAMH) to perform experimental measurements of the unsteady pressure field on propeller turbine model runner blades. The main objective of the project was to measure the pressure fluctuations on a wide band of frequencies, both in a blade-to-blade channel and on the pressure and suction side of the same blade, to provide validation data for CFD computations. To do so, a 32 channels telemetric data transmission system was used to extract the signal of 31 pressure transducers and two strain gages from the rotating part at an acquisition frequency of 5 KHz. The miniature piezoelectric pressure transducers were placed on two adjacent runner blades according to an estimated pressure distribution coming from flow simulations. Two suction sides and one pressure side were instrumented. The strain gages were mounted in full-bridge on both pressure and suction sides to measure the blade span wise deflection. In order to provide boundary conditions for flow simulations, the test bench conditions during the measurements were acquired. The measurements were made in different operating conditions ranging from part load, where a cavitating vortex occurs, to full load under different heads. The results enabled the identification and the quantification of the major known sources of pressure fluctuation as well as some unexpected ones. The paper first presents the experimental methodology discussing relevant topics such as telemetric system setup, transducers calibration and errors analysis. The main results are then presented to illustrate the relative amplitude of the main source of pressure fluctuations under different operating conditions. The discussion and conclusion addresses the important observations stemming from the data analysis and illustrates that most of the results can be correlated with the known behavior of hydraulic turbines while some require further investigation.

Main rotor-body action for virtual blades model

NASA Astrophysics Data System (ADS)

Kusyumov, Alexander; Kusyumov, Sergey; Mikhailov, Sergey; Romanova, Elena; Phayzullin, Konstantin; Lopatin, Evgeny; Barakos, G.

2018-06-01

This research aims to investigate a virtual blade model and assess rotor influence on helicopter fuselage aerodynamics. The rotor disk is discretized in the azimuthal direction, and a time-varied pressure jump is applied in regions occupied by the blades. To obtain the pressure jump, an actuator disk is employed using uniform and non-uniform blade load distribution, based on momentum theory.

FORTRAN program for calculating velocities and streamlines on the hub-shroud mid-channel flow surface of an axial- or mixed-flow turbomachine. 1: User's manual

NASA Technical Reports Server (NTRS)

Katsanis, T.

1973-01-01

A FORTRAN 4 computer program has been developed that obtains a subsonic or shock-free transonic flow solution on the hub-shroud mid-channel flow surface of a turbomachine. The blade row may be fixed or rotating, and may be twisted and leaned. Flow may be axial or mixed, up to 45 deg from axial. Upstream and downstream flow variables may vary from hub to shroud, and provision is made to correct for loss of stagnation pressure. The results include velocities, streamlines, and flow angles on the flow surface; and approximate blade surface velocities. Subsonic solutions are obtained by a finite-difference stream-function solution. Transonic solutions are obtained by a velocity-gradient method, using information from a finite-difference stream-function solution at a reduced mass flow.

FORTRAN program for calculating velocities and streamlines on the hub-shroud mid-channel flow surface of an axial-or mixed-flow turbomachine. 2: Programmer's manual

NASA Technical Reports Server (NTRS)

Katsanis, T.; Mcnally, W. D.

1974-01-01

A FORTRAN-IV computer program, MERIDL, has been developed that obtains a subsonic or shock-free transonic flow solution on the hub-shroud mid-channel flow surface of a turbomachine. The blade row may be fixed or rotating and may be twisted and leaned. Flow may be axial or mixed, up to 45 deg from axial. Upstream and downstream flow variables can vary from hub to shroud, and provision is made to correct for loss of stagnation pressure. The results include velocities, streamlines, and flow angles on the flow surface and approximate blade surface velocities. Subsonic solutions are obtained by a finite-difference stream-function solution. Transonic solutions are obtained by a velocity-gradient method, using information from a finite-difference stream-function solution at a reduced mass flow.

High Temperature Ceramic Guide Vane Temperature and Pressure Distribution Calculation for Flow with Cooling Jets

NASA Technical Reports Server (NTRS)

Srivastava, Rakesh

2004-01-01

A ceramic guide vane has been designed and tested for operation under high temperature. Previous efforts have suggested that some cooling flow may be required to alleviate the high temperatures observed near the trailing edge region. The present report describes briefly a three-dimensional viscous analysis carried out to calculate the temperature and pressure distribution on the blade surface and in the flow path with a jet of cooling air exiting from the suction surface near the trailing edge region. The data for analysis was obtained from Dr. Craig Robinson. The surface temperature and pressure distribution along with a flowfield distribution is shown in the results. The surface distribution is also given in a tabular form at the end of the document.

Compact Assumption Applied to the Monopole Term of Farassat's Formulations

NASA Technical Reports Server (NTRS)

Lopes, Leonard V.

2015-01-01

Farassat's formulations provide an acoustic prediction at an observer location provided a source surface, including motion and flow conditions. This paper presents compact forms for the monopole term of several of Farassat's formulations. When the physical surface is elongated, such as the case of a high aspect ratio rotorcraft blade, compact forms can be derived which are shown to be a function of the blade cross sectional area by reducing the computation from a surface integral to a line integral. The compact forms of all formulations are applied to two example cases: a short span wing with constant airfoil cross section moving at three forward flight Mach numbers and a rotor at two advance ratios. Acoustic pressure time histories and power spectral densities of monopole noise predicted from the compact forms of all the formulations at several observer positions are shown to compare very closely to the predictions from their non-compact counterparts. A study on the influence of rotorcraft blade shape on the high frequency portion of the power spectral density shows that there is a direct correlation between the aspect ratio of the airfoil and the error incurred by using the compact form. Finally, a prediction of pressure gradient from the non-compact and compact forms of the thickness term of Formulation G1A shows that using the compact forms results in a 99.6% improvement in computation time, which will be critical when noise is incorporated into a design environment.

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.

Investigation of Positively Curved Blade in Compressor Cascade Based on Transition Model

NASA Astrophysics Data System (ADS)

Chen, Shaowen; Lan, Yunhe; Zhou, Zhihua; Wang, Songtao

2016-06-01

Experiment and numerical simulation of flow transition in a compressor cascade with positively curved blade is carried out in a low speed. In the experimental investigation, the outlet aerodynamic parameters are measured using a five-hole aerodynamic probe, and an ink-trace flow visualization is applied to the cascade surface. The effects of transition flow on the boundary layer development, three-dimensional flow separation and aerodynamic performance are studied. The feasibility of a commercial computational fluid dynamic code is validated and the numerical results show a good agreement with experimental data. The blade-positive curving intensifies the radial force from the endwalls to the mid-span near the suction surface, which leads to the smaller scope of the intermittent region, the lesser extents of turbulence intensity and the shorter radial height of the separation bubble near the endwalls, but has little influence on the flow near the mid-span. The large passage vortex is divided into two smaller shedding vortexes under the impact of the radial pressure gradient due to the positively curved blade. The new concentrated shedding vortex results in an increase in the turbulence intensity and secondary flow loss of the corresponding region.

MERIDL- VELOCITIES AND STREAMLINES ON THE HUB-SHROUD MIDCHANNEL STREAM SURFACE OF AN AXIAL, RADIAL, OR MIXED FLOW TURBOMACHINE OR ANNULAR DUCT

NASA Technical Reports Server (NTRS)

Katsanis, T.

1994-01-01

This computer program was developed for calculating the subsonic or transonic flow on the hub-shroud mid-channel stream surface of a single blade row of a turbomachine. The design and analysis of blades for compressors and turbines ideally requires methods for analyzing unsteady, three-dimensional, turbulent viscous flow through a turbomachine. Since an exact solution is impossible at present, solutions on two-dimensional surfaces are calculated to obtain a quasi-three dimensional solution. When three-dimensional effects are important, significant information can be obtained from a solution on a cross-sectional surface of the passage normal to the flow. With this program, a solution to the equations of flow on the meridional surface can be carried out. This solution is chosen when the turbomachine under consideration has significant variation in flow properties in the hubshroud direction, especially when input is needed for use in blade-to-blade calculations. The program can also perform flow calculations for annular ducts without blades. This program should prove very useful in the design and analysis of any turbomachine. This program calculates a solution for two-dimensional, adiabatic shockfree flow. The flow must be essentially subsonic, but there may be local areas of supersonic flow. To obtain the solution, this program uses both the finite difference and the quasi-orthogonal (velocity gradient) methods combined in a way that takes maximum advantage of both. The finite-difference method solves a finite-difference equation along the meridional stream surface in a very efficient manner but is limited to subsonic velocities. This approach must be used in cases where the blade aspect ratios are above one, cases where the passage is curved, and cases with low hub-tip-ratio blades. The quasi-orthogonal method solves the velocity gradient equation on the meridional surface and is used if it is necessary to extend the range of solutions into the transonic regime. In general the blade row may be fixed or rotating and the blades may be twisted and leaned. The flow may be axial, radial, or mixed. The upstream and downstream flow conditions can vary from hub to shroud with provisions made for an approximate correction for loss of stagnation pressure. Also, viscous forces are neglected along solution mesh lines running from hub to tip. The capabilities of this program include handling of nonaxial flows without restriction, annular ducts without blades, and specified streamwise loss distributions. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 360 computer with a central memory requirement of approximately 700K of 8 bit bytes. This core requirement can be reduced depending on the size of the problem and the desired solution accuracy. This program was developed in 1977.

Surface modeling method for aircraft engine blades by using speckle patterns based on the virtual stereo vision system

NASA Astrophysics Data System (ADS)

Yu, Zhijing; Ma, Kai; Wang, Zhijun; Wu, Jun; Wang, Tao; Zhuge, Jingchang

2018-03-01

A blade is one of the most important components of an aircraft engine. Due to its high manufacturing costs, it is indispensable to come up with methods for repairing damaged blades. In order to obtain a surface model of the blades, this paper proposes a modeling method by using speckle patterns based on the virtual stereo vision system. Firstly, blades are sprayed evenly creating random speckle patterns and point clouds from blade surfaces can be calculated by using speckle patterns based on the virtual stereo vision system. Secondly, boundary points are obtained in the way of varied step lengths according to curvature and are fitted to get a blade surface envelope with a cubic B-spline curve. Finally, the surface model of blades is established with the envelope curves and the point clouds. Experimental results show that the surface model of aircraft engine blades is fair and accurate.

Measurement of the Coolant Channel Temperatures and Pressures of a Cooled Radial-Inflow Turbine

NASA Technical Reports Server (NTRS)

Dicicco, L. Danielle; Nowlin, Brent C.; Tirres, Lizet

1994-01-01

Instrumentation has been installed on the surface of a cooled radial-inflow turbine. Thermocouples and miniature integrated sensor pressure transducers were installed to measure steady state coolant temperatures, blade wall temperatures, and coolant pressures. These measurements will eventually be used to determine the heat transfer characteristics of the rotor. This paper will describe the procedures used to install and calibrate the instrumentation and the testing methods followed. A limited amount of data will compare the measured values to the predicted values.

Knowledge-based system for detailed blade design of turbines

NASA Astrophysics Data System (ADS)

Goel, Sanjay; Lamson, Scott

1994-03-01

A design optimization methodology that couples optimization techniques to CFD analysis for design of airfoils is presented. This technique optimizes 2D airfoil sections of a blade by minimizing the deviation of the actual Mach number distribution on the blade surface from a smooth fit of the distribution. The airfoil is not reverse engineered by specification of a precise distribution of the desired Mach number plot, only general desired characteristics of the distribution are specified for the design. Since the Mach number distribution is very complex, and cannot be conveniently represented by a single polynomial, it is partitioned into segments, each of which is characterized by a different order polynomial. The sum of the deviation of all the segments is minimized during optimization. To make intelligent changes to the airfoil geometry, it needs to be associated with features observed in the Mach number distribution. Associating the geometry parameters with independent features of the distribution is a fairly complex task. Also, for different optimization techniques to work efficiently the airfoil geometry needs to be parameterized into independent parameters, with enough degrees of freedom for adequate geometry manipulation. A high-pressure, low reaction steam turbine blade section was optimized using this methodology. The Mach number distribution was partitioned into pressure and suction surfaces and the suction surface distribution was further subdivided into leading edge, mid section and trailing edge sections. Two different airfoil representation schemes were used for defining the design variables of the optimization problem. The optimization was performed by using a combination of heuristic search and numerical optimization. The optimization results for the two schemes are discussed in the paper. The results are also compared to a manual design improvement study conducted independently by an experienced airfoil designer. The turbine blade optimization system (TBOS) is developed using the described methodology of coupling knowledge engineering with multiple search techniques for blade shape optimization. TBOS removes a major bottleneck in the design cycle by performing multiple design optimizations in parallel, and improves design quality at the same time. TBOS not only improves the design but also the designers' quality of work by taking the mundane repetitive task of design iterations away and leaving them more time for innovative design.

A Numerical Study of the Effect of Wake Passing on Turbine Blade Film Cooling

NASA Technical Reports Server (NTRS)

Heidmann, James D.

1995-01-01

Time-accurate and steady three-dimensional viscous turbulent numerical simulations were performed to study the effect of upstream blade wake passing unsteadiness on the performance of film cooling on a downstream axial turbine blade. The simulations modeled the blade as spanwise periodic and of infinite span. Both aerodynamic and heat transfer quantities were explored. A showerhead film cooling arrangement typical of modern gas turbine engines was employed. Showerhead cooling was studied because of its anticipated strong sensitivity to upstream flow fluctuations. The wake was modeled as a region of zero axial velocity on the upstream computational boundary which translated with each iteration. This model is compatible with a planned companion experiment in which the wakes will be produced by a rotating row of cylindrical rods upstream of an annular turbine cascade. It was determined that a steady solution with appropriate upstream swirl and stagnation pressure predicted the span-average film effectiveness quite well. The major difference is a 2 to 3 percent overprediction of span-average film effectiveness by the steady simulation on the pressure surface and in the showerhead region. Local overpredictions of up to 8 percent were observed in the showerhead region. These differences can be explained by the periodic relative lifting of the boundary layer and enhanced mixing in the unsteady simulations.

Comparison of CFD simulations to non-rotating MEXICO blades experiment in the LTT wind tunnel of TUDelft

NASA Astrophysics Data System (ADS)

Zhang, Ye; van Zuijlen, Alexander; van Bussel, Gerard

2014-06-01

In this paper, three dimensional flow over non-rotating MEXICO blades is simulated by CFD methods. The numerical results are compared with the latest MEXICO wind turbine blades measurements obtained in the low speed low turbulence (LTT) wind tunnel of Delft University of Technology. This study aims to validate CFD codes by using these experimental data measured in well controlled conditions. In order to avoid use of wind tunnel corrections, both the blades and the wind tunnel test section are modelled in the simulations. The ability of Menter's k - ω shear stress transport (SST) turbulence model is investigated at both attached flow and massively separated flow cases. Steady state Reynolds averaged Navier Stokes (RANS) equations are solved in these computations. The pressure distribution at three measured sections are compared under the conditions of different inflow velocities and a range of angles of attack. The comparison shows that at attached flow condition, good agreement can be obtained for all three airfoil sections. Even with massively separated flow, still fairly good pressure distribution comparison can be found for the DU and NACA airfoil sections, although the RISØ section shows poor comparison. At the near stall case, considerable deviations exists on the forward half part of the upper surface for all three sections.

Influence of Cleats-Surface Interaction on the Performance and Risk of Injury in Soccer: A Systematic Review

PubMed Central

Macedo, Rui; Montes, António Mesquita

2017-01-01

Objective To review the influence of cleats-surface interaction on the performance and risk of injury in soccer athletes. Design Systematic review. Data Sources Scopus, Web of science, PubMed, and B-on. Eligibility Criteria Full experimental and original papers, written in English that studied the influence of soccer cleats on sports performance and injury risk in artificial or natural grass. Results Twenty-three articles were included in this review: nine related to performance and fourteen to injury risk. On artificial grass, the soft ground model on dry and wet conditions and the turf model in wet conditions are related to worse performance. Compared to rounded studs, bladed ones improve performance during changes of directions in both natural and synthetic grass. Cleat models presenting better traction on the stance leg improve ball velocity while those presenting a homogeneous pressure across the foot promote better kicking accuracy. Bladed studs can be considered less secure by increasing plantar pressure on lateral border. The turf model decrease peak plantar pressure compared to other studded models. Conclusion The soft ground model provides lower performance especially on artificial grass, while the turf model provides a high protective effect in both fields. PMID:28684897

Influence of Cleats-Surface Interaction on the Performance and Risk of Injury in Soccer: A Systematic Review.

PubMed

Silva, Diogo C F; Santos, Rubim; Vilas-Boas, João Paulo; Macedo, Rui; Montes, António Mesquita; Sousa, Andreia S P

2017-01-01

To review the influence of cleats-surface interaction on the performance and risk of injury in soccer athletes. Systematic review. Scopus, Web of science, PubMed, and B-on. Full experimental and original papers, written in English that studied the influence of soccer cleats on sports performance and injury risk in artificial or natural grass. Twenty-three articles were included in this review: nine related to performance and fourteen to injury risk. On artificial grass, the soft ground model on dry and wet conditions and the turf model in wet conditions are related to worse performance. Compared to rounded studs, bladed ones improve performance during changes of directions in both natural and synthetic grass. Cleat models presenting better traction on the stance leg improve ball velocity while those presenting a homogeneous pressure across the foot promote better kicking accuracy. Bladed studs can be considered less secure by increasing plantar pressure on lateral border. The turf model decrease peak plantar pressure compared to other studded models. The soft ground model provides lower performance especially on artificial grass, while the turf model provides a high protective effect in both fields.

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.

Computational methods in the prediction of advanced subsonic and supersonic propeller induced noise: ASSPIN users' manual

NASA Technical Reports Server (NTRS)

Dunn, M. H.; Tarkenton, G. M.

1992-01-01

This document describes the computational aspects of propeller noise prediction in the time domain and the use of high speed propeller noise prediction program ASSPIN (Advanced Subsonic and Supersonic Propeller Induced Noise). These formulations are valid in both the near and far fields. Two formulations are utilized by ASSPIN: (1) one is used for subsonic portions of the propeller blade; and (2) the second is used for transonic and supersonic regions on the blade. Switching between the two formulations is done automatically. ASSPIN incorporates advanced blade geometry and surface pressure modelling, adaptive observer time grid strategies, and contains enhanced numerical algorithms that result in reduced computational time. In addition, the ability to treat the nonaxial inflow case has been included.

Numerical simulation on the cavitation of waterjet propulsion pump

NASA Astrophysics Data System (ADS)

Xia, C. Z.; Cheng, L.; Shang, Y. N.; Zhou, J. R.; Yang, F.; Jin, Y.

2016-05-01

Waterjet propulsion system is widely used in high speed vessels with advantages of simple transmission mechanism, low noise underwater and good manoeuvrability. Compared with the propeller, waterjet propulsion can be used flow stamping to increasing cavitation resistance at high speed. But under certain conditions, such as low ship speed or high ship speed, cavitation problem still exists. If water-jet propulsion pump is run in cavitation condition for a long time, then the cavitation will cause a great deal of noise CFD is applied to analysis and predict the process of production and development of cavitation in waterjet propulsion pump. Based on the cavitation model of Zwart-Gerber-Belamri and a mixture of homogeneous flow model, commercial CFD software CFX was taken for characteristics of cavitation under the three operating conditions. Commercial software ANSYS 14.0 is used to build entity model, mesh and numerical simulation. The grid independence analysis determine the grid number of mixed flow pump model is about 1.6 million and the grid number of water-jet pump system unit is about 2.7 million. The cavitation characteristics of waterjet pump under three operating conditions are studied. The results show that the cavitation development trend is similar design and small rate of flow condition .Under the design conditions Cavitation bubbles are mainly gathered in suction surface of blade near the inlet side of the hub under the primary stage, and gradually extended to the water side in the direction of the rim with the loss of the inlet total pressure. Cavitation appears in hub before the blade rim, but the maximum value of gas content in blade rim is bigger than that in hub. Under large flow conditions, bubble along the direction of wheel hub extends to the rim gradually. Cavitation is found in the pressure surface of blade near the hub region under the critical point of cavitation nearby. When NPSHa is lower than critical point, the area covering by bubbles is about 40% in the suction surface of blade. It means that the critical point of cavitation of pump system is not the accrue point of install cavitation but cavitation has been developed to a certain stage.

Pressure Sensitive Paint Measurements on 15% Scale Rotor Blades in Hover

NASA Technical Reports Server (NTRS)

Wong, Oliver D.; Watkins, Anthony Neal; Ingram, JoAnne L.

2005-01-01

This paper describes a proof of concept test to examine the feasibility of using pressure sensitive paint (PSP) to measure the pressure distributions on a rotor in hover. The test apparatus consisted of the US Army 2-meter Rotor Test Stand (2MRTS) and 15% scale swept tip rotor blades. Two camera/rotor separations were examined: 0.76 and 1.35 radii. The outer 15% of each blade was painted with PSP. Intensity and lifetime based PSP measurement techniques were attempted. Data were collected from all blades at thrust coefficients ranging from 0.004 to 0.009.

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.

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.

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.

Comparison of Calculated and Experimental Temperatures and Coolant Pressure Losses for a Cascade of Small Air-Cooled Turbine Rotor Blades

NASA Technical Reports Server (NTRS)

Stepka, Francis S

1958-01-01

Average spanwise blade temperatures and cooling-air pressure losses through a small (1.4-in, span, 0.7-in, chord) air-cooled turbine blade were calculated and are compared with experimental nonrotating cascade data. Two methods of calculating the blade spanwise metal temperature distributions are presented. The method which considered the effect of the length-to-diameter ratio of the coolant passage on the blade-to-coolant heat-transfer coefficient and assumed constant coolant properties based on the coolant bulk temperature gave the best agreement with experimental data. The agreement obtained was within 3 percent at the midspan and tip regions of the blade. At the root region of the blade, the agreement was within 3 percent for coolant flows within the turbulent flow regime and within 10 percent for coolant flows in the laminar regime. The calculated and measured cooling-air pressure losses through the blade agreed within 5 percent. Calculated spanwise blade temperatures for assumed turboprop engine operating conditions of 2000 F turbine-inlet gas temperature and flight conditions of 300 knots at a 30,000-foot altitude agreed well with those obtained by the extrapolation of correlated experimental data of a static cascade investigation of these blades.

Numerical investigation of the effects of ITD length on low pressure nozzle

NASA Astrophysics Data System (ADS)

Liu, Guang; Liu, Jun; Liu, Hongrui; Wang, Pei; Du, Qiang

2017-06-01

The advantage of high efficiency, low SFC (Specific Fuel Consumption), ultra-high bypass ratio turbofan engine attracts more and more attention in modern commercial engine. The intermediate turbine duct (ITD), which connects high pressure turbine (HPT) with low pressure turbine (LPT), has a critical impact on the overall performance of turbine by guiding flow coming from HPT to LPT without too much loss. Therefore, it becomes more and more urgent to master the technique of designing aggressive, even super-aggressive ITD. Much more concerns have been concentrated on the duct. However, in order to further improve turbine, LPT nozzle is arranged into ITD to shorten low pressure axle. With such design concept, it is obvious that LPT nozzle flow field is easily influenced by upstream duct structure, but receives much less interests on the contrary. In this paper, numerical method is used to investigate the effects of length of ITD with upstream swirl blades on LPT nozzle. Nine models with the same swirl and nozzle blades, while the length of ITD is the only parameter to be changed, will be discussed. Finally, several conclusions and advices for designers are summarized. After changing axial length of ducts, inlet and outlet flow field of nozzle differs, correspondingly. On the other hand, the shearing stress on nozzle blade (suction and pressure) surface presents individual feature under various inlet flow. In addition to that, "Clocking-like effect" is described in this paper, which will contribute much to the pressure loss and should be paid enough attention.

A flight investigation of blade section aerodynamics for a helicopter main rotor having NLR-1T airfoil sections

NASA Technical Reports Server (NTRS)

Morris, C. E. K., Jr.; Stevens, D. D.; Tomaine, R. L.

1980-01-01

A flight investigation was conducted using a teetering-rotor AH-1G helicopter to obtain data on the aerodynamic behavior of main-rotor blades with the NLR-1T blade section. The data system recorded blade-section aerodynamic pressures at 90 percent rotor radius as well as vehicle flight state, performance, and loads. The test envelope included hover, forward flight, and collective-fixed maneuvers. Data were obtained on apparent blade-vortex interactions, negative lift on the advancing blade in high-speed flight and wake interactions in hover. In many cases, good agreement was achieved between chordwise pressure distributions predicted by airfoil theory and flight data with no apparent indications of blade-vortex interactions.

Progress in the utilization of an oxide-dispersion-strengthened alloy for small engine turbine blades

NASA Technical Reports Server (NTRS)

Beatty, T. G.; Millan, P. P.

1984-01-01

The conventional means of improving gas turbine engine performance typically involves increasing the turbine inlet temperature; however, at these higher operational temperatures the high pressure turbine blades require air-cooling to maintain durability. Air-cooling imposes design, material, and economic constraints not only on the turbine blades but also on engine performance. The use of uncooled turbine blades at increased operating temperatures can offer significantly improved performance in small gas turbine engines. A program to demonstrate uncooled MA6000 high pressure turbine blades in a GTEC TFE731 turbofan engine is being conducted. The project goals include demonstration of the advantages of using uncooled MA6000 turbine blades as compared with cast directionally solidified MAR-M 247 blades.

Heated Surface Temperatures Measured by Infrared Detector in a Cascade Environment

NASA Technical Reports Server (NTRS)

Boyle, Robert J.

2002-01-01

Investigators have used infrared devices to accurately measure heated surface temperatures. Several of these applications have been for turbine heat transfer studies involving film cooling and surface roughness, typically, these measurements use an infrared camera positioned externally to the test section. In cascade studies, where several blades are used to ensure periodic flow, adjacent blades block the externally positioned camera's views of the test blade. To obtain a more complete mapping of the surface temperatures, researchers at the NASA Glenn Research Center fabricated a probe with an infrared detector to sense the blade temperatures. The probe size was kept small to minimize the flow disturbance. By traversing and rotating the probe, using the same approach as for total pressure surveys, one can find the blade surface temperatures. Probe mounted infrared detectors are appropriate for measuring surface temperatures where an externally positioned infrared camera is unable to completely view the test object. This probe consists of a 8-mm gallium arsenide (GaAs) lens mounted in front of a mercury-cadmium-zinc-tellurium (HgCdZnTe) detector. This type of photovoltaic detector was chosen because of its high sensitivity to temperature when the detector is uncooled. The particular application is for relatively low surface temperatures, typically ambient to 100 C. This requires a detector sensitive at long wavelengths. The detector is a commercial product enclosed in a 9-mm-diameter package. The GaAs lens material was chosen because of its glass-like hardness and its good long-wavelength transmission characteristics. When assembled, the 6.4-mm probe stem is held in the traversing actuator. Since the entire probe is above the measurement plane, the flow field disturbance in the measurement plane is minimized. This particular probe body is somewhat wider than necessary, because it was designed to have replaceable detectors and lenses. The signal for the detector is fed through the hollow probe body. The detector's signal goes to an externally mounted preamplifier. The detector assembly, along with a preamplifier, is calibrated as a function of the surface temperature for various detector temperatures. The output voltage is a function of both the detector and object temperatures.

Broadband Noise Reduction of a Low-Speed Fan Noise Using Trailing Edge Blowing

NASA Technical Reports Server (NTRS)

Sutliff, Daniel L.

2005-01-01

An experimental proof-of-concept test was conducted to demonstrate reduction of rotor-stator interaction noise through the use of rotor-trailing edge blowing. The velocity deficit from the viscous wake of the rotor blades was reduced by injecting air into the wake from a continuous trailing edge slot. Hollow blades with interior guide vanes create flow channels through which externally supplied air flows from the blade root to the trailing edge. A previous paper documented the substantial tonal reductions of this Trailing Edge Rotor Blowing (TERB) fan. This report documents the broadband characteristics of TERB. The Active Noise Control Fan (ANCF), located at the NASA Glenn Research Center, was used as the proof-of-concept test bed. Two-component hotwire data behind the rotor, unsteady surface pressures on the stator vane, and farfield directivity acoustic data were acquired at blowing rates of 1.1, 1.5, and 1.8 percent of the total fan mass flow. The results indicate a substantial reduction in the rotor wake turbulent velocity and in the stator vane unsteady surface pressures. Based on the physics of the noise generation, these indirect measurements indicate the prospect of broadband noise reduction. However, since the broadband noise generated by the ANCF is rotor-dominated, any change in the rotor-stator interaction broadband noise levels is barely distinguishable in the farfield measurements.

An investigation of the flow characteristics in the blade endwall corner region

NASA Technical Reports Server (NTRS)

Hazarika, Birinchi K.; Raj, Rishi S.

1987-01-01

Studies were undertaken to determine the structure of the flow in the blade end wall corner region simulated by attaching two uncambered airfoils on either side of a flat plate with a semicircular leading edge. Detailed measurements of the corner flow were obtained with conventional pressure probes, hot wire anemometry, and flow visualization. The mean velocity profiles and six components of the Reynolds stress tensor were obtained with an inclined single sensor hot wire probe whereas power spectra were obtained with a single sensor oriented normal to the flow. Three streamwise vortices were identified based on the surface streamlines, distortion of total pressure profiles, and variation of mean velocity components in the corner. A horseshoe vortex formed near the leading edge of the airfoil. Within a short distance downstream, a corner vortex was detected between the horseshoe vortex and the surfaces forming the corner. A third vortex was formed at the rear portion of the corner between the corner vortex and the surface of the flat plate. Turbulent shear stress and production of turbulence are negligibly small. A region of negative turbulent shear stress was also observed near the region of low turbulence intensity from the vicinity of the flat plate.

Unified aeroacoustics analysis for high speed turboprop aerodynamics and noise. Volume 3: Application of theory for blade loading, wakes, noise, and wing shielding

NASA Technical Reports Server (NTRS)

Hanson, D. B.; Mccolgan, C. J.; Ladden, R. M.; Klatte, R. J.

1991-01-01

Results of the program for the generation of a computer prediction code for noise of advanced single rotation, turboprops (prop-fans) such as the SR3 model are presented. The code is based on a linearized theory developed at Hamilton Standard in which aerodynamics and acoustics are treated as a unified process. Both steady and unsteady blade loading are treated. Capabilities include prediction of steady airload distributions and associated aerodynamic performance, unsteady blade pressure response to gust interaction or blade vibration, noise fields associated with thickness and steady and unsteady loading, and wake velocity fields associated with steady loading. The code was developed on the Hamilton Standard IBM computer and has now been installed on the Cray XMP at NASA-Lewis. The work had its genesis in the frequency domain acoustic theory developed at Hamilton Standard in the late 1970s. It was found that the method used for near field noise predictions could be adapted as a lifting surface theory for aerodynamic work via the pressure potential technique that was used for both wings and ducted turbomachinery. In the first realization of the theory for propellers, the blade loading was represented in a quasi-vortex lattice form. This was upgraded to true lifting surface loading. Originally, it was believed that a purely linear approach for both aerodynamics and noise would be adequate. However, two sources of nonlinearity in the steady aerodynamics became apparent and were found to be a significant factor at takeoff conditions. The first is related to the fact that the steady axial induced velocity may be of the same order of magnitude as the flight speed and the second is the formation of leading edge vortices which increases lift and redistribute loading. Discovery and properties of prop-fan leading edge vortices were reported in two papers. The Unified AeroAcoustic Program (UAAP) capabilites are demonstrated and the theory verified by comparison with the predictions with data from tests at NASA-Lewis. Steady aerodyanmic performance, unsteady blade loading, wakes, noise, and wing and boundary layer shielding are examined.

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

Effects of wake and shock passing on the heat transfer to a film cooled transonic turbine blade

NASA Astrophysics Data System (ADS)

Rigby, M. J.

An attempt is made to further the understanding of film cooling process in an engine environment. The environment in a gas turbine is unsteady. A source of unsteadiness, the cutting of nozzle guide vane (NGV) wakes and shock waves by the rotor, was modeled experimentally. The influence of the unsteady wakes and shock waves on the heat transfer to a film cooled rotor blade was studied for five film cooling configurations using a rotating bar apparatus in front of a 2-D cascade. Heat transfer measurements were made using thin film gauges placed at the mid-span of the test blade. Schlieren photography was used to study the behavior of the coolant film and the movement of the unsteady shock waves and wakes. The effect of simulated NGV wake passing observed on the uncooled airfoil is to promote an intermittent transition of the suction surface. The effect of the wake on the turbulent pressure surface is small. With injection on the suction surface, the film acts as a boundary layer trip which offsets the rise in heat transfer due to the wake. The simulated NGV trailing edge shock wave had a dramatic effect on the suction surface heat transfer.

Rotational Augmentation on a 2.3 MW Rotor Blade with Thick Flatback Airfoil Cross-Sections: Preprint

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

Schreck, S.; Fingersh, L.; Siegel, K.

2013-01-01

Rotational augmentation was analyzed for a 2.3 MW wind turbine, which was equipped with thick flatback airfoils at inboard radial locations and extensively instrumented for acquisition of time varying surface pressures. Mean aerodynamic force and surface pressure data were extracted from an extensive field test database, subject to stringent criteria for wind inflow and turbine operating conditions. Analyses of these data showed pronounced amplification of aerodynamic forces and significant enhancements to surface pressures in response to rotational influences, relative to two-dimensional, stationary conditions. Rotational augmentation occurrence and intensity in the current effort was found to be consistent with that observedmore » in previous research. Notably, elevated airfoil thickness and flatback design did not impede rotational augmentation.« less

High-speed propeller noise predictions - Effects of boundary conditions used in blade loading calculations

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Clark, B. J.; Groeneweg, J. F.

1987-01-01

The acoustics of an advanced single rotation SR-3 propeller at cruise conditions are studied employing a time-domain approach. The study evaluates the acoustic significance of the differences in blade pressures computed using nonreflecting rather than hard wall boundary conditions in the three-dimensional Euler code solution. The directivities of the harmonics of the blade passing frequency tone and the effects of chordwise loading on tone directivity are examined. The results show that the maximum difference in the computed sound pressure levels due to the use of blade pressure distributions obtained with the nonreflecting rather than the hard wall boundary conditions is about 1.5 dB. The blade passing frequency tone directivity obtained in the present study shows good agreement with jetstar flight data.

High-speed propeller noise predictions: Effects of boundary conditions used in blade loading calculations

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Clark, B. J.; Groeneweg, J. F.

1987-01-01

The acoustics of an advanced single rotation SR-3 propeller at cruise conditions are studied employing a time-domain approach. The study evaluates the acoustic significance of the differences in blade pressures computed using nonreflecting rather than hard wall boundary conditions in the three-dimensional Euler code solution. The directivities of the harmonics of the blade passing frequency tone and the effects of chordwise loading on tone directivity are examined. The results show that the maximum difference in the computed sound pressure levels due to the use of blade pressure distributions obtained with the nonreflecting rather than the hard wall boundary conditions is about 1.5 dB. The blade passing frequency tone directivity obtained in the present study shows good agreement with jetstar flight data.

Investigation of the unsteady pressure distribution on the blades of an axial flow fan

NASA Technical Reports Server (NTRS)

Henderson, R. E.; Franke, G. F.

1978-01-01

The unsteady response of a stator blade caused by the interaction of the stator with the wakes of an upstream rotor was investigated. Unsteady pressure distributions were measured using a blade instrumented with a series miniature pressure transducers. The influence of several geometrical and flow parameters - rotor/stator spacing, stator solidity and stator incidence angle - were studied to determine the unsteady response of the stator to these parameters. A major influence on the stator unsteady response is due to the stator solidity. At high solidities the blade-to-blade interference has a larger contribution. While the range of rotor/stator spacings investigated had a minor influence, the effect of stator incidence angle is significant. The data indicate the existence of an optimum positive incidence which minimizes the unsteady response.

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.

LDV measurement of boundary layer on rotating blade surface in wind tunnel

NASA Astrophysics Data System (ADS)

Maeda, Takao; Kamada, Yasunari; Murata, Junsuke; Suzuki, Daiki; Kaga, Norimitsu; Kagisaki, Yosuke

2014-12-01

Wind turbines generate electricity due to extracting energy from the wind. The rotor aerodynamics strongly depends on the flow around blade. The surface flow on the rotating blade affects the sectional performance. The wind turbine surface flow has span-wise component due to span-wise change of airfoil section, chord length, twisted angle of blade and centrifugal force on the flow. These span-wise flow changes the boundary layer on the rotating blade and the sectional performance. Hence, the thorough understanding of blade surface flow is important to improve the rotor performance. For the purpose of clarification of the flow behaviour around the rotor blade, the velocity in the boundary layer on rotating blade surface of an experimental HAWT was measured in a wind tunnel. The velocity measurement on the blade surface was carried out by a laser Doppler velocimeter (LDV). As the results of the measurement, characteristics of surface flow are clarified. In optimum tip speed operation, the surface flow on leading edge and r/R=0.3 have large span-wise velocity which reaches 20% of sectional inflow velocity. The surface flow inboard have three dimensional flow patterns. On the other hand, the flow outboard is almost two dimensional in cross sectional plane.

Controlled laboratory testing of arthroscopic shaver systems: do blades, contact pressure, and speed influence their performance?

PubMed

Wieser, Karl; Erschbamer, Matthias; Neuhofer, Stefan; Ek, Eugene T; Gerber, Christian; Meyer, Dominik C

2012-10-01

The purposes of this study were (1) to establish a reproducible, standardized testing protocol to evaluate the performance of different shaver systems and blades in a controlled, laboratory setting, and (2) to determine the optimal use of different blades with respect to the influence of contact pressure and speed of blade rotation. A holding device was developed for reproducible testing of soft-tissue (tendon and meniscal) resection performance in a submerged environment, after loading of the shaver with interchangeable weights. The Karl Storz Powershaver S2 (Karl Storz, Tuttlingen, Germany), the Stryker Power Shaver System (Stryker, Kalamazoo, MI), and the Dyonics Power Shaver System (Smith & Nephew, Andover, MA) were tested, with different 5.5-mm shaver blades and varied contact pressure and rotation speed. For quality testing, serrated shaver blades were evaluated at 40× image magnification. Overall, more than 150 test cycles were performed. No significant differences could be detected between comparable blade types from different manufacturers. Shavers with a serrated inner blade and smooth outer blade performed significantly better than the standard smooth resectors (P < .001). Teeth on the outer layer of the blade did not lead to any further improvement of resection (P = .482). Optimal contact pressure ranged between 6 and 8 N, and optimal speed was found to be 2,000 to 2,500 rpm. Minimal blunting of the shaver blades occurred after soft-tissue resection; however, with bone resection, progressive blunting of the shaver blades was observed. Arthroscopic shavers can be tested in a controlled setting. The performance of the tested shaver types appears to be fairly independent of the manufacturer. For tendon resection, a smooth outer blade and serrated inner blade were optimal. This is one of the first established independent and quantitative assessments of arthroscopic shaver systems and blades. We believe that this study will assist the surgeon in choosing the optimal tool for the desired effect. Copyright © 2012 Arthroscopy Association of North America. Published by Elsevier Inc. All rights reserved.

A proposed through-flow inverse method for the design of mixed-flow pumps

NASA Technical Reports Server (NTRS)

Borges, Joao Eduardo

1991-01-01

A through-flow (hub-to-shroud) truly inverse method is proposed and described. It uses an imposition of mean swirl, i.e., radius times mean tangential velocity, given throughout the meridional section of the turbomachine as an initial design specification. In the present implementation, it is assumed that the fluid is inviscid, incompressible, and irrotational at inlet and that the blades are supposed to have zero thickness. Only blade rows that impart to the fluid a constant work along the space are considered. An application of this procedure to design the rotor of a mixed-flow pump is described in detail. The strategy used to find a suitable mean swirl distribution and the other design inputs is also described. The final blade shape and pressure distributions on the blade surface are presented, showing that it is possible to obtain feasible designs using this technique. Another advantage of this technique is the fact that it does not require large amounts of CPU time.

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.

Laboratory Experiments on the Effects of Blade Strike from Hydrokinetic Energy Technologies on Larval and Juvenile Freshwater Fishes

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

Schweizer, Peter E; Cada, Glenn F; Bevelhimer, Mark S

2012-03-01

There is considerable interest in the development of marine and hydrokinetic energy projects in rivers, estuaries, and coastal ocean waters of the United States. Hydrokinetic (HK) technologies convert the energy of moving water in river or tidal currents into electricity, without the impacts of dams and impoundments associated with conventional hydropower or the extraction and combustion of fossil fuels. The Federal Energy Regulatory Commission (FERC) maintains a database that displays the geographical distribution of proposed HK projects in inland and tidal waters (FERC 2012). As of March 2012, 77 preliminary permits had been issued to private developers to study HKmore » projects in inland waters, the development of which would total over 8,000 MW. Most of these projects are proposed for the lower Mississippi River. In addition, the issuance of another 27 preliminary permits for HK projects in inland waters, and 3 preliminary permits for HK tidal projects (totaling over 3,100 MW) were under consideration by FERC. Although numerous HK designs are under development (see DOE 2009 for a description of the technologies and their potential environmental effects), the most commonly proposed current-based projects entail arrays of rotating devices, much like submerged wind turbines, that are positioned in the high-velocity (high energy) river channels. The many diverse HK designs imply a diversity of environmental impacts, but a potential impact common to most is the risk for blade strike to aquatic organisms. In conventional hydropower generation, research on fish passage through reaction turbines at low-head dams suggested that strike and mortality for small fish could be low. As a consequence of the large surface area to mass ratio of small fish, the drag forces in the boundary layer flow at the surface of a rotor blade may pull small fish around the leading edge of a rotor blade without making physical contact (Turnpenny 1998, Turnpenny et al. 2000). Although there is concern that small, fragile fish early life stages may be unable to avoid being struck by the blades of hydrokinetic turbines, we found no empirical data in the published literature that document survival of earliest life-stage fish in passage by rotor blades. In addition to blade strike, research on passage of fish through conventional hydropower turbines suggested that fish mortalities from passage through the rotor swept area could also occur due to shear stresses and pressure chances in the water column (Cada et al. 1997, Turnpenny 1998). However, for most of the proposed HK turbine designs the rotors are projected to operate a lower RPM (revolutions per minute) than observed from conventional reaction turbines; the associated shear stress and pressure changes are expected to be lower and pose a smaller threat to fish survival (DOE 2009). Only a limited number of studies have been conducted to examine the risk of blade strike from hydrokinetic technologies to fish (Turnpenny et al. 1992, Normandeau et al. 2009, Seitz et al. 2011, EPRI 2011); the survival of drifting or weakly swimming fish (especially early life stages) that encounter rotor blades from hydrokinetic (HK) devices is currently unknown. Our study addressed this knowledge gap by testing how fish larvae and juveniles encountered different blade profiles of hydrokinetic devices and how such encounters influenced survivorship. We carried out a laboratory study designed to improve our understanding of how fish larvae and juvenile fish may be affected by encounters with rotor blades from HK turbines in the water column of river and ocean currents. (For convenience, these early life stages will be referred to as young of the year, YOY). The experiments developed information needed to quantify the risk (both probability and consequences) of rotor-blade strike to YOY fish. In particular, this study attempted to determine whether YOY drifting in a high-velocity flow directly in the path of the blade leading edge will make contact with the rotor blade or will bypass the blade while entrained in the boundary layer of water flowing over the blade surface. The study quantified both immediate and delayed mortalities (observed immediately, 3 hours, and 24 hours after encountering the blade) among freshwater YOY fish resulting from contact with the blade or turbulent flows in the wake of the blade.« less

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors, part 8

NASA Technical Reports Server (NTRS)

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

1974-01-01

An experimental investigation was conducted with an 0.8 hub/tip ratio, single-stage, axial flow compressor to determine the potential of tandem-airfoil blading for improving the efficiency and stable operating range of compressor stages. The investigation included testing of a baseline stage with single-airfoil blading and two tandem-blade stages. The overall performance of the baseline stage and the tandem-blade stage with a 20-80% loading split was considerably below the design prediction. The other tandem-blade stage, which had a rotor with a 50-50% loading split, came within 4.5% of the design pressure rise (delta P(bar)/P(bar) sub 1) and matched the design stage efficiency. The baseline stage with single-airfoil blading, which was designed to account for the actual rotor inlet velocity profile and the effects of axial velocity ratio and secondary flow, achieved the design predicted performance. The corresponding tandem-blade stage (50-50% loading split in both blade rows) slightly exceeded the design pressure rise but was 1.5 percentage points low in efficiency. The tandem rotors tested during both phases demonstrated higher pressure rise and efficiency than the corresponding single-airfoil rotor, with identical inlet and exit airfoil angles.

Development of High-Powered Steam Turbines by OAO NPO Central Research and Design Institute for Boilers and Turbines

NASA Astrophysics Data System (ADS)

Mikhailov, V. E.; Khomenok, L. A.; Kovalev, I. A.

2018-01-01

The article provides an overview of the developments by OAO NPO TsKTI aimed at improvement of components and assemblies of new-generation turbine plants for ultra-supercritical steam parameters to be installed at the power-generating facilities in service. The list of the assemblies under development includes cylinder shells, the cylinder's flow paths and rotors, seals, bearings, and rotor cooling systems. The authors consider variants of the shafting-cylinder configurations for which advanced high-pressure and intermediate-pressure cylinders with reactive blading and low-pressure cylinders of conventional design and with counter-current steam flows are proposed and high-pressure rotors, which can increase the economic efficiency and reduce the overall turbine plant dimensions. Materials intended for the equipment components that operate at high temperatures and a steam cooling technique that allows the use of cheaper steel grades owing to the reduction in the metal's working temperature are proposed. A new promising material for the bearing surfaces is described that enables the operation at higher unit pressures. The material was tested on a full-scale test bench at OAO NPO TsKTI and a turbine in operation. Ways of controlling the erosion of the blades in the moisture-steam turbine compartments by the steam heating of the hollow guide blades are considered. To ensure the dynamic stability of the shafting, shroud and diaphragm seals that prevent the development of the destabilizing circulatory forces of the steam flow were devised and trialed. Advanced instrumentation and software are proposed to monitor the condition of the blading and thermal stresses under transient conditions, to diagnose the vibration processes, and to archive the obtained data. Attention is paid to the normalization of the electromagnetic state of the plant in order to prevent the electrolytic erosion of the plant components. The instrumentation intended for monitoring the relevant electric parameters is described.

Documentation and Control of Flow Separation on a Low Pressure Turbine Linear Cascade of Pak-B Blades Using Plasma Actuators

NASA Technical Reports Server (NTRS)

Corke, Thomas c.; Thomas, FLint, O.; Huang, Junhui

2007-01-01

This work involved the documentation and control of flow separation that occurs over low pressure turbine (LPT) blades at low Reynolds numbers. A specially constructed linear cascade was utilized to study the flow field over a generic LPT cascade consisting of Pratt & Whitney "Pak-B" shaped blades. Flow visualization, surface pressure measurements, LDV measurements, and hot-wire anemometry were conducted to examine the flow fields with and without separation control. Experimental conditions were chosen to give a range of chord Reynolds numbers (based on axial chord and inlet velocity) from 10,000 to 100,000, and a range of freestream turbulence intensities from u'/U(infinity) = 0.08 to 2.85 percent. The blade pressure distributions were measured and used to identify the region of separation that depends on Reynolds number and the turbulence intensity. Separation control was performed using dielectric barrier discharge (DBD) plasma actuators. Both steady and unsteady actuation were implemented and found to work well. The comparison between the steady and unsteady actuators showed that the unsteady actuators worked better than the steady ones. For the steady actuators, it was found that the separated region is significantly reduced. For the unsteady actuators, where the signal was pulsed, the separation was eliminated. The total pressure losses (a low Reynolds number) was reduced by approximately a factor of two. It was also found that lowest plasma duty cycle (10 percent in this work) was as effective as the highest plasma duty cycle (50 percent in this work). The mechanisms of the steady and unsteady plasma actuators were studied. It was suggested by the experimental results that the mechanism for the steady actuators is turbulence tripping, while the mechanism for the unsteady actuators is to generate a train of spanwise structures that promote mixing.

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.

The compressible aerodynamics of rotating blades based on an acoustic formulation

NASA Technical Reports Server (NTRS)

Long, L. N.

1983-01-01

An acoustic formula derived for the calculation of the noise of moving bodies is applied to aerodynamic problems. The acoustic formulation is a time domain result suitable for slender wings and bodies moving at subsonic speeds. A singular integral equation is derived in terms of the surface pressure which must then be solved numerically for aerodynamic purposes. However, as the 'observer' is moved onto the body surface, the divergent integrals in the acoustic formulation are semiconvergent. The procedure for regularization (or taking principal values of divergent integrals) is explained, and some numerical examples for ellipsoids, wings, and lifting rotors are presented. The numerical results show good agreement with available measured surface pressure data.

78 FR 69987 - Airworthiness Directives; Erickson Air-Crane Incorporated Helicopters (Type Certificate...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-11-22

... to require recurring checks of the Blade Inspection Method (BIM) indicator on each blade to determine whether the BIM indicator is signifying that the blade pressure may have been compromised by a blade crack... check procedures for BIM blades installed on the Model S-64E and S-64F helicopters. Several blade spars...

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

14 CFR Appendix A to Part 43 - Major Alterations, Major Repairs, and Preventive Maintenance

Code of Federal Regulations, 2012 CFR

2012-01-01

... alterations: (i) Changes in blade design. (ii) Changes in hub design. (iii) Changes in the governor or control... alterations: (i) Wings. (ii) Tail surfaces. (iii) Fuselage. (iv) Engine mounts. (v) Control system. (vi... the basic design of the fuel, oil, cooling, heating, cabin pressurization, electrical, hydraulic, de...

14 CFR Appendix A to Part 43 - Major Alterations, Major Repairs, and Preventive Maintenance

Code of Federal Regulations, 2014 CFR

2014-01-01

... alterations: (i) Changes in blade design. (ii) Changes in hub design. (iii) Changes in the governor or control... alterations: (i) Wings. (ii) Tail surfaces. (iii) Fuselage. (iv) Engine mounts. (v) Control system. (vi... the basic design of the fuel, oil, cooling, heating, cabin pressurization, electrical, hydraulic, de...

14 CFR Appendix A to Part 43 - Major Alterations, Major Repairs, and Preventive Maintenance

Code of Federal Regulations, 2011 CFR

2011-01-01

... alterations: (i) Changes in blade design. (ii) Changes in hub design. (iii) Changes in the governor or control... alterations: (i) Wings. (ii) Tail surfaces. (iii) Fuselage. (iv) Engine mounts. (v) Control system. (vi... the basic design of the fuel, oil, cooling, heating, cabin pressurization, electrical, hydraulic, de...

14 CFR Appendix A to Part 43 - Major Alterations, Major Repairs, and Preventive Maintenance

Code of Federal Regulations, 2010 CFR

2010-01-01

... alterations: (i) Changes in blade design. (ii) Changes in hub design. (iii) Changes in the governor or control... alterations: (i) Wings. (ii) Tail surfaces. (iii) Fuselage. (iv) Engine mounts. (v) Control system. (vi... the basic design of the fuel, oil, cooling, heating, cabin pressurization, electrical, hydraulic, de...

Fundamental Ice Crystal Accretion Physics Studies

NASA Technical Reports Server (NTRS)

Struk, Peter M.; Broeren, Andy P.; Tsao, Jen-Ching; Vargas, Mario; Wright, William B.; Currie, Tom; Knezevici, Danny; Fuleki, Dan

2012-01-01

Due to numerous engine power-loss events associated with high-altitude convective weather, ice accretion within an engine due to ice crystal ingestion is being investigated. The National Aeronautics and Space Administration (NASA) and the National Research Council (NRC) of Canada are starting to examine the physical mechanisms of ice accretion on surfaces exposed to ice-crystal and mixed-phase conditions. In November 2010, two weeks of testing occurred at the NRC Research Altitude Facility utilizing a single wedge-type airfoil designed to facilitate fundamental studies while retaining critical features of a compressor stator blade or guide vane. The airfoil was placed in the NRC cascade wind tunnel for both aerodynamic and icing tests. Aerodynamic testing showed excellent agreement compared with CFD data on the icing pressure surface and allowed calculation of heat transfer coefficients at various airfoil locations. Icing tests were performed at Mach numbers of 0.2 to 0.3, total pressures from 93 to 45 kPa, and total temperatures from 5 to 15 C. Ice and liquid water contents ranged up to 20 and 3 g/m3, respectively. The ice appeared well adhered to the surface in the lowest pressure tests (45 kPa) and, in a particular case, showed continuous leading-edge ice growth to a thickness greater than 15 mm in 3 min. Such widespread deposits were not observed in the highest pressure tests, where the accretions were limited to a small area around the leading edge. The suction surface was typically ice-free in the tests at high pressure, but not at low pressure. The icing behavior at high and low pressure appeared to be correlated with the wet-bulb temperature, which was estimated to be above 0 C in tests at 93 kPa and below 0 C in tests at lower pressure, the latter enhanced by more evaporative cooling of water. The authors believe that the large ice accretions observed in the low pressure tests would undoubtedly cause the aerodynamic performance of a compressor component such as a stator blade to degrade significantly, and could damage downstream components if shed.

Fundamental Ice Crystal Accretion Physics Studies

NASA Technical Reports Server (NTRS)

Currie, Tom; Knezevici, Danny; Fuleki, Dan; Struk, Peter M.; Broeren, Andy P.; Tsao, Jen-ching; Vargas, Mario; Wright, William

2011-01-01

Due to numerous engine power-loss events associated with high-altitude convective weather, ice accretion within an engine due to ice-crystal ingestion is being investigated. The National Aeronautics and Space Administration (NASA) and the National Research Council (NRC) of Canada are starting to examine the physical mechanisms of ice accretion on surfaces exposed to ice-crystal and mixed-phase conditions. In November 2010, two weeks of testing occurred at the NRC Research Altitude Facility utilizing a single wedge-type airfoil designed to facilitate fundamental studies while retaining critical features of a compressor stator blade or guide vane. The airfoil was placed in the NRC cascade wind tunnel for both aerodynamic and icing tests. Aerodynamic testing showed excellent agreement compared with CFD data on the icing pressure surface and allowed calculation of heat transfer coefficients at various airfoil locations. Icing tests were performed at Mach numbers of 0.2 to 0.3, total pressures from 93 to 45 kPa, and total temperatures from 5 to 15 C. Ice and liquid water contents ranged up to 20 and 3 grams per cubic meter, respectively. The ice appeared well adhered to the surface in the lowest pressure tests (45 kPa) and, in a particular case, showed continuous leading-edge ice growth to a thickness greater than 15 millimeters in 3 minutes. Such widespread deposits were not observed in the highest pressure tests, where the accretions were limited to a small area around the leading edge. The suction surface was typically ice-free in the tests at high pressure, but not at low pressure. The icing behavior at high and low pressure appeared to be correlated with the wet-bulb temperature, which was estimated to be above 0 C in tests at 93 kPa and below 0 C in tests at lower pressure, the latter enhanced by more evaporative cooling of water. The authors believe that the large ice accretions observed in the low pressure tests would undoubtedly cause the aerodynamic performance of a compressor component such as a stator blade to degrade significantly, and could damage downstream components if shed.

Weak Defect Identification for Centrifugal Compressor Blade Crack Based on Pressure Sensors and Genetic Algorithm.

PubMed

Li, Hongkun; He, Changbo; Malekian, Reza; Li, Zhixiong

2018-04-19

The Centrifugal compressor is a piece of key equipment for petrochemical factories. As the core component of a compressor, the blades suffer periodic vibration and flow induced excitation mechanism, which will lead to the occurrence of crack defect. Moreover, the induced blade defect usually has a serious impact on the normal operation of compressors and the safety of operators. Therefore, an effective blade crack identification method is particularly important for the reliable operation of compressors. Conventional non-destructive testing and evaluation (NDT&E) methods can detect the blade defect effectively, however, the compressors should shut down during the testing process which is time-consuming and costly. In addition, it can be known these methods are not suitable for the long-term on-line condition monitoring and cannot identify the blade defect in time. Therefore, the effective on-line condition monitoring and weak defect identification method should be further studied and proposed. Considering the blade vibration information is difficult to measure directly, pressure sensors mounted on the casing are used to sample airflow pressure pulsation signal on-line near the rotating impeller for the purpose of monitoring the blade condition indirectly in this paper. A big problem is that the blade abnormal vibration amplitude induced by the crack is always small and this feature information will be much weaker in the pressure signal. Therefore, it is usually difficult to identify blade defect characteristic frequency embedded in pressure pulsation signal by general signal processing methods due to the weakness of the feature information and the interference of strong noise. In this paper, continuous wavelet transform (CWT) is used to pre-process the sampled signal first. Then, the method of bistable stochastic resonance (SR) based on Woods-Saxon and Gaussian (WSG) potential is applied to enhance the weak characteristic frequency contained in the pressure pulsation signal. Genetic algorithm (GA) is used to obtain optimal parameters for this SR system to improve its feature enhancement performance. The analysis result of experimental signal shows the validity of the proposed method for the enhancement and identification of weak defect characteristic. In the end, strain test is carried out to further verify the accuracy and reliability of the analysis result obtained by pressure pulsation signal.

Weak Defect Identification for Centrifugal Compressor Blade Crack Based on Pressure Sensors and Genetic Algorithm

PubMed Central

Li, Hongkun; He, Changbo

2018-01-01

The Centrifugal compressor is a piece of key equipment for petrochemical factories. As the core component of a compressor, the blades suffer periodic vibration and flow induced excitation mechanism, which will lead to the occurrence of crack defect. Moreover, the induced blade defect usually has a serious impact on the normal operation of compressors and the safety of operators. Therefore, an effective blade crack identification method is particularly important for the reliable operation of compressors. Conventional non-destructive testing and evaluation (NDT&E) methods can detect the blade defect effectively, however, the compressors should shut down during the testing process which is time-consuming and costly. In addition, it can be known these methods are not suitable for the long-term on-line condition monitoring and cannot identify the blade defect in time. Therefore, the effective on-line condition monitoring and weak defect identification method should be further studied and proposed. Considering the blade vibration information is difficult to measure directly, pressure sensors mounted on the casing are used to sample airflow pressure pulsation signal on-line near the rotating impeller for the purpose of monitoring the blade condition indirectly in this paper. A big problem is that the blade abnormal vibration amplitude induced by the crack is always small and this feature information will be much weaker in the pressure signal. Therefore, it is usually difficult to identify blade defect characteristic frequency embedded in pressure pulsation signal by general signal processing methods due to the weakness of the feature information and the interference of strong noise. In this paper, continuous wavelet transform (CWT) is used to pre-process the sampled signal first. Then, the method of bistable stochastic resonance (SR) based on Woods-Saxon and Gaussian (WSG) potential is applied to enhance the weak characteristic frequency contained in the pressure pulsation signal. Genetic algorithm (GA) is used to obtain optimal parameters for this SR system to improve its feature enhancement performance. The analysis result of experimental signal shows the validity of the proposed method for the enhancement and identification of weak defect characteristic. In the end, strain test is carried out to further verify the accuracy and reliability of the analysis result obtained by pressure pulsation signal. PMID:29671821

Effect of blade outlet angle on radial thrust of single-blade centrifugal pump

NASA Astrophysics Data System (ADS)

Nishi, Y.; Fukutomi, J.; Fujiwara, R.

2012-11-01

Single-blade centrifugal pumps are widely used as sewage pumps. However, a large radial thrust acts on a single blade during pump operation because of the geometrical axial asymmetry of the impeller. This radial thrust causes vibrations of the pump shaft, reducing the service life of bearings and shaft seal devices. Therefore, to ensure pump reliability, it is necessary to quantitatively understand the radial thrust and clarify the behavior and generation mechanism. This study investigated the radial thrust acting on two kinds of single-blade centrifugal impellers having different blade outlet angles by experiments and computational fluid dynamics (CFD) analysis. Furthermore, the radial thrust was modeled by a combination of three components, inertia, momentum, and pressure, by applying an unsteady conservation of momentum to this impeller. As a result, the effects of the blade outlet angle on both the radial thrust and the modeled components were clarified. The total head of the impeller with a blade outlet angle of 16 degrees increases more than the impeller with a blade outlet angle of 8 degrees at a large flow rate. In this case, since the static pressure of the circumference of the impeller increases uniformly, the time-averaged value of the radial thrust of both impellers does not change at every flow rate. On the other hand, since the impeller blade loading becomes large, the fluctuation component of the radial thrust of the impeller with the blade outlet angle of 16 degrees increases. If the blade outlet angle increases, the fluctuation component of the inertia component will increase, but the time-averaged value of the inertia component is located near the origin despite changes in the flow rate. The fluctuation component of the momentum component becomes large at all flow rates. Furthermore, although the time-averaged value of the pressure component is almost constant, the fluctuation component of the pressure component becomes large at a large flow rate. In addition to the increase of the fluctuation component of this pressure component, because the fluctuation component of the inertia and momentum components becomes large (as mentioned above), the radial thrust increases at a large flow rate, as is the case for the impeller with a large blade outlet angle.

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.

Data for Design of Entrance Vanes from Two-Dimensional Tests of Airfoils in Cascade

NASA Technical Reports Server (NTRS)

Zimmey, Charles M.; Lappi, Viola M.

1945-01-01

As a part of a program of the NACA directed toward increasing the efficiency of compressors and turbines, data were obtained for application to the design of entrance vanes for axfax-flow compressors or turbines. A series of blower-blade sections with relatively high critical speeds have been developed for turning air efficiently from 0 deg to 80 deg starting with an axial direction. Tests were made of five NACA 65-series blower blades (modified NACA 65(216)-010 airfoils) and of four experimentally designed blower blades in a stationary cascade at low Mach numbers. The turning effectiveness and the pressure distributions of these blade sections at various angles of attack were evaluated over a range of solidities near 1. Entrance-vane design charts are presented that give a blade section and angle of attack for any desired turning angle. The blades thus obtained operate with peak-free pressure distributions. Approximate critical Mach numbers were calculated from the pressure distributions.

Computational fluid dynamics simulation of sound propagation through a blade row.

PubMed

Zhao, Lei; Qiao, Weiyang; Ji, Liang

2012-10-01

The propagation of sound waves through a blade row is investigated numerically. A wave splitting method in a two-dimensional duct with arbitrary mean flow is presented, based on which pressure amplitude of different wave mode can be extracted at an axial plane. The propagation of sound wave through a flat plate blade row has been simulated by solving the unsteady Reynolds average Navier-Stokes equations (URANS). The transmission and reflection coefficients obtained by Computational Fluid Dynamics (CFD) are compared with semi-analytical results. It indicates that the low order URANS scheme will cause large errors if the sound pressure level is lower than -100 dB (with as reference pressure the product of density, main flow velocity, and speed of sound). The CFD code has sufficient precision when solving the interaction of sound wave and blade row providing the boundary reflections have no substantial influence. Finally, the effects of flow Mach number, blade thickness, and blade turning angle on sound propagation are studied.

Start-up circuit upgrading to reduce the erosion of the rotor blades of the last stages of steam turbines and prevent the mass strips of stellite plates

NASA Astrophysics Data System (ADS)

Bozhko, V. V.; Gorin, A. V.; Zaitsev, I. V.; Kovalev, I. A.; Nosovitskii, I. A.; Orlik, V. G.; Lomagin, S. N.; Chernov, V. P.

2017-03-01

At turbine starts with low steam flow rates in idle mode, the low-pressure rotor blades consume energy, causing the ventilation heating of the stages and creating higher depression in them than in the condenser. This leads to the return steam flows in the exhaust of the low-pressure cylinder (LPC), reducing the heat due to the moisture of starting steam damps and cooling injections. It is shown that, as a result of upgrading with the transition to fully milled shroud platforms of rotor blades, the depression in the stages decreases and their cooling efficiency is reduced due to the removal of an elastic turn of the rotor blades under the action of centrifugal forces and seal of them by periphery. Heating the rotor blades of the last stages exceeds the temperature threshold of soldering resistance of stellite plates (150°C), and their mass strips begin. The start-up circuit providing both the temperature retention of the last stages lower the soldering resistance threshold due to overwetting the steam damps up to saturation condition and the high degree of removal from the dump steam of excessive erosive-dangerous condensed moisture was proposed, applied, and tested at the operating power unit. The investment in the development and application of the new start-up circuit are compensated in the course of a year owing to guaranteed prevention of the strips of stellite plates that lengthens the service life of the rotor blades of the last stages as well as increase of the rotor blade efficiency due to the sharp decrease of erosive wear of the profiles and reduction of their surface roughness. This reduces the annual consumption of equivalent fuel by approximately 1000 t for every 100 MW of installed capacity.

Jet Engine Fan Response to Inlet Distortions Generated by Ingesting Boundary Layer Flow

NASA Astrophysics Data System (ADS)

Giuliani, James Edward

Future civil transport designs may incorporate engines integrated into the body of the aircraft to take advantage of efficiency increases due to weight and drag reduction. Additional increases in engine efficiency are predicted if the inlets ingest the lower momentum boundary layer flow that develops along the surface of the aircraft. Previous studies have shown, however, that the efficiency benefits of Boundary Layer Ingesting (BLI) inlets are very sensitive to the magnitude of fan and duct losses, and blade structural response to the non-uniform flow field that results from a BLI inlet has not been studied in-depth. This project represents an effort to extend the modeling capabilities of TURBO, an existing rotating turbomachinery unsteady analysis code, to include the ability to solve the external and internal flow fields of a BLI inlet. The TURBO code has been a successful tool in evaluating fan response to flow distortions for traditional engine/inlet integrations. Extending TURBO to simulate the external and inlet flow field upstream of the fan will allow accurate pressure distortions that result from BLI inlet configurations to be computed and used to analyze fan aerodynamics and structural response. To validate the modifications for the BLI inlet flow field, an experimental NASA project to study flush-mounted S-duct inlets with large amounts of boundary layer ingestion was modeled. Results for the flow upstream and in the inlet are presented and compared to experimental data for several high Reynolds number flows to validate the modifications to the solver. Once the inlet modifications were validated, a hypothetical compressor fan was connected to the inlet, matching the inlet operating conditions so that the effect on the distortion could be evaluated. Although the total pressure distortion upstream of the fan was symmetrical for this geometry, the pressure rise generated by the fan blades was not, because of the velocity non-uniformity of the distortion. Total pressure profiles at various axial locations are computed to identify the overall distortion pattern, how the distortion evolves through the blade passages and mixes out downstream of the blades, and where any critical performance concerns might be. Stall cells are identified that are stationary in the absolute frame and are fixed to the inlet distortion. Flow paths around the blades are examined to study the stall mechanism. Rather than a static airfoil stall, it is observed that the non-uniform pressure loading promotes a three-dimensional dynamic stall. The stall occurs at a point of rapid incidence angle oscillation, observed when a blade passes through the distortion, and re-attaches when the blade leaves the distortion.

Three-Dimensional Flow Field Measurements in a Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

Giel, P. W.; Thurman, D. R.; Lopez, I.; Boyle, R. J.; VanFossen, G. J.; Jett, T. A.; Camperchioli, W. P.; La, H.

1996-01-01

Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 x 10(exp 6) and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 x 10(exp 6) and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.

Formulation of blade-flutter spectral analyses in stationary reference frame

NASA Technical Reports Server (NTRS)

Kurkov, A. P.

1984-01-01

Analytic representations are developed for the discrete blade deflection and the continuous tip static pressure fields in a stationary reference frame. Considered are the sampling rates equal to the rotational frequency, equal to blade passing frequency, and for the pressure, equal to a multiple of the blade passing frequency. For the last two rates the expressions for determining the nodal diameters from the spectra are included. A procedure is presented for transforming the complete unsteady pressure field into a rotating frame of reference. The determination of the true flutter frequency by using two sensors is described. To illustrate their use, the developed procedures are used to interpret selected experimental results.

The Optimization of Four-Stage Low Pressure Turbine with Outlet Guide Vane

NASA Astrophysics Data System (ADS)

Matveev, V.; Baturin, O.; Popov, G.

2018-01-01

The goal of the research was to improve efficiency of four-stage low-pressure turbine with outlet guide vane (LPT) at the design point by optimization the shape of all turbine stator vanes and the stagger angles of all rotor blades. The LPT mathematical model was created by using NUMECA FineTurbo software. Several constraints were imposed the mass flow rate through the LPT and the total pressure ratio at the design point may vary within ±0.5% from the original. Parameters of the stator blade shape and rotor blades stagger angles were obtained. This new geometry of the LPT blades produce a 0.8% increase in efficiency at the design point.

Experimental Investigation of Inter-Blade Vortices in a Model Francis Turbine

NASA Astrophysics Data System (ADS)

LIU, Demin; LIU, Xiaobing; ZHAO, Yongzhi

2017-07-01

The inter-blade vortex in a Francis turbine becomes one of the main hydraulic factors that are likely to cause blade erosion at deep part load operating conditions. However, the causes and the mechanism of inter-blade vortex are still under investigation according to present researches. Thus the causes of inter-blade vortex and the effect of different hydraulic parameters on the inter-blade vortex are investigated experimentally. The whole life cycle of the inter-blade vortex is observed by a high speed camera. The test results illustrate the whole life cycle of the inter-blade vortex from generation to separation and even to fading. It is observed that the inter-blade vortex becomes stronger with the decreasing of flow and head, which leads to pressure fluctuation. Meanwhile, the pressure fluctuations in the vane-less area and the draft tube section become stronger when inter-blade vortices exist in the blade channel. The turbine will be damaged if operating in the inter-blade vortex zone, so its operating range must be far away from that zone. This paper reveals the main cause of the inter-blade vortex which is the larger incidence angle between the inflow angle and the blade angle on the leading edge of the runner at deep part load operating conditions.

SSME Turbopump Turbine Computations

NASA Technical Reports Server (NTRS)

Jorgenson, P. G. E.

1985-01-01

A two-dimensional viscous code was developed to be used in the prediction of the flow in the SSME high-pressure turbopump blade passages. The rotor viscous code (RVC) employs a four-step Runge-Kutta scheme to solve the two-dimensional, thin-layer Navier-Stokes equations. The Baldwin-Lomax eddy-viscosity model is used for these turbulent flow calculations. A viable method was developed to use the relative exit conditions from an upstream blade row as the inlet conditions to the next blade row. The blade loading diagrams are compared with the meridional values obtained from an in-house quasithree-dimensional inviscid code. Periodic boundary conditions are imposed on a body-fitted C-grid computed by using the GRAPE GRids about Airfoils using Poisson's Equation (GRAPE) code. Total pressure, total temperature, and flow angle are specified at the inlet. The upstream-running Riemann invariant is extrapolated from the interior. Static pressure is specified at the exit such that mass flow is conserved from blade row to blade row, and the conservative variables are extrapolated from the interior. For viscous flows the noslip condition is imposed at the wall. The normal momentum equation gives the pressure at the wall. The density at the wall is obtained from the wall total temperature.

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.

Simulation of air admission in a propeller hydroturbine during transient events

NASA Astrophysics Data System (ADS)

Nicolle, J.; Morissette, J.-F.

2016-11-01

In this study, multiphysic simulations are carried out in order to model fluid loading and structural stresses on propeller blades during startup and runaway. It is found that air admission plays an important role during these transient events and that biphasic simulations are therefore required. At the speed no load regime, a large air pocket with vertical free surface forms in the centre of the runner displacing the water flow near the shroud. This significantly affects the torque developed on the blades and thus structural loading. The resulting pressures are applied to a quasi-static structural model and good agreement is obtained with experimental strain gauge data.

An experimental study on the effects of blade row interactions on aerodynamic loss mechanisms in a multistage compressor

NASA Astrophysics Data System (ADS)

Smith, Natalie Rochelle

While the gas turbine engine has existed for nearly 80 years, much of the complex aerodynamics which governs compressor performance is still not well understood. The unsteady flow field consists of periodic blade row interactions from the wakes and potential fields of each blade and vane. Vane clocking is the relative circumferential indexing of adjacent vane rows with the same vane count, and it is one method to change blade row interactions. Though the potential of performance benefits with vane clocking is known, the driving flow physics have yet to be identified. This research examines the effects of blade row interactions on embedded stator total pressure loss and boundary layer transition in the Purdue 3-stage axial compressor. The inlet guide vane, Stator 1, and Stator 2 all have 44 vanes which enable vane clocking of the embedded stage, while the rotors have different blade counts producing amplitude modulation of the unsteady interactions. A detailed investigation of corrected conditions is presented to establish repeatable, compressor performance year-round in a facility utilizing ambient inlet conditions. Without proper humidity accounting of compressor corrected conditions and an understanding of the potential for inlet temperature changes to affect clearances due to thermal growth, measurements of small performance changes in detailed research studies could be indiscernible. The methodology and implementation of a powder-paint flow visualization technique along with the illuminated flow physics are presented in detail. This method assists in understanding the loss development in the compressor by highlighting stator corner separations and endwall flow patterns. Effects of loading condition, rotor tip clearance height, and stator wake and rotor tip leakage interactions are shown with this technique. Vane clocking effects on compressor performance were quantified for nine loading conditions and six clocking configurations - the largest vane clocking dataset in the open literature. These data show that vane clocking effects are small at low loading conditions, including peak efficiency operation, but become stronger as loading increases, and then eventually lessen at near stall operation. Additionally, stator wake profiles and flow visualization reveal that total pressure loss changes are due to a corner separation modulation between clocking configurations. To further address these clocking trends, high-frequency response data were acquired at the Stator 2 inlet and along the Stator 2 surface. The unsteadiness at the Stator 2 inlet was quantified with detailed radial traverses for the different clocking configurations. These data show the effects of interactions between the Stator 1 wake and Rotor 2 tip leakage flow, which result in significantly different inlet flow conditions for Stator 2. The high unsteadiness and blockage region formed by the rotor tip leakage flow changes in size and shape between clocking configurations. Finally, measurements of the Stator 2 surface flows were acquired to investigate the vane clocking effects on unsteady surface pressures and boundary layer transition. These data reveal that Stator 2 performance is influenced by blade row interactions including rotor-rotor interactions, stator wake-rotor tip leakage flow interactions, and vane clocking.

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.

Influence of blade leading edge geometry and upstream blowing on the heat/mass transfer in a turbine cascade

NASA Astrophysics Data System (ADS)

Papa, Marco

The effect of secondary flows on mass transfer from a simulated gas turbine blade and hubwall is investigated. Measurements performed using naphthalene sublimation provide non-dimensional mass transfer coefficients, in the form of Sherwood numbers, that can be converted to heat transfer coefficients through the use of an analogy. Tests are conducted in a linear cascade composed of five blades having the profile of a first stage rotor blade of a high-pressure turbine aircraft engine. Detailed mass transfer maps on the airfoil and endwall surfaces allow the identification of significant flow features that are in good agreement with existing secondary flow models. These results are well-suited for validation of numerical codes, as they are obtained with an accurate technique that does not suffer from conduction or radiation errors and allows the imposition of precise boundary conditions. The performance of a RANS (Reynolds Averaged Navier-Stokes) numerical code that simulates the flow and heat/mass transfer in the cascade using the SST (Shear Stress Transport) k-o model is evaluated through a comparison with the experimental results. Tests performed with a modified blade leading edge show that the introduction of a fillet at the junction with the endwall reduces the effects of the horseshoe vortex in the first part of the passage, while no measurable changes in mass transfer are observed further downstream. Air injected through a slot located upstream of the cascade simulates the engine wheelspace coolant injection between the stator and the rotor. Local mass transfer data obtained injecting naphthalene-free and naphthalene-saturated air are reduced to derive maps of cooling effectiveness on the blade and endwall. Oil dot tests show the surface flow on the endwall. The surface downstream of the gap is coplanar to the upstream surface in the baseline configuration and is shifted to form a forward and backward facing step to investigate the effects of component misalignments. Sufficiently high injection rates alter the structure of the secondary flows and significantly improve the cooling performance.

Numerical analysis of flow in ultra micro centrifugal compressor -influence of meridional configuration

NASA Astrophysics Data System (ADS)

Kaneko, Masanao; Tsujita, Hoshio; Hirano, Toshiyuki

2013-04-01

A single stage ultra micro centrifugal compressor constituting ultra micro gas turbine is required to operate at high rotational speed in order to achieve the pressure ratio which establishes the gas turbine cycle. As a consequence, the aerodynamic losses can be increased by the interaction of a shock wave with the boundary layer on the blade surface. Moreover, the centrifugal force which exceeds the allowable stress of the impeller material can act on the root of blades. On the other hand, the restrictions of processing technology for the downsizing of impeller not only relatively enlarge the size of tip clearance but also make it difficult to shape the impeller with the three-dimensional blade. Therefore, it is important to establish the design technology for the impeller with the two-dimensional blade which possesses the sufficient aerodynamic performance and enough strength to bear the centrifugal force caused by the high rotational speed. In this study, the flow in two types of impeller with the two-dimensional blade which have different meridional configuration was analyzed numerically. The computed results clarified the influence of the meridional configuration on the loss generations in the impeller passage.

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.

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.

Implementation of a Transition Model in a NASA Code and Validation Using Heat Transfer Data on a Turbine Blade

NASA Technical Reports Server (NTRS)

Ameri, Ali A.

2012-01-01

The purpose of this report is to summarize and document the work done to enable a NASA CFD code to model laminar-turbulent transition process on an isolated turbine blade. The ultimate purpose of the present work is to down-select a transition model that would allow the flow simulation of a variable speed power turbine to be accurately performed. The flow modeling in its final form will account for the blade row interactions and their effects on transition which would lead to accurate accounting for losses. The present work only concerns itself with steady flows of variable inlet turbulence. The low Reynolds number k- model of Wilcox and a modified version of the same model will be used for modeling of transition on experimentally measured blade pressure and heat transfer. It will be shown that the k- model and its modified variant fail to simulate the transition with any degree of accuracy. A case is thus made for the adoption of more accurate transition models. Three-equation models based on the work of Mayle on Laminar Kinetic Energy were explored. The three-equation model of Walters and Leylek was thought to be in a relatively mature state of development and was implemented in the Glenn-HT code. Two-dimensional heat transfer predictions of flat plate flow and two-dimensional and three-dimensional heat transfer predictions on a turbine blade were performed and reported herein. Surface heat transfer rate serves as sensitive indicator of transition. With the newly implemented model, it was shown that the simulation of transition process is much improved over the baseline k- model for the single Reynolds number and pressure ratio attempted; while agreement with heat transfer data became more satisfactory. Armed with the new transition model, total-pressure losses of computed three-dimensional flow of E3 tip section cascade were compared to the experimental data for a range of incidence angles. The results obtained, form a partial loss bucket for the chosen blade. In time the loss bucket will be populated with losses at additional incidences. Results obtained thus far will be discussed herein.

The Derivation of the Gradient of the Acoustic Pressure on a Moving Surface for Application to the Fast Scattering Code (FSC)

NASA Technical Reports Server (NTRS)

Farassat, F.; Brentner, Kenneth S.

2005-01-01

In this paper we develop an analytic expression for calculation of the the acoustic pressure from a rotating blade on a moving surface for application to the Fast Scattering Code (FSC). The analytic result is intended to be used in the helicopter noise prediction code PSU-WOPWOP. One of the goals of the derivation is obtaining a result that will not use any more information than are needed for the prediction of the thickness and loading noise. The result derived here achieves this goal and its incorporation in PSU-WOPWOP is straight forward and attainable.

System for determining aerodynamic imbalance

NASA Technical Reports Server (NTRS)

Churchill, Gary B. (Inventor); Cheung, Benny K. (Inventor)

1994-01-01

A system is provided for determining tracking error in a propeller or rotor driven aircraft by determining differences in the aerodynamic loading on the propeller or rotor blades of the aircraft. The system includes a microphone disposed relative to the blades during the rotation thereof so as to receive separate pressure pulses produced by each of the blades during the passage thereof by the microphone. A low pass filter filters the output signal produced by the microphone, the low pass filter having an upper cut-off frequency set below the frequency at which the blades pass by the microphone. A sensor produces an output signal after each complete revolution of the blades, and a recording display device displays the outputs of the low pass filter and sensor so as to enable evaluation of the relative magnitudes of the pressure pulses produced by passage of the blades by the microphone during each complete revolution of the blades.

Space Shuttle Main Engine High Pressure Fuel Turbopump Turbine Blade Cracking

NASA Technical Reports Server (NTRS)

Lee, Henry

1988-01-01

The analytical results from two-dimensional (2D) and three-dimensional (3D) finite element model investigations into the cracking of Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump (HPFTP) first- and second-stage turbine blades are presented. Specifically, the initiation causes for transverse cracks on the pressure side of the firststage blade fir tree lobes and face/corner cracks on the downstream fir tree face of the second-state blade are evaluated. Because the blade material, MAR-M-246 Hf (DS), is highly susceptible to hydrogen embrittlement in the -100 F to 400 F thermal environment, a steady-state condition (full power level = 109 percent) rather than a start-up or shut-down transient was considered to be the most likely candidate for generating a high-strain state in the fir tree areas. Results of the analyses yielded strain levels on both first- and second-stage blade fir tree regions that are of a magnitude to cause hydrogen assisted low cycle fatigue cracking. Also evident from the analysis is that a positive margin against fir tree cracking exists for the planned design modifications, which include shot peening for both first- and second-stage blade fir tree areas.

Aerodynamic Effects of Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie Brynn; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The high turbulence study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Downstream total pressure and exit angle data were acquired for ten incidence angles ranging from +15.8 to 51.0. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12105 to 2.12106 and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 0.25 - 0.4 for the low Tu tests and 8- 15 for the high Tu study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitchyaw probe located in a survey plane 7 axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Gas turbine engine

DOEpatents

Lawlor, Shawn P.; Roberts, II, William Byron

2016-03-08

A gas turbine engine with a compressor rotor having compressor impulse blades that delivers gas at supersonic conditions to a stator. The stator includes a one or more aerodynamic ducts that each have a converging portion and a diverging portion for deceleration of the selected gas to subsonic conditions and to deliver a high pressure oxidant containing gas to flameholders. The flameholders may be provided as trapped vortex combustors, for combustion of a fuel to produce hot pressurized combustion gases. The hot pressurized combustion gases are choked before passing out of an aerodynamic duct to a turbine. Work is recovered in a turbine by expanding the combustion gases through impulse blades. By balancing the axial loading on compressor impulse blades and turbine impulse blades, asymmetrical thrust is minimized or avoided.

Turbine Blade Temperature Measurements Using Thin Film Temperature Sensors

NASA Technical Reports Server (NTRS)

Grant, H. P.; Przybyszewski, J. S.; Claing, R. G.

1981-01-01

The development of thin film temperature sensors is discussed. The technology for sputtering 2 micron thin film platinum versus platinum 10 percent rhodium thermocouples on alumina forming coatings was improved and extended to applications on actual turbine blades. Good adherence was found to depend upon achieving a proper morphology of the alumina surface. Problems of adapting fabrication procedures to turbine blades were uncovered, and improvements were recommended. Testing at 1250 K at one atmosphere pressure was then extended to a higher Mach No. (0.5) in combustor flow for 60 hours and 71 thermal cycles. The mean time to failure was 47 hours accumulated during 1 hour exposures in the combustor. Calibration drift was about 0.1 percent per hour, attributable to oxidation of the rhodium in the thin films. An increase in film thickness and application of a protective overcoat are recommended to reduce drift in actual engine testing.

Design and Development of Turbodrill Blade Used in Crystallized Section

PubMed Central

Yu, Wang; Jianyi, Yao; Zhijun, Li

2014-01-01

Turbodrill is a type of hydraulic axial turbomachinery which has a multistage blade consisting of stators and rotors. In this paper, a turbodrill blade that can be applied in crystallized section under high temperature and pressure conditions is developed. On the basis of Euler equations, the law of energy transfer is analyzed and the output characteristics of turbodrill blade are proposed. Moreover, considering the properties of the layer and the bole-hole conditions, the radical size, the geometrical dimension, and the blade profile are optimized. A computational model of a single-stage blade is built on the ANSYS CFD into which the three-dimensional model of turbodrill is input. In light of the distribution law of the pressure and flow field, the functions of the turbodrill blade are improved and optimized. The turbodrill blade optimization model was verified based on laboratory experiments. The results show that the design meets the deep hard rock mineral exploration application and provides good references for further study. PMID:25276857

Experimental investigation on centrifugal compressor blade crack classification using the squared envelope spectrum.

PubMed

Li, Hongkun; Zhang, Xuefeng; Xu, Fujian

2013-09-18

Centrifugal compressors are a key piece of equipment for modern production. Among the components of the centrifugal compressor, the impeller is a pivotal part as it is used to transform kinetic energy into pressure energy. Blade crack condition monitoring and classification has been broadly investigated in the industrial and academic area. In this research, a pressure pulsation (PP) sensor arranged in close vicinity to the crack area and the corresponding casing vibration signals are used to monitor blade crack information. As these signals cannot directly demonstrate the blade crack, the method employed in this research is based on the extraction of weak signal characteristics that are induced by blade cracking. A method for blade crack classification based on the signals monitored by using a squared envelope spectrum (SES) is presented. Experimental investigations on blade crack classification are carried out to verify the effectiveness of this method. The results show that it is an effective tool for blade crack classification in centrifugal compressors.

Experimental Investigation on Centrifugal Compressor Blade Crack Classification Using the Squared Envelope Spectrum

PubMed Central

Li, Hongkun; Zhang, Xuefeng; Xu, Fujian

2013-01-01

Centrifugal compressors are a key piece of equipment for modern production. Among the components of the centrifugal compressor, the impeller is a pivotal part as it is used to transform kinetic energy into pressure energy. Blade crack condition monitoring and classification has been broadly investigated in the industrial and academic area. In this research, a pressure pulsation (PP) sensor arranged in close vicinity to the crack area and the corresponding casing vibration signals are used to monitor blade crack information. As these signals cannot directly demonstrate the blade crack, the method employed in this research is based on the extraction of weak signal characteristics that are induced by blade cracking. A method for blade crack classification based on the signals monitored by using a squared envelope spectrum (SES) is presented. Experimental investigations on blade crack classification are carried out to verify the effectiveness of this method. The results show that it is an effective tool for blade crack classification in centrifugal compressors. PMID:24051521

Experimental vibration damping characteristics of the third-stage rotor of a three-stage transonic axial-flow compressor

NASA Technical Reports Server (NTRS)

Newman, Frederick A.

1988-01-01

Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gage output power spectra. The combined damping consists of the aerodynanmic damping and the structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given corrected speed, weight flow, and pressure ratio while the structural and mechanical damping is assumed to remain constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third-stage rotor blade aerodynamic damping is presented and discussed for the design equivalent speed with the stator blades reset for maximum efficiency. The compressor overall performance and experimental Campbell diagrams for the third-stage rotor blade row are also presented.

Experimental Vibration Damping Characteristics of the Third-stage Rotor of a Three-stage Transonic Axial-flow Compressor

NASA Technical Reports Server (NTRS)

Newman, Frederick A.

1988-01-01

Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gage output power spectra. The combined damping consists of the aerodynamic damping and the structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given corrected speed, weight flow, and pressure ratio while the structural and mechanical damping is assumed to remain constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third-stage rotor blade aerodynamic damping is presented and discussed for the design equivalent speed with the stator blades reset for maximum efficiency. The compressor overall preformance and experimental Campbell diagrams for the third-stage rotor blade row are also presented.

Experimental and Numerical Investigation of Losses in Low-Pressure Turbine Blade Rows

NASA Technical Reports Server (NTRS)

Dorney, Daniel J.; Lake, James P.; King, Paul I.; Ashpis, David E.

2000-01-01

Experimental data and numerical simulations of low-pressure turbines have shown that unsteady blade row interactions and separation can have a significant impact on the turbine efficiency. Measured turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Several recent studies have revealed that the performance of low-pressure turbine blades is a strong function of the Reynolds number. In the current investigation, experiments and simulations have been performed to study the behavior of a low-pressure turbine blade at several Reynolds numbers. Both the predicted and experimental results indicate increased cascade losses as the Reynolds number is reduced to the values associated with aircraft cruise conditions. In addition, both sets of data show that tripping the boundary layer helps reduce the losses at lower Reynolds numbers. Overall, the predicted aerodynamic and performance results exhibit fair agreement with experimental data.

Model helicopter rotor high-speed impulsive noise: Measured acoustics and blade pressures

NASA Technical Reports Server (NTRS)

Boxwell, D. A.; Schmitz, F. H.; Splettstoesser, W. R.; Schultz, K. J.

1983-01-01

A 1/17-scale research model of the AH-1 series helicopter main rotor was tested. Model-rotor acoustic and simultaneous blade pressure data were recorded at high speeds where full-scale helicopter high-speed impulsive noise levels are known to be dominant. Model-rotor measurements of the peak acoustic pressure levels, waveform shapes, and directively patterns are directly compared with full-scale investigations, using an equivalent in-flight technique. Model acoustic data are shown to scale remarkably well in shape and in amplitude with full-scale results. Model rotor-blade pressures are presented for rotor operating conditions both with and without shock-like discontinuities in the radiated acoustic waveform. Acoustically, both model and full-scale measurements support current evidence that above certain high subsonic advancing-tip Mach numbers, local shock waves that exist on the rotor blades ""delocalize'' and radiate to the acoustic far-field.

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.

Laser Scanning System for Pressure and Temperature Paints

NASA Technical Reports Server (NTRS)

Sullivan, John

1997-01-01

Acquiring pressure maps of aerodynamic surfaces is very important for improving and validating the performance of aerospace vehicles. Traditional pressure measurements are taken with pressure taps embedded in the model surface that are connected to transducers. While pressure taps allow highly accurate measurements to be acquired, they do have several drawbacks. Pressure taps do not give good spatial resolution due to the need for individual pressure tubes, compounded by limited space available inside models. Also, building a model proves very costly if taps are needed because of the large amount of labor necessary to drill, connect and test each one. The typical cost to install one tap is about $200. Recently, a new method for measuring pressure on aerodynamic surfaces has been developed utilizing a technology known as pressure sensitive paints (PSP). Using PSP, pressure distributions can be acquired optically with high spatial resolution and simple model preparation. Flow structures can be easily visualized using PSP, but are missed using low spatial resolution arrays of pressure taps. PSP even allows pressure distributions to be found on rotating machinery where previously this has been extremely difficult or even impossible. The goal of this research is to develop a laser scanning system for use with pressure sensitive paints that allows accurate pressure measurements to be obtained on various aerodynamic surfaces ranging from wind tunnel models to high speed jet engine compressor blades.

The numerical study of the rake angle of impeller blade in centrifugal compressor

NASA Astrophysics Data System (ADS)

Drozdov, A.; Galerkin, Y.

2017-08-01

Investigated impellers have blade surfaces formed by straight generatrix. Blade profiles on shroud and disc surfaces are optimized by velocity diagram control (inviscid, quasi-three dimensional calculations). The blade profiles at hub and shroud blade-to-blade surfaces must be coordinated. A designer can choose the generatrix position at a trailing edge for it. The position is defined by the rake angle that is the angle between a trailing edge generatrix and a meridional plane. Two stages with 3D impellers, vaneless diffusers and return channels were investigated. Seven candidates of impellers of these stages with rake angles in range plus-minus 30 degrees were designed and investigated by quasi-three-dimensional inviscid calculation. CFD-calculations were made for the stages with these impellers. The optimal rake angle is minus 20 degrees for the high flow rate impeller due to lesser blade surface area and favorable meridian velocity field. Zero rake angle is optimal for the medium flow rate impeller where blade surface area is not so important. The combination of inviscid and viscid calculations is the informative instrument for further studies.

Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines

NASA Technical Reports Server (NTRS)

Gallagher, Edward J. (Inventor); Monzon, Byron R. (Inventor); Bugaj, Shari L. (Inventor)

2018-01-01

A gas turbine engine includes a core flow passage, a bypass flow passage, and a propulsor arranged at an inlet of the bypass flow passage and the core flow passage. The propulsor includes a row of propulsor blades. The row includes no more than 20 of the propulsor blades. The propulsor has a pressure ratio between about 1.2 and about 1.7 across the propulsor blades.

Eutectic Composite Turbine Blade Development

DTIC Science & Technology

1976-11-01

turbine blades for aircraft engines . An MC carbide fiber reinforced eutectic alloy, NiTaC-13...composites in turbine blades for aircraft engines . An MC carbide fiber reinforced eutectic alloy, NiTaC-13 and the low pressure turbine blade of the...identified that appeared to have potential for application to aircraft engine turbine blade hardware. The potential benefits offered by these materials

The Uses and Abuses of the Acoustic Analogy in Helicopter Rotor Noise Prediction

NASA Technical Reports Server (NTRS)

Farassat, F.; Brentner, Kenneth S.

1987-01-01

This paper is theoretical in nature and addresses applications of the acoustic analogy in helicopter rotor noise prediction. It is argued that in many instances the acoustic analogy has not been used with care in rotor noise studies. By this it is meant that approximate or inappropriate formulations have been used. By considering various mechanisms of noise generation, such abuses are identified and the remedy is suggested. The mechanisms discussed are thickness, loading, quadrupole, and blade-vortex interaction noise. The quadrupole term of the Ffowcs Williams-Hawkings equation is written in a new form which separates the contributions of regions of high gradients such as shock surfaces. It is shown by order of magnitude studies that such regions are capable of producing noise with the same directivity as the thickness noise. The inclusion of this part of quadrupole sources in current acoustic codes is quite practical. Some of the difficulties with the use of loading noise formulations of the first author in predictions of blade-vortex interaction noise are discussed. It appears that there is a need for development of new theoretical results based on the acoustic analogy in this area. Because of the impulsive character of the blade surface pressure, a time scale of integration different from that used in loading and thickness computations must he used in a computer code for prediction of blade-vortex interaction noise.

End wall flow characteristics and overall performance of an axial flow compressor stage

NASA Technical Reports Server (NTRS)

Sitaram, N.; Lakshminarayana, B.

1983-01-01

This review indicates the possible future directions for research on endwall flows in axial flow compressors. Theoretical investigations on the rotor blade endwall flows in axial flow compressors reported here include the secondary flow calculation and the development of the momentum integral equations for the prediction of the annulus wall boundary layer. The equations for secondary vorticity at the rotor exit are solved analytically. The solution includes the effects of rotation and the viscosity. The momentum integral equations derived include the effect of the blade boundary layers. The axial flow compressor facility of the Department of Aerospace Engineering at The Pennsylvania State University, which is used for the experimental investigations of the endwall flows, is described in some detail. The overall performance and other preliminary experimental results are presented. Extensive radial flow surveys are carried out at the design and various off design conditions. These are presented and interpreted in this report. The following experimental investigations of the blade endwall flows are carried out. (1) Rotor blade endwall flows: The following measurements are carried out at four flow coefficients. (a) The rotor blade static pressures at various axial and radial stations (with special emphasis near the blade tips). (b) The hub wall static pressures inside the rotor blade passage at various axial and tangential stations. (2) IGV endwall flows: The following measurements are carried out at the design flow coefficient. (a) The boundary layer profiles at various axial and tangential stations inside the blade passage and at the blade exit. (b) Casing static pressures and limiting streamline angles inside the blade passage.

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.

The influence of wedge diffuser blade number and divergence angle on the performance of a high pressure ratio centrifugal compressor

NASA Astrophysics Data System (ADS)

Wang, Yi; Han, Ge; Lu, Xingen; Zhu, Junqiang

2018-02-01

Wedge diffuser is widely used in centrifugal compressors due to its high performance and compact size. This paper is aimed to research the influence of wedge diffuser blade number and divergence angle on centrifugal compressor performance. The impact of wedge diffuser blade number on compressor stage performance is investigated, and then the wedge diffusers with different divergence angle are studied by varying diffuser wedge angle and blade number simultaneously. It is found that wedge diffuser with 27 blades could have about 0.8% higher adiabatic efficiency and 0.14 higher total pressure ratio than the wedge diffuser with 19 blades and the best compressor performance is achieved when diffuser divergence angle is 8.3°.These results could give some advices on centrifugal compressor design.

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.

Composite turbine blade design options for Claude (open) cycle OTEC power systems

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

Penney, T R

1985-11-01

Small-scale turbine rotors made from composites offer several technical advantages for a Claude (open) cycle ocean thermal energy conversion (OTEC) power system. Westinghouse Electric Corporation has designed a composite turbine rotor/disk using state-of-the-art analysis methods for large-scale (100-MW/sub e/) open cycle OTEC applications. Near-term demonstrations using conventional low-pressure turbine blade shapes with composite material would achieve feasibility and modern credibility of the open cycle OTEC power system. Application of composite blades for low-pressure turbo-machinery potentially improves the reliability of conventional metal blades affected by stress corrosion.

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.

Performance of Single-Stage Turbine of Mark 25 Torpedo Power Plant with Two Nozzles and Three Rotor-Blade Designs

NASA Technical Reports Server (NTRS)

Schum, Harold J.; Whitney, Warren J.

1949-01-01

A single-stage modification of the turbine from a Mark 25 torpedo power plant was investigated to determine the performance with two nozzles and three rotor-blade designs. The performance was evaluated in terms of brake, rotor, and blade efficiencies at pressure ratios of 8, 15 (design), and 20. The blade efficiencies with the two nozzles are compared with those obtained with four other nozzles previously investigated with the same three rotor-blade designs. Blade efficiency with the cast nozzle of rectangular cross section (J) was higher than that with the circular reamed nozzle (K) at all speeds and pressure ratios with a rotor having a 0.45-inch 17 degree-inlet-angle blades. The efficiencies for both these nozzles were generally low compared with those of the four other nozzles previously investigated in combination with this rotor. At pressure ratios of 15 and 20, the blade efficiencies with nozzle K and the two rotors with 0.40-inch blades having different inlet angles were higher than with the four other nozzles, but the efficiency with nozzle J was generally low. Increasing the blade inlet angle from 17 degrees to 20 degrees had little effect on turbine performance, whereas changing the blade length from 0.40 to 0.45 inch had a marked effect. Although a slight correlation of efficiency with nozzle size was noted for the rotor with 0.45-inch 17 degree-inlet-angle blades, no such effect was discernible ,for the two rotors with 0.40-inch blades.Losses in the supersonic air stream resulting from the complex flow path in the small air passages are probably a large percentage of the total losses, and apparently the effects of changing nozzle size and shape within the limits investigated are of secondary importance.

Blade number impact on pressure and performance of archimedes screw turbine using CFD

NASA Astrophysics Data System (ADS)

Maulana, Muhammad Ilham; Syuhada, Ahmad; Nawawi, Muhammad

2018-02-01

Many rivers in Indonesia can be used as source of mini/micro hydro power plant using low head turbine. The most suitable type of turbine used in fluid flow with low head is the Archimedes screw turbine. The Archimedes screw hydro turbine is a relative newcomer to the small-scale hydropower that can work efficiently on heads as low as 10 meter. In this study, the performance of Archimedes water turbines that has different blade numbers that are thoroughly evaluated to obtain proper blade configuration. For this purpose, numerical simulations are used to predict the pressure changes that occur along the turbine. The simulation results show that turbines with an amount of two blades have more sloping pressure distribution so that it has better stability.

Does a medial retraction blade transmit direct pressure to pharyngeal/esophageal wall during anterior cervical surgery?

PubMed

Han, In Ho; Lee, Su Heon; Lee, Jae Min; Kim, Hwan Soo; Nam, Kyoung Hyup; Duetzmann, Stephan; Park, Jon; Choi, Byung Kwan

2015-01-01

A prospective study of 25 patients who underwent anterior cervical surgery. To assess retraction pressure and the exposure of pharyngeal/esophageal (P/E) wall to the medial retractor blade to clarify whether medial retraction causes direct pressure transmission to the P/E wall. Retraction pressure on P/E walls has been used to explain the relation between the retraction pressure and dysphagia or the efficacies of new retractor blades. However, it is doubtful whether the measured pressure represent real retraction pressure on the P/E wall because exposure of the P/E in the surgical field could be reduced by the shielding effect of thyroid cartilage. Epi- and endoesophageal pressures were serially measured using online pressure transducers 15 minutes before retraction, immediately after retraction, and 30 minutes after retraction. To measure the extent of P/E wall exposure to pressure transducer, we used posterior border of thyroid cartilage as a landmark. Intraoperative radiograph was used to mark the position of the posterior border of thyroid cartilage. We checked out the marked location on retractors by measuring the distance from distal retractor tip. The mean epiesophageal pressure significantly increased after retraction (0 mmHg: 88.7 ± 19.6 mmHg: 81.9 ± 15.3 mmHg). The mean endoesophageal pressure minimally changed after retraction (9.0 ± 6.6 mmHg: 15.7 ± 13.8 mmHg: 17.0 ± 14.3 mmHg). The mean location of the posterior border of thyroid cartilage was 7.3 ± 3.5 mm on the retractor blade from the tip, which means epiesophageal pressure was measured against the posterior border of thyroid cartilage, not against the P/E wall. We suggest that a medial retraction blade does not transmit direct pressure on P/E wall due to minimal wall exposure and intervening thyroid cartilage. Our result should be considered when measuring retraction pressure during anterior cervical surgery or designing novel retractor systems.

Study of inducer load and stress, volume 2

NASA Technical Reports Server (NTRS)

1972-01-01

A program of analysis, design, fabrication and testing has been conducted to develop computer programs for predicting rocket engine turbopump inducer hydrodynamic loading, stress magnitude and distribution, and vibration characteristics. Methods of predicting blade loading, stress, and vibration characteristics were selected from a literature search and used as a basis for the computer programs. An inducer, representative of typical rocket engine inducers, was designed, fabricated, and tested with special instrumentation selected to provide measurements of blade surface pressures and stresses. Data from the tests were compared with predicted values and the computer programs were revised as required to improve correlation. For Volume 1 see N71-20403. For Volume 2 see N71-20404.

Effects of rising angle on upstream blades and intermediate turbine duct

NASA Astrophysics Data System (ADS)

Liu, Jun; Wang, Pei; Du, Qiang; Liu, Guang; Zhu, Junqiang

2016-08-01

With the improvement of requirement, design and manufacture technology, aero-engines for the future are characterized by further reduction in fuel consumption, cost, but increment in propulsion efficiency, which leads to ultra-high bypass ratio. The intermediate turbine duct (ITD), which connects the high pressure turbine (HPT) with the low pressure turbine (LPT), has a critical impact on the overall performances of such future engines. Therefore, it becomes more and more urgent to master the design technique of aggressive, even super-aggressive ITDs. Over the last years, a lot of research works about the flow mechanism in the diffuser ducts were carried out. Many achievements were reported, but further investigation should be performed. With the aid of numerical method, this paper focuses on the change of performance and flow field of ITD, as well as nearby turbines, brought by rising angle (RA). Eight ITDs with the same area ratio and length, but different RAs ranges from 8 degrees to 45 degrees, are compared. According to the investigation, flow field, especially outlet Ma of swirl blade is influenced by RA under potential effect, which is advisable for designers to modify HPT rotor blades after changing ITD. In addition to that, low velocity area moves towards upstream until the first bend as RA increases, while pressure loss distribution at S2 stream surface shows that hub boundary layer is more sensitive to RA, and casing layer keeps almost constant. On the other hand, the overall total pressure loss could keep nearly equivalent among different RA cases, which implies the importance of optimization.

On the Cutting Performance of Segmented Diamond Blades when Dry-Cutting Concrete.

PubMed

Sánchez Egea, A J; Martynenko, V; Martínez Krahmer, D; López de Lacalle, L N; Benítez, A; Genovese, G

2018-02-09

The objective of the present study is to analyze and compare the cutting performance of segmented diamond blades when dry-cutting concrete. A cutting criteria is proposed to characterize the wear of the blades by measuring the variation of the external diameter and the weight loss of the blade. The results exhibit the cutting blade SB-A, which has twice the density of diamonds and large contact area, exhibits less wear even though the material removal rate is higher compared with the other two cutting blades. Additionally, the surface topography of the different blades is evaluated to examine the impact of wear depending on the surface profile and the distribution of the diamonds in the blade's matrix. Large number of diamonds pull-out are found in blades type SB-C, which additionally shows the worst wear resistant capability. As a conclusion, the cutting efficiency of the blade is found to be related to the density of embedded diamonds and the type of the surface profile of the cutting blade after reaching the stop criteria.

Computational fluid dynamics study of the variable-pitch split-blade fan concept

NASA Technical Reports Server (NTRS)

Kepler, C. E.; Elmquist, A. R.; Davis, R. L.

1992-01-01

A computational fluid dynamics study was conducted to evaluate the feasibility of the variable-pitch split-blade supersonic fan concept. This fan configuration was conceived as a means to enable a supersonic fan to switch from the supersonic through-flow type of operation at high speeds to a conventional fan with subsonic inflow and outflow at low speeds. During this off-design, low-speed mode of operation, the fan would operate with a substantial static pressure rise across the blade row like a conventional transonic fan; the front (variable-pitch) blade would be aligned with the incoming flow, and the aft blade would remain fixed in the position set by the supersonic design conditions. Because of these geometrical features, this low speed configuration would inherently have a large amount of turning and, thereby, would have the potential for a large total pressure increase in a single stage. Such a high-turning blade configuration is prone to flow separation; it was hoped that the channeling of the flow between the blades would act like a slotted wing and help alleviate this problem. A total of 20 blade configurations representing various supersonic and transonic configurations were evaluated using a Navier Stokes CFD program called ADAPTNS because of its adaptive grid features. The flow fields generated by this computational procedure were processed by another data reduction program which calculated average flow properties and simulated fan performance. These results were employed to make quantitative comparisons and evaluations of blade performance. The supersonic split-blade configurations generated performance comparable to a single-blade supersonic, through-flow fan configuration. Simulated rotor total pressure ratios of the order of 2.5 or better were achieved for Mach 2.0 inflow conditions. The corresponding fan efficiencies were approximately 75 percent or better. The transonic split-blade configurations having large amounts of turning were able to generate large amounts of total turning and achieve simulated total pressure ratios of 3.0 or better with subsonic inflow conditions. These configurations had large losses and low fan efficiencies in the 70's percent. They had large separated regions and low velocity wakes. Additional turning and diffusion of this flow in a subsequent stator row would probably be very inefficient. The high total pressure ratios indicated by the rotor performance would be substantially reduced by the stators, and the stage efficiency would be substantially lower. Such performance leaves this dual-mode fan concept less attractive than originally postulated.

Pressure Mapping and Efficiency Analysis of an EPPLER 857 Hydrokinetic Turbine

NASA Astrophysics Data System (ADS)

Clark, Tristan

A conceptual energy ship is presented to provide renewable energy. The ship, driven by the wind, drags a hydrokinetic turbine through the water. The power generated is used to run electrolysis on board, taking the resultant hydrogen back to shore to be used as an energy source. The basin efficiency (Power/thrust*velocity) of the Hydrokinetic Turbine (HTK) plays a vital role in this process. In order to extract the maximum allowable power from the flow, the blades need to be optimized. The structural analysis of the blade is important, as the blade will undergo high pressure loads from the water. A procedure for analysis of a preliminary Hydrokinetic Turbine blade design is developed. The blade was designed by a non-optimized Blade Element Momentum Theory (BEMT) code. Six simulations were run, with varying mesh resolution, turbulence models, and flow region size. The procedure was developed that provides detailed explanation for the entire process, from geometry and mesh generation to post-processing analysis tools. The efficiency results from the simulations are used to study the mesh resolution, flow region size, and turbulence models. The results are compared to the BEMT model design targets. Static pressure maps are created that can be used for structural analysis of the blades.

Energy efficient engine high-pressure turbine single crystal vane and blade fabrication technology report

NASA Technical Reports Server (NTRS)

Giamei, A. F.; Salkeld, R. W.; Hayes, C. W.

1981-01-01

The objective of the High-Pressure Turbine Fabrication Program was to demonstrate the application and feasibility of Pratt & Whitney Aircraft-developed two-piece, single crystal casting and bonding technology on the turbine blade and vane configurations required for the high-pressure turbine in the Energy Efficient Engine. During the first phase of the program, casting feasibility was demonstrated. Several blade and vane halves were made for the bonding trials, plus solid blades and vanes were successfully cast for materials evaluation tests. Specimens exhibited the required microstructure and chemical composition. Bonding feasibility was demonstrated in the second phase of the effort. Bonding yields of 75 percent for the vane and 30 percent for the blade were achieved, and methods for improving these yield percentages were identified. A bond process was established for PWA 1480 single crystal material which incorporated a transient liquid phase interlayer. Bond properties were substantiated and sensitivities determined. Tooling die materials were identified, and an advanced differential thermal expansion tooling concept was incorporated into the bond process.

Investigation the Amplitude Uniformity on the Surface of the Wide-Blade Ultrasonic Plastic Welding Horn

NASA Astrophysics Data System (ADS)

Hai Nguyen, Thanh; Thanh Quang, Quang; Luat Tran, Cong; Loc Nguyen, Huu

2017-10-01

Ultrasonic welding has been applied for joining thermoplastic components due to their advantages such as clean, fast and reliable. The basic principle is to use the mechanical energy of ultrasonic frequency vibration to produce the molten pool at the interface of the joined components under high pressure to create solid-state welding joints. Depending on the specific application, the ultrasonic horn is designed to generate suitable amplitudes on the surface of the welding zone. Uniformity of the amplitudes can be a challenge as the welding area increases. Therefore, design a welding horn in order to obtain the uniform amplitudes at the large area is significant difficult. This work presents a method for obtaining the uniform amplitudes at the working surface of the stepped wide-blade horn. Finite element method is used to analyze the amplitude distribution at the horn surface of 250 × 34 mm2 with working frequency of 15 kHz and aluminum alloy 7075. The uniformity of amplitude is obtained by changing the shape of the horn.

Large Eddy Simulation of Ducted Propulsors in Crashbac

NASA Astrophysics Data System (ADS)

Jang, Hyunchul; Mahesh, Krishnan

2008-11-01

Flow around a ducted marine propulsor is computed using the large eddy simulation methodology under crashback conditions. Crashback is an operating condition where a propulsor rotates in the reverse direction while the vessel moves in the forward direction. It is characterized by massive flow separation and highly unsteady propeller loads, which affect both blade life and maneuverability. The simulations are performed on unstructured grids using the algorithm developed by Mahesh at al. (2004, J. Comput. Phys 197). The flow is computed at the advance ratio J=-0.7 and Reynolds number Re=480,000 based on the propeller diameter. Average and RMS values of the unsteady loads such as thrust, torque, and side force on the blades and duct are compared to experiment. It is seen that even though effects of the duct on thrust and torque are not large enough, those on the side force are significant. The rms of side forces is much higher in the presence of the duct. Pressure distributions on blade surfaces and duct surface are examined and used to explain this effect. This work was supported by the United States Office of Naval Research under ONR Grant N00014-05-1-0003.

Experimental Study of Boundary Layer Behavior in a Simulated Low Pressure Turbine. Degree awarded by the University of Toledo, May 1998

NASA Technical Reports Server (NTRS)

Shyne, Rickey J.

1998-01-01

A detailed investigation of the flow physics occurring on the suction side of a simulated Low Pressure Turbine (LPT) blade was performed. A contoured upper wall was designed to simulate the pressure distribution of an actual LPT blade onto a flat plate. The experiments were carried out at Reynolds numbers of 100,000 and 250,000 with three levels of freestream turbulence. Freestream turbulence levels ranging from 0.8% to 3% was used in this experiment. Smoke-wire flow visualization data was used to confirm that the boundary layer was separated and formed a bubble. Hot-wires (single and x-wire) and surface mounted hot-film gases and static pressure taps were used to map the flowfield. The transition process over the separated flow region is observed to be similar to a laminar free shear layer flow with the formation of a large coherent eddy structure. For each condition, the locations defining the separation bubble were determined by careful examination of pressure and mean velocity profile data. Transition onset location and length determined from intermittency profiles decrease as freestream turbulence levels increase. Additionally, the length and height of the laminar separation bubbles were observed to be inversely proportional to the levels of freestream turbulence.

Unsteady potential flow past a propeller blade section

NASA Technical Reports Server (NTRS)

Takallu, M. A.

1990-01-01

An analytical study was conducted to predict the effect of an oscillating stream on the time dependent sectional pressure and lift coefficients of a model propeller blade. The assumption is that as the blade sections encounter a wake, the actual angles of attack vary in a sinusoidal manner through the wake, thus each blade is exposed to an unsteady stream oscillating about a mean value at a certain reduced frequency. On the other hand, an isolated propeller at some angle of attack can experience periodic changes in the value of the flow angle causing unsteady loads on the blades. Such a flow condition requires the inclusion of new expressions in the formulation of the unsteady potential flow around the blade sections. These expressions account for time variation of angle of attack and total shed vortices in the wake of each airfoil section. It was found that the final expressions for the unsteady pressure distribution on each blade section are periodic and that the unsteady circulation and lift coefficients exhibit a hysteresis loop.

Effect of Cricoid Pressure on Laryngeal View During Macintosh, McGrath MAC X-Blade and GlideScope Video Laryngoscopies

PubMed Central

Arslan, Zehra İpek; Solak, Mine

2017-01-01

Objective Cricoid pressure is useful in fasted patients requiring emergency intubation. We compared the effect of cricoid pressure on laryngeal view during Macintosh, McGrath MAC X-Blade and GlideScope video laryngoscopy. Methods After obtaining approval from the Human Research Ethics Committee and written informed consent from patients, we enrolled 120 patients (American Society of Anesthesiologists I–II, age 18–65 years) undergoing elective surgery that required endotracheal intubation in this prospective randomised study. Patients were divided into three groups (Macintosh, McGrath MAC X-Blade and GlideScope). Results Demographic and airway variables were similar in the groups. Cormack-Lehane grades were improved or unchanged on using cricoid pressure in Macintosh and McGrath MAC X-Blade groups. However, laryngeal views worsened in 12 patients (30%), remained unchanged in 26 patients (65%) and improved in 2 patients (5%) in the GlideScope group (p<0.001). Insertion and intubation times for Macintosh and McGrath MAC X-Blade video laryngoscopes were similar. Insertion times for GlideScope and Macintosh video laryngoscopes were similar, but were longer than those for the McGrath MAC X-Blade video laryngoscope (p=0.02). Tracheal intubation took longer with the GlideScope video laryngoscope than with the other devices (p<0.001 and p=0.003). Mean arterial pressures after insertion increased significantly in Macintosh and GlideScope groups (p=0.004 and p=0.001, respectively) compared with post-induction values. Heart rates increased after insertion in all three groups compared with post-induction values (p<0.001). Need for optimisation manoeuvres and postoperative minor complications were comparable in all three groups. Conclusion Although all three devices are useful for normal or difficult intubation, cricoid pressure improved Cormack-Lehane grades of Macintosh and McGrath MAC X-Blade video laryngoscopes but statistically significantly worsened that of the GlideScope video laryngoscope. PMID:29359076

Harmful cleats of football boots: a biomechanical evaluation.

PubMed

Bentley, J A; Ramanathan, A K; Arnold, G P; Wang, W; Abboud, R J

2011-09-01

Football players wear boots of varying cleat designs with some preferring the bladed cleats while others opting for the conventional studded cleats. The current study compares biomechanically the boots with differing cleat designs and their effect on feet, if any. Twenty-nine healthy male volunteers were recruited from amateur football teams. They were asked to perform three trials each of two activities: a straight run and a run cutting at a 60° angle wearing bladed and studded Adidas®-F series boots on artificial turf. Plantar pressure values were recorded using the Pedar®-X in-shoe pressure measuring device. Peak pressure and pressure-time integral were analysed over 11 clinically relevant areas under the foot. While the in-shoe pressure and pressure-time integral were higher under the medial half of the foot with studded boots, they were higher under the lateral half of the foot with the bladed design. The studded boots can be considered safer as the pressure distribution across the foot and the pattern of centre of pressure progression mimicked the normal motif, whereas the bladed boots could potentially be deemed relatively more harmful due to the unnatural increased loading under the lateral half of the foot, predisposing the foot to injuries. Copyright © 2010 European Foot and Ankle Society. Published by Elsevier Ltd. All rights reserved.

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.

Structural Evaluation of a Space Shuttle Main Engine (SSME) High Pressure Fuel Turbopump Turbine Blade

NASA Technical Reports Server (NTRS)

Abdul-Aziz, Ali

1996-01-01

Thermal and structural finite-element analyses were performed on the first high pressure fuel turbopump turbine blade of the space shuttle main engine (SSME). A two-dimensional (2-D) finite-element model of the blade and firtree disk attachment was analyzed using the general purpose MARC (finite-element) code. The loading history applied is a typical test stand engine cycle mission, which consists of a startup condition with two thermal spikes, a steady state and a shutdown transient. The blade material is a directionally solidified (DS) Mar-M 246 alloy, the blade rotor is forged with waspalloy material. Thermal responses under steady-state and transient conditions were calculated. The stresses and strains under the influence of mechanical and thermal loadings were also determined. The critical regions that exhibited high stresses and severe localized plastic deformation were the blade-rotor gaps.

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.

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.

Rotating hot-wire investigation of the vortex responsible for blade-vortex interaction noise

NASA Technical Reports Server (NTRS)

Fontana, Richard Remo

1988-01-01

This distribution of the circumferential velocity of the vortex responsible for blade-vortex interaction noise was measured using a rotating hot-wire rake synchronously meshed with a model helicopter rotor at the blade passage frequency. Simultaneous far-field acoustic data and blade differential pressure measurements were obtained. Results show that the shape of the measured far-field acoustic blade-vortex interaction signature depends on the blade-vortex interaction geometry. The experimental results are compared with the Widnall-Wolf model for blade-vortex interaction noise.

A Neural Network Aero Design System for Advanced Turbo-Engines

NASA Technical Reports Server (NTRS)

Sanz, Jose M.

1999-01-01

An inverse design method calculates the blade shape that produces a prescribed input pressure distribution. By controlling this input pressure distribution the aerodynamic design objectives can easily be met. Because of the intrinsic relationship between pressure distribution and airfoil physical properties, a Neural Network can be trained to choose the optimal pressure distribution that would meet a set of physical requirements. Neural network systems have been attempted in the context of direct design methods. From properties ascribed to a set of blades the neural network is trained to infer the properties of an 'interpolated' blade shape. The problem is that, especially in transonic regimes where we deal with intrinsically non linear and ill posed problems, small perturbations of the blade shape can produce very large variations of the flow parameters. It is very unlikely that, under these circumstances, a neural network will be able to find the proper solution. The unique situation in the present method is that the neural network can be trained to extract the required input pressure distribution from a database of pressure distributions while the inverse method will still compute the exact blade shape that corresponds to this 'interpolated' input pressure distribution. In other words, the interpolation process is transferred to a smoother problem, namely, finding what pressure distribution would produce the required flow conditions and, once this is done, the inverse method will compute the exact solution for this problem. The use of neural network is, in this context, highly related to the use of proper optimization techniques. The optimization is used essentially as an automation procedure to force the input pressure distributions to achieve the required aero and structural design parameters. A multilayered feed forward network with back-propagation is used to train the system for pattern association and classification.

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.

Minnowbrook VI: 2009 Workshop on Flow Physics and Control for Internal and External Aerodynamics

NASA Technical Reports Server (NTRS)

LaGraff, John E.; Povinelli, Louis A.; Gostelow, J. Paul; Glauser, Mark

2010-01-01

Topics covered include: Flow Physics and control for Internal and External Aerodynamics (not in TOC...starts on pg13); Breaking CFD Bottlenecks in Gas-Turbine Flow-Path Design; Streamwise Vortices on the Convex Surfaces of Circular Cylinders and Turbomachinery Blading; DNS and Embedded DNS as Tools for Investigating Unsteady Heat Transfer Phenomena in Turbines; Cavitation, Flow Structure and Turbulence in the Tip Region of a Rotor Blade; Development and Application of Plasma Actuators for Active Control of High-Speed and High Reynolds Number Flows; Active Flow Control of Lifting Surface With Flap-Current Activities and Future Directions; Closed-Loop Control of Vortex Formation in Separated Flows; Global Instability on Laminar Separation Bubbles-Revisited; Very Large-Scale Motions in Smooth and Rough Wall Boundary Layers; Instability of a Supersonic Boundary-Layer With Localized Roughness; Active Control of Open Cavities; Amplitude Scaling of Active Separation Control; U.S. Air Force Research Laboratory's Need for Flow Physics and Control With Applications Involving Aero-Optics and Weapon Bay Cavities; Some Issues Related to Integrating Active Flow Control With Flight Control; Active Flow Control Strategies Using Surface Pressure Measurements; Reduction of Unsteady Forcing in a Vaned, Contra-Rotating Transonic Turbine Configuration; Active Flow Control Stator With Coanda Surface; Controlling Separation in Turbomachines; Flow Control on Low-Pressure Turbine Airfoils Using Vortex Generator Jets; Reduced Order Modeling Incompressible Flows; Study and Control of Flow Past Disk, and Circular and Rectangular Cylinders Aligned in the Flow; Periodic Forcing of a Turbulent Axisymmetric Wake; Control of Vortex Breakdown in Critical Swirl Regime Using Azimuthal Forcing; External and Turbomachinery Flow Control Working Group; Boundary Layers, Transitions and Separation; Efficiency Considerations in Low Pressure Turbines; Summary of Conference; and Final Plenary Session Transcript.

On the Cutting Performance of Segmented Diamond Blades when Dry-Cutting Concrete

PubMed Central

Martynenko, V.; Martínez Krahmer, D.; Benítez, A.; Genovese, G.

2018-01-01

The objective of the present study is to analyze and compare the cutting performance of segmented diamond blades when dry-cutting concrete. A cutting criteria is proposed to characterize the wear of the blades by measuring the variation of the external diameter and the weight loss of the blade. The results exhibit the cutting blade SB-A, which has twice the density of diamonds and large contact area, exhibits less wear even though the material removal rate is higher compared with the other two cutting blades. Additionally, the surface topography of the different blades is evaluated to examine the impact of wear depending on the surface profile and the distribution of the diamonds in the blade’s matrix. Large number of diamonds pull-out are found in blades type SB-C, which additionally shows the worst wear resistant capability. As a conclusion, the cutting efficiency of the blade is found to be related to the density of embedded diamonds and the type of the surface profile of the cutting blade after reaching the stop criteria. PMID:29425125

75 FR 12661 - Airworthiness Directives; General Electric Company CF6-45 and CF6-50 Series Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2010-03-17

... borescope inspections of the high-pressure turbine (HPT) rotor stage 1 and stage 2 blades for wear and... already been done. Borescope Inspection of High-Pressure Turbine (HPT) Stage 1 and Stage 2 Rotor Blades (f... turbofan engines with certain low-pressure turbine (LPT) stage 3 disks installed. This AD requires...

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.

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.

Eye injury risk associated with remote control toy helicopter blades.

PubMed

Alphonse, Vanessa D; Kemper, Andrew R; Rowson, Steven; Duma, Stefan M

2012-01-01

Eye injuries can be caused by a variety of consumer products and toys. Recently, indoor remote controlled (RC) toy helicopters have become very popular. The purpose of this study is to quantify eye injury risk associated with five commercially available RC toy helicopter blades. An experimental matrix of 25 tests was developed to test five different RC toy helicopter blades at full battery power on six postmortem human eyes. A pressure sensor inserted through the optic nerve measured intraocular pressure. Corneal abrasion was assessed post-impact using fluorescein dye. Intraocular pressure was correlated to injury risk for hyphema, lens damage, retinal damage, and globe rupture using published risk functions. All tests resulted in corneal abrasions; however, no other injuries were observed. The 25 tests produced an increase intraocular pressure between 15.2 kPa and 99.3 kPa (114.3 mmHg and 744.7 mmHg). Calculated blade velocities ranged between 16.0 m/s and 25.4 m/s. Injury risk for hyphema was a maximum of 0.2%. Injury risk for lens damage, retinal damage, and globe rupture was 0.0% for all tests. Blade design parameters such as length and mass did not affect the risk of eye injury. This is the first study to quantify the risk of eye injury from RC toy helicopter blades. While corneal abrasions were observed, more serious eye injuries were neither observed nor predicted to have occurred. Results from this study are critical for establishing safe design thresholds for RC toy helicopter blades so that more serious injuries can be prevented.

Predicting the Rotor-Stator Interaction Acoustics of a Ducted Fan Engine

NASA Technical Reports Server (NTRS)

Biedron, Robert T.; Rumsey, Christopher L.; Podboy, Gary G.; Dunn, M. H.

2001-01-01

A Navier-Stokes computation is performed for a ducted-fan configuration with the goal of predicting rotor-stator noise generation without having to resort to heuristic modeling. The calculated pressure field in the inlet region is decomposed into classical infinite-duct modes, which are then used in either a hybrid finite-element/Kirchhoff surface method or boundary integral equation method to calculate the far field noise. Comparisons with experimental data are presented, including rotor wake surveys and far field sound pressure levels for two blade passage frequency (BPF) tones.

76 FR 2605 - Airworthiness Directives; Rolls-Royce plc RB211-Trent 800 Series Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2011-01-14

... blade root lubrication coating system may be reduced dependant on the extent of previous running with... Trent 895 powered Boeing 777-200 aircraft experienced release of a low pressure (LP) compressor blade which failed due to fatigue cracking in the root section of the blade. The released blade (undercut root...

Investigation of Effectiveness of Air-Heating a Hollow Steel Propeller for Protection Against Icing. 2: 50% Impartitioned Blades

NASA Technical Reports Server (NTRS)

Perkins, Porter J.; Mulholland, Donald R.

1948-01-01

The icing protection afforded an internal air-heated propeller blade by radial partitioning at 50-percent chord to confine the heated air to the forward half of the blade was determined in the NACA Cleveland icing research tunnel. A modified production-model hollow steel propeller, was used for the investigation. Temperatures of the blade surfaces for several heating rates were measured under various tunnel Icing' conditions. Photographic observations of ice formations on blade surfaces and blade heat-exchanger effectiveness were obtained. With 50-percent partitioning of the blades, adequate icing protection at 1050 rpm was obtained with a heating rate of 26,000 Btu per hour per blade at the blade shank using an air temperature of 400 F with a flow rate of 280 pounds per hour per blade, which is one-third less heat than was found necessary for similar Ice protection with unpartitioned blades. The chordwise distribution of the applied heat, as determined by surface temperature measurements, was considered unsatisfactory with much of the heat dissipated well back of the leading edge. Heat-exchanger effectiveness of approximately 56 percent also Indicated poor utilization of available heat. This effectiveness was, however, 9 percent greater than that obtained from unpartitioned blades.

Aeroelastic Considerations For Rotorcraft Primary Control with On-Blade Elevons

NASA Technical Reports Server (NTRS)

Ormiston, Robert A.; Rutkowski, Michael (Technical Monitor)

2001-01-01

Replacing the helicopter rotor swashplate and blade pitch control system with on-blade elevon control surfaces for primary flight control may significantly reduce weight and drag to improve mission performance. Simplified analyses are used to examine the basic aeroelastic characteristics of such rotor blades, including pitch and flap dynamic response, elevon reversal, and elevon control effectiveness. The profile power penalty associated with deflections of elevon control surfaces buried within the blade planform is also evaluated. Results suggest that with aeroelastic design for pitch frequencies in the neighborhood of 2/rev, reasonable elevon control effectiveness may be achieved and that, together with collective pitch indexing, the aerodynamic profile power penalty of on-blade control surface deflections may be minimized.

Effect of Blade-surface Finish on Performance of a Single-stage Axial-flow Compressor

NASA Technical Reports Server (NTRS)

Moses, Jason J; Serovy, George, K

1951-01-01

A set of modified NACA 5509-34 rotor and stator blades was investigated with rough-machine, hand-filed, and highly polished surface finishes over a range of weight flows at six equivalent tip speeds from 672 to 1092 feet per second to determine the effect of blade-surface finish on the performance of a single-stage axial-flow compressor. Surface-finish effects decreased with increasing compressor speed and with decreasing flow at a given speed. In general, finishing blade surfaces below the roughness that may be considered aerodynamically smooth on the basis of an admissible-roughness formula will have no effect on compressor performance.

Unsteady Flow Field in a Multistage Axial Flow Compressor

NASA Technical Reports Server (NTRS)

Suryavamshi, N.; Lakshminarayana, B.; Prato, J.

1997-01-01

The flow field in a multistage compressor is three-dimensional, unsteady, and turbulent with substantial viscous effects. Some of the specific phenomena that has eluded designers include the effects of rotor-stator and rotor-rotor interactions and the physics of mixing of velocity, pressure, temperature and velocity fields. An attempt was made, to resolve experimentally, the unsteady pressure and temperature fields downstream of the second stator of a multistage axial flow compressor which will provide information on rotor-stator interaction effects and the nature of the unsteadiness in an embedded stator of a three stage axial flow compressor. Detailed area traverse measurements using pneumatic five hole probe, thermocouple probe, semi-conductor total pressure probe (Kulite) and an aspirating probe downstream of the second stator were conducted at the peak efficiency operating condition. The unsteady data was then reduced through an ensemble averaging technique which splits the signal into deterministic and unresolved components. Auto and cross correlation techniques were used to correlate the deterministic total temperature and velocity components (acquired using a slanted hot-film probe at the same measurement locations) and the gradients, distributions and relative weights of each of the terms of the average passage equation were then determined. Based on these measurements it was observed that the stator wakes, hub leakage flow region, casing endwall suction surface corner region, and the casing endwall region away from the blade surfaces were the regions of highest losses in total pressure, lowest efficiency and highest levels of unresolved unsteadiness. The deterministic unsteadiness was found to be high in the hub and casing endwall regions as well as on the pressure side of the stator wake. The spectral distribution of hot-wire and kulite voltages shows that at least eight harmonics of all three rotor blade passing frequencies are present at this measurement location. In addition to the basic three rotor blade passing frequencies (R1, R2 and R3) and their harmonics, various difference frequencies such as (2R1 -R2) and (2R3-R2) and their harmonics are also observed. These difference frequencies are due to viscous and potential interactions between rotors 1, 2 and 3 which are sensed by both the total pressure and aspirating probes at this location. Significant changes occur to the stator exit flow features with passage of the rotor upstream of the stator. Because of higher convection speeds of the rotor wake on the suction surface of the downstream stator than on the pressure side, the chopped rotor wake was found to be arriving at different times on either side of the stator wake. As the rotor passes across the stator.

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to -51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Aerodynamic Effects of High Turbulence Intensity on a Variable-Speed Power-Turbine Blade With Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. These results are compared to previous measurements made in a low turbulence environment. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at these large incidence and Reynolds number variations complements the existing database. Downstream total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8deg to -51.0deg. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12×10(exp 5) to 2.12×10(exp 6) and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

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.

Low-speed wind tunnel tests of a 50.8-centimeter (20-in.) 1.15-pressure-ratio fan engine model

NASA Technical Reports Server (NTRS)

Wesoky, H. L.; Abbott, J. M.; Albers, J. A.; Dietrich, D. A.

1974-01-01

At a typical STOL aircraft takeoff and landing velocity, wind tunnel aerodynamic and acoustic measurements demonstrated that an inlet lip-area contraction ratio of 1.35 was superior to a ratio of 1.26 at high incidence angles. A 17 percent reduction in net thrust and an increase of 9 decibels in sound pressure level at the blade passing frequency resulted from inlet flow separation at an incidence angle of 50 deg with the 1.26-contraction-ratio inlet. Reverse-thrust forces obtained with blade rotation through the feathered angle were 1.8 times larger than with blade rotation through the flat angle. Reverse-thrust force was reduced from 30 to 50 percent and sound pressure level increased from 3 to 7 decibels at the blade passing frequency between the wind-tunnel-off condition and a typical STOL aircraft landing velocity.

"Space slitter" for film or tape

NASA Technical Reports Server (NTRS)

Johnson, W. H.

1978-01-01

Device cuts film or tape into strips by guiding film in channel under cutting blades. Device is operated by lifting pressure bar to insert blades into film. Film is then pulled through blades. Cutter has potential uses in advertising, commercial art, and publishing fields.

The turbomachine blading design using S2-S1 approach

NASA Technical Reports Server (NTRS)

Luu, T. S.; Bencherif, L.; Viney, B.; Duc, J. M. Nguyen

1991-01-01

The boundary conditions corresponding to the design problem when the blades being simulated by the bound vorticity distribution are presented. The 3D flow is analyzed by the two steps S2 - S1 approach. In the first step, the number of blades is supposed to be infinite, the vortex distribution is transformed into an axisymmetric one, so that the flow field can be analyzed in a meridional plane. The thickness distribution of the blade producing the flow channel striction is taken into account by the modification of metric tensor in the continuity equation. Using the meridional stream function to define the flow field, the mass conservation is satisfied automatically. The governing equation is deduced from the relation between the azimuthal component of the vorticity and the meridional velocity. The value of the azimuthal component of the vorticity is provided by the hub to shroud equilibrium condition. This step leads to the determination of the axisymmetric stream sheets as well as the approximate camber surface of the blade. In the second step, the finite number of blades is taken into account, the inverse problem corresponding to the blade to blade flow confined in each stream sheet is analyzed. The momentum equation implies that the free vortex of the absolute velocity must be tangential to the stream sheet. The governing equation for the blade to blade flow stream function is deduced from this condition. At the beginning, the upper and the lower surfaces of the blades are created from the camber surface obtained from the first step with the assigned thickness distribution. The bound vorticity distribution and the penetrating flux conservation applied on the presumed blade surface constitute the boundary conditions of the inverse problem. The detection of this flux leads to the rectification of the geometry of the blades.

An interactive grid generation procedure for axial and radial flow turbomachinery

NASA Technical Reports Server (NTRS)

Beach, Timothy A.

1989-01-01

A combination algebraic/elliptic technique is presented for the generation of three dimensional grids about turbo-machinery blade rows for both axial and radial flow machinery. The technique is built around use of an advanced engineering workstation to construct several two dimensional grids interactively on predetermined blade-to-blade surfaces. A three dimensional grid is generated by interpolating these surface grids onto an axisymmetric grid. On each blade-to-blade surface, a grid is created using algebraic techniques near the blade to control orthogonality within the boundary layer region and elliptic techniques in the mid-passage to achieve smoothness. The interactive definition of bezier curves as internal boundaries is the key to simple construction. This procedure lends itself well to zonal grid construction, an important example being the tip clearance region. Calculations done to date include a space shuttle main engine turbopump blade, a radial inflow turbine blade, and the first stator of the United Technologies Research Center large scale rotating rig. A finite Navier-Stokes solver was used in each case.

Quiet Clean Short-haul Experimental Engine (QCSEE). The aerodynamic and mechanical design of the QCSEE over-the-wing fan

NASA Technical Reports Server (NTRS)

1976-01-01

The aerodynamic and mechanical design of a fixed-pitch 1.36 pressure ratio fan for the over-the-wing (OTW) engine is presented. The fan has 28 blades. Aerodynamically, the fan blades were designed for a composite blade, but titanium blades were used in the experimental fan as a cost savings measure.

75 FR 38052 - Airworthiness Directives; Pratt & Whitney JT8D-209, -217, -217A, -217C, and -219 Series Turbofan...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-07-01

... refurbished or replaced. (p) Whenever a used (service run) blade is reinstalled in a rotor, the previous used...- pressure turbine (LPT) blades for shroud notch wear and replacement of the blade if wear limits are... are proposing this AD to prevent turbine blade failures that could result in uncontained engine debris...

76 FR 24798 - Airworthiness Directives; Rolls-Royce plc (RR) RB211-Trent 875-17, RB211-Trent 877-17, RB211...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-05-03

... blade root lubrication coating system may be reduced dependant on the extent of previous running with... Trent 895 powered Boeing 777-200 aircraft experienced release of a low pressure (LP) compressor blade which failed due to fatigue cracking in the root section of the blade. The released blade (undercut root...

75 FR 26681 - Airworthiness Directives; Turbomeca S.A. Arriel 2B and 2B1 Turboshaft Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2010-05-12

... displacement of the HP blade relative to the disk in case of blade lock rupture or opening, therefore... repetitive borescope inspection of the engine for rearward displacement of the high-pressure turbine (HP... axial displacement of the HP blade relative to the disk in case of blade lock rupture or opening...

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.

Steady and Unsteady Velocity Measurements in a Small Turbocharger Turbine with Computational Validation

NASA Astrophysics Data System (ADS)

Karamanis, N.; Palfreyman, D.; Arcoumanis, C.; Martinez-Botas, R. F.

2006-07-01

The detailed flow characteristics of three high-pressure-ratio mixed-flow turbines were investigated under both steady and pulsating flow conditions. Two rotors featured a constant inlet blade angle, one with 12 blades and the second with 10. The third rotor was shorter and had a nominally constant incidence angle. The rotors find application on an automotive high-speed large commercial diesel turbocharger. The steady flow entering and exiting the blades has been quantified by a laser Doppler velocimetry system. The measurements were performed at a plane 3.0-mm ahead of the rotor leading edge and 9.5-mm downstream the rotor trailing edge. The turbine test conditions corresponded to the peak efficiency point at two rotational speeds, 29,400 and 41,300-rpm. The results were resolved in a blade-to-blade sense to examine fully the nature of the flow at turbocharger representative conditions. A correlation between the combined effects of incidence and exit flow angle with the isentropic efficiency has been verified. Regarding pulsating flow, the velocity data and their corresponding instantaneous velocity triangles were resolved in a blade-to-blade sense to understand better the complex phenomenon. The results highlighted the potential of a nominally constant incidence design to absorb better the inadequacy of the volute to discharge the exhaust gas uniformly along the blade leading edge. A double vortex rotating in a clockwise sense propagated on the plane normal to the meridional direction. This should be attributed to the effect of the passing blade that was acting as a blockage to the flow. The phenomenon was more pronounced near the suction and pressure surfaces of the blade, but diminished at the mid-passage region where the flow exhibited its best level of guidance. The full mixed flow turbine stage under transient conditions was modelled firstly with a 'steady' inlet and secondly with a 'pulsating' inlet boundary condition. In both cases comparison was made to experiment performance and LDV measurements. With the steady inlet boundary condition, a high level of accuracy was achieved when compared to the experimental performance and velocity field. The velocity along the leading edge showed the same discrepancy as the single passage analysis that is with the radial and axial component from mid span to the blade tip. At the trailing edge features identified in the experimental data are identified in the numerical results; the velocity field appears more 'diffused' across the plane as per the experimental data than from the single passage analysis. With the pulsating inlet boundary, the predicted velocity traces in the volute and close to the turbine lead and trailing edge show excellent agreement in both form (against time) and magnitude.

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.

An experimental investigation of gapwise periodicity and unsteady aerodynamic response in an oscillating cascade. 1: Experimental and theoretical results. [turbine blades

NASA Technical Reports Server (NTRS)

Carta, F. O.

1982-01-01

Tests were conducted on a linear cascade of airfoils oscillating in pitch to measure the unsteady pressure response on selected blades along the leading edge plane of the cascade, over the chord of the center blade, and on the sidewall in the plane of the leading edge. The pressure data were reduced to Fourier coefficient form for direct comparison, and were also processed to yield integrated loads and, particularly, the aerodynamic damping coefficient. Results from the unsteady Verdon/Caspar theory for cascaded blades with nonzero thickness and camber were compared with the experimental measurements. The three primary results are: (1) from the leading edge plane blade data, the cascade was judged to be periodic in unsteady flow over the range of parameters tested; (2) the interblade phase angle was found to be the single most important parameter affecting the stability of the oscillating cascade blades; and (3) the real blade theory and the experiment were in excellent agreement for the several cases chosen for comparison.

Advanced turbo-prop airplane interior noise reduction-source definition

NASA Technical Reports Server (NTRS)

Magliozzi, B.; Brooks, B. M.

1979-01-01

Acoustic pressure amplitudes and phases were measured in model scale on the surface of a rigid semicylinder mounted in an acoustically treated wind tunnel near a prop-fan (an advanced turboprop with many swept blades) model. Operating conditions during the test simulated those of a prop-fan at 0.8 Mach number cruise. Acoustic pressure amplitude and phase contours were defined on the semicylinder surface. Measurements obtained without the semi-cylinder in place were used to establish the magnitude of pressure doubling for an aircraft fuselage located near a prop-fan. Pressure doubling effects were found to be 6dB at 90 deg incidence decreasing to no effect at grazing incidence. Comparisons of measurements with predictions made using a recently developed prop-fan noise prediction theory which includes linear and non-linear source terms showed good agreement in phase and in peak noise amplitude. Predictions of noise amplitude and phase contours, including pressure doubling effects derived from test, are included for a full scale prop-fan installation.

Heat Transfer Measurements and Predictions on a Power Generation Gas Turbine Blade

NASA Technical Reports Server (NTRS)

Giel, Paul W.; Bunker, Ronald S.; VanFossen, G. James; Boyle, Robert J.

2000-01-01

Detailed heat transfer measurements and predictions are given for a power generation turbine rotor with 129 deg of nominal turning and an axial chord of 137 mm. Data were obtained for a set of four exit Reynolds numbers comprised of the design point of 628,000, -20%, +20%, and +40%. Three ideal exit pressure ratios were examined including the design point of 1.378, -10%, and +10%. Inlet incidence angles of 0 deg and +/-2 deg were also examined. Measurements were made in a linear cascade with highly three-dimensional blade passage flows that resulted from the high flow turning and thick inlet boundary layers. Inlet turbulence was generated with a blown square bar grid. The purpose of the work is the extension of three-dimensional predictive modeling capability for airfoil external heat transfer to engine specific conditions including blade shape, Reynolds numbers, and Mach numbers. Data were obtained by a steady-state technique using a thin-foil heater wrapped around a low thermal conductivity blade. Surface temperatures were measured using calibrated liquid crystals. The results show the effects of strong secondary vortical flows, laminar-to-turbulent transition, and also show good detail in the stagnation region.

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.

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.

Turbine blade damping device with controlled loading

DOEpatents

Marra, John J

2013-09-24

A damping structure for a turbomachine rotor. The damping structure including an elongated snubber element including a first snubber end rigidly attached to a first blade and extending toward an adjacent second blade, and an opposite second snubber end positioned adjacent to a cooperating surface associated with the second blade. The snubber element has a centerline extending radially inwardly in a direction from the first blade toward the second blade along at least a portion of the snubber element between the first and second snubber ends. Rotational movement of the rotor effects relative movement between the second snubber end and the cooperating surface to position the second snubber end in frictional engagement with the cooperating surface with a predetermined damping force determined by a centrifugal force on the snubber element.

Turbine blade damping device with controlled loading

DOEpatents

Marra, John J.

2015-09-29

A damping structure for a turbomachine rotor. The damping structure including an elongated snubber element including a first snubber end rigidly attached to a first blade and extending toward an adjacent second blade, and an opposite second snubber end positioned adjacent to a cooperating surface associated with the second blade. The snubber element has a centerline extending radially inwardly in a direction from the first blade toward the second blade along at least a portion of the snubber element between the first and second snubber ends. Rotational movement of the rotor effects relative movement between the second snubber end and the cooperating surface to position the second snubber end in frictional engagement with the cooperating surface with a predetermined damping force determined by a centrifugal force on the snubber element.

TX-100 manufacturing final project report.

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

Ashwill, Thomas D.; Berry, Derek S.

This report details the work completed under the TX-100 blade manufacturing portion of the Carbon-Hybrid Blade Developments: Standard and Twist-Coupled Prototype project. The TX-100 blade is a 9 meter prototype blade designed with bend-twist coupling to augment the mitigation of peak loads during normal turbine operation. This structural coupling was achieved by locating off axis carbon fiber in the outboard portion of the blade skins. The report will present the tooling selection, blade production, blade instrumentation, blade shipping and adapter plate design and fabrication. The baseline blade used for this project was the ERS-100 (Revision D) wind turbine blade. Themore » molds used for the production of the TX-100 were originally built for the production of the CX-100 blade. The same high pressure and low pressure skin molds were used to manufacture the TX-100 skins. In order to compensate for the difference in skin thickness between the CX-100 and the TX-100, however, a new TX-100 shear web plug and mold were required. Both the blade assembly fixture and the root stud insertion fixture used for the CX-100 blades could be utilized for the TX-100 blades. A production run of seven TX-100 prototype blades was undertaken at TPI Composites during the month of October, 2004. Of those seven blades, four were instrumented with strain gauges before final assembly. After production at the TPI Composites facility in Rhode Island, the blades were shipped to various test sites: two blades to the National Wind Technology Center at the National Renewable Energy Laboratory in Boulder, Colorado, two blades to Sandia National Laboratory in Albuquerque, New Mexico and three blades to the United States Department of Agriculture turbine field test facility in Bushland, Texas. An adapter plate was designed to allow the TX-100 blades to be installed on existing Micon 65/13M turbines at the USDA site. The conclusion of this program is the kick-off of the TX-100 blade testing at the three testing facilities.« less

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.

Spiral inlets for steam turbines

NASA Astrophysics Data System (ADS)

Škach, Radek; Uher, Jan

2017-09-01

This paper deals with the design process of special nozzle blades for spiral inlets. Spiral inlets are used for the first stages of high pressure and intermediate pressure steam turbines with both reaction and impulse blades when throttling or sliding pressure control is applied. They improve the steam flow uniformity from the inlet pipe and thus decrease the aerodynamic losses. The proposed evaluation of the inlet angle is based on the free vortex law.

Experimental Determination of Aerodynamic Damping in a Three-Stage Transonic Axial-Flow Compressor. Degree awarded by Case Western Reserve Univ.

NASA Technical Reports Server (NTRS)

Newman, Frederick A.

1988-01-01

Rotor blade aerodynamic damping is experimentally determined in a three-stage transonic axial flow compressor having design aerodynamic performance goals of 4.5:1 pressure ratio and 65.5 lbm/sec weight flow. The combined damping associated with each mode is determined by a least squares fit of a single degree of freedom system transfer function to the nonsynchronous portion of the rotor blade strain gauge output power spectra. The combined damping consists of aerodynamic and structural and mechanical damping. The aerodynamic damping varies linearly with the inlet total pressure for a given equivalent speed, equivalent mass flow, and pressure ratio while structural and mechanical damping are assumed to be constant. The combined damping is determined at three inlet total pressure levels to obtain the aerodynamic damping. The third stage rotor blade aerodynamic damping is presented and discussed for 70, 80, 90, and 100 percent design equivalent speed. The compressor overall performance and experimental Campbell diagrams for the third stage rotor blade row are also presented.

A flight investigation of blade-section aerodynamics for a helicopter main rotor having 10-64C airfoil sections

NASA Technical Reports Server (NTRS)

Morris, C. E. K., Jr.

1981-01-01

Pressure data at 90 percent blade radius were obtained for a helicopter main rotor with 10-64C blade sections during flight. Concurrent measurements ere made of vehicle flight state, performance and some rotor loads. The test envelope included hover, level flight from about 65 to 162 knots, climb and descent, and collective fixed maneuvers. Good agreement is shown between some sets of airfoil pressure distributions obtained in flight and those from two-dimensional wind-tunnel tests or theoretical calculations.

Measurement of Gust Response on a Turbine Cascade

NASA Technical Reports Server (NTRS)

Kurkov, A. P.; Lucci, B. L.

1995-01-01

The paper presents benchmark experimental data on a gust response of an annular turbine cascade. The experiment was particularly designed to provide data for comparison with the results of a typical linearized gust-response analysis. Reduced frequency, Mach number, and incidence were varied independently. Except for the lowest reduced frequency, the gust velocity distribution was nearly sinusoidal. For the high inlet-velocity series of tests, the cascade was near choking. The mean flow was documented by measuring blade surface pressures and the cascade exit flow. High-response pressure transducers were used to measure the unsteady pressure distribution. Inlet-velocity components and turbulence parameters were measured using hot wire. In addition to the synchronous time-average pressure spectra, typical power spectra are included for several representative conditions.

A Laboratory Scale Vortex Generator for Simulation of Martian Dust Devils.

NASA Astrophysics Data System (ADS)

Balme, M.; Greeley, R.; Mickelson, B.; Iversen, J.; Beardmore, G.; Metzger, S.

2001-12-01

Martian dust particles are a few microns in diameter. Current Martian ambient wind speeds appear to be insufficient to lift such fine particles and are marginal to entrain even the optimum particles sizes for threshold (100-160mm diameter). Instead, dust devils were suggested as a local source of airborne particles and have been observed on Mars both from orbit and from lander data. Dust devils lift particles through enhanced local wind speeds and by a pressure drop often associated with the vortex which provides `lift'. This study seeks to 1) quantify the relative importance of enhanced wind speed versus pressure drop lift in dust devil entrainment threshold; 2) measure the mass transport potential of dust devils; 3) investigate the effects of surface roughness and topography on dust devil morphology; 4) quantify the overall effects of low atmospheric pressure on the formation, structure and entrainment processes of dust devils. To investigate the particle lifting properties of dust devils, a laboratory vortex generator was fabricated. It consists of a large vertical cylinder (45 and 75cm in diameter) containing a motor-driven rotor comprised of four vertical blades. Beneath the cylinder is a 2.4 by 2.4 m tabletop containing 14 differential pressure transducer ports used to measure the surface pressure structure of the vortex. Both the distance between the cylinder and the tabletop and the height of the blades within the cylinder can be varied. By controlling these variables and the angular velocity of the blades, a wide range of geometries and intensities of atmospheric vortices can be achieved. The apparatus is portable for use both under terrestrial atmospheric conditions and in the NASA-Ames Research Center Mars Surface Wind Tunnel facility to simulate Martian atmospheric conditions. The laboratory simulation is preferable to a numerical model because direct measurements of dust lifting threshold can be made and holds several advantages over terrestrial field measurements in that it is convenient, easily instrumented and, most importantly, can be moved to a low-pressure environment. Terrestrial field data are necessary, however, to validate the laboratory simulation as a good approximation of reality. Field measurements show that both pressure and velocity structure of the laboratory-generated vortex are similar to terrestrial dust devils. Initial threshold tests under terrestrial conditions show that the geometry of the vortex plays a key role in the angular velocity required to entrain material: smaller vortices have lower angular velocities at threshold. This is thought to be due to the smaller inflow boundary layer associated with narrow vortices and hence enhanced shear stress. However, calculations show that the shear stresses at the surface are at least two orders of magnitude less than the upward force caused by the pressure drop at the center of the vortex. This leads to the tentative conclusion that the actual particle lifting action of the `lift' force is minimal. A full program of experiments using this apparatus is under way to confirm these initial findings and a sequence of experiments under Martian conditions is being planned.

Turbomachinery noise

NASA Astrophysics Data System (ADS)

Groeneweg, John F.; Sofrin, Thomas G.; Rice, Edward J.; Gliebe, Phillip R.

1991-08-01

Summarized here are key advances in experimental techniques and theoretical applications which point the way to a broad understanding and control of turbomachinery noise. On the experimental side, the development of effective inflow control techniques makes it possible to conduct, in ground based facilities, definitive experiments in internally controlled blade row interactions. Results can now be valid indicators of flight behavior and can provide a firm base for comparison with analytical results. Inflow control coupled with detailed diagnostic tools such as blade pressure measurements can be used to uncover the more subtle mechanisms such as rotor strut interaction, which can set tone levels for some engine configurations. Initial mappings of rotor wake-vortex flow fields have provided a data base for a first generation semiempirical flow disturbance model. Laser velocimetry offers a nonintrusive method for validating and improving the model. Digital data systems and signal processing algorithms are bringing mode measurement closer to a working tool that can be frequently applied to a real machine such as a turbofan engine. On the analytical side, models of most of the links in the chain from turbomachine blade source to far field observation point have been formulated. Three dimensional lifting surface theory for blade rows, including source noncompactness and cascade effects, blade row transmission models incorporating mode and frequency scattering, and modal radiation calculations, including hybrid numerical-analytical approaches, are tools which await further application.

Development of multidisciplinary design optimization procedures for smart composite wings and turbomachinery blades

NASA Astrophysics Data System (ADS)

Jha, Ratneshwar

Multidisciplinary design optimization (MDO) procedures have been developed for smart composite wings and turbomachinery blades. The analysis and optimization methods used are computationally efficient and sufficiently rigorous. Therefore, the developed MDO procedures are well suited for actual design applications. The optimization procedure for the conceptual design of composite aircraft wings with surface bonded piezoelectric actuators involves the coupling of structural mechanics, aeroelasticity, aerodynamics and controls. The load carrying member of the wing is represented as a single-celled composite box beam. Each wall of the box beam is analyzed as a composite laminate using a refined higher-order displacement field to account for the variations in transverse shear stresses through the thickness. Therefore, the model is applicable for the analysis of composite wings of arbitrary thickness. Detailed structural modeling issues associated with piezoelectric actuation of composite structures are considered. The governing equations of motion are solved using the finite element method to analyze practical wing geometries. Three-dimensional aerodynamic computations are performed using a panel code based on the constant-pressure lifting surface method to obtain steady and unsteady forces. The Laplace domain method of aeroelastic analysis produces root-loci of the system which gives an insight into the physical phenomena leading to flutter/divergence and can be efficiently integrated within an optimization procedure. The significance of the refined higher-order displacement field on the aeroelastic stability of composite wings has been established. The effect of composite ply orientations on flutter and divergence speeds has been studied. The Kreisselmeier-Steinhauser (K-S) function approach is used to efficiently integrate the objective functions and constraints into a single envelope function. The resulting unconstrained optimization problem is solved using the Broyden-Fletcher-Goldberg-Shanno algorithm. The optimization problem is formulated with the objective of simultaneously minimizing wing weight and maximizing its aerodynamic efficiency. Design variables include composite ply orientations, ply thicknesses, wing sweep, piezoelectric actuator thickness and actuator voltage. Constraints are placed on the flutter/divergence dynamic pressure, wing root stresses and the maximum electric field applied to the actuators. Numerical results are presented showing significant improvements, after optimization, compared to reference designs. The multidisciplinary optimization procedure for the design of turbomachinery blades integrates aerodynamic and heat transfer design objective criteria along with various mechanical and geometric constraints on the blade geometry. The airfoil shape is represented by Bezier-Bernstein polynomials, which results in a relatively small number of design variables for the optimization. Thin shear layer approximation of the Navier-Stokes equation is used for the viscous flow calculations. Grid generation is accomplished by solving Poisson equations. The maximum and average blade temperatures are obtained through a finite element analysis. Total pressure and exit kinetic energy losses are minimized, with constraints on blade temperatures and geometry. The constrained multiobjective optimization problem is solved using the K-S function approach. The results for the numerical example show significant improvements after optimization.

Apparatus and process for removing a predetermined portion of reflective material from mirror

DOEpatents

Perry, Stephen J.; Steinmetz, Lloyd L.

1994-01-01

An apparatus and process are disclosed for removal of a stripe of soft reflective material of uniform width from the surface of a mirror by using a blade having a large included angle to inhibit curling of the blade during the cutting operation which could result in damage to the glass substrate of the mirror. The cutting blade is maintained at a low blade angle with respect to the mirror surface to produce minimal chipping along the cut edge and to minimize the force exerted on the coating normal to the glass surface which could deform the flat mirror. The mirror is mounted in a cutting mechanism containing a movable carriage on which the blade is mounted to provide very accurate straightness of the travel of the blade along the mirror.

Water Flow Testing and Unsteady Pressure Analysis of a Two-Bladed Liquid Oxidizer Pump Inducer

NASA Technical Reports Server (NTRS)

Schwarz, Jordan B.; Mulder, Andrew; Zoladz, Thomas

2011-01-01

The unsteady fluid dynamic performance of a cavitating two-bladed oxidizer turbopump inducer was characterized through sub-scale water flow testing. While testing a novel inlet duct design that included a cavitation suppression groove, unusual high-frequency pressure oscillations were observed. With potential implications for inducer blade loads, these high-frequency components were analyzed extensively in order to understand their origins and impacts to blade loading. Water flow testing provides a technique to determine pump performance without the costs and hazards associated with handling cryogenic propellants. Water has a similar density and Reynolds number to liquid oxygen. In a 70%-scale water flow test, the inducer-only pump performance was evaluated. Over a range of flow rates, the pump inlet pressure was gradually reduced, causing the flow to cavitate near the pump inducer. A nominal, smooth inducer inlet was tested, followed by an inlet duct with a circumferential groove designed to suppress cavitation. A subsequent 52%-scale water flow test in another facility evaluated the combined inducer-impeller pump performance. With the nominal inlet design, the inducer showed traditional cavitation and surge characteristics. Significant bearing loads were created by large side loads on the inducer during synchronous cavitation. The grooved inlet successfully mitigated these loads by greatly reducing synchronous cavitation, however high-frequency pressure oscillations were observed over a range of frequencies. Analytical signal processing techniques showed these oscillations to be created by a rotating, multi-celled train of pressure pulses, and subsequent CFD analysis suggested that such pulses could be created by the interaction of rotating inducer blades with fluid trapped in a cavitation suppression groove. Despite their relatively low amplitude, these high-frequency pressure oscillations posed a design concern due to their sensitivity to flow conditions and test scale. The amplitude and frequency of oscillations varied considerably over the pump s operating space, making it difficult to predict blade loads.

An Investigation of End-Wall Vortex Cavitation in a High Reynolds Number Axial-Flow Pump

DTIC Science & Technology

1989-08-01

groove treatment, and various winglet configurations. Notwithstanding the differences between compressor and turbine flow fields, the deep- groove...treatment and pressure-surface winglet reduced the leakage by about 15 percent as reported by Booth et al. [20]. Additional numerical and experimental...the blade appears stationary because the eye blends the individual pictures together similar to a motion picture projection. However, at lower

Analysis of an axial compressor blade vibration based on wave reflection theory

NASA Technical Reports Server (NTRS)

Owczarek, J. A.

1983-01-01

The paper describes application of the theory of wave reflection in turbomachines to rotor blade vibrations measured in an axial compressor stage. The blade vibrations analyzed could not be predicted using various flutter prediction techniques. The wave reflection theory, first advanced in 1966, is expanded, and more general equations for the rotor blade excitation frequencies are derived. The results of the analysis indicate that all examined rotor blade vibrations can be explained by forced excitations caused by reflecting waves (pressure pulses). Wave reflections between the rotor blades and both the upstream and downstream stator vanes had to be considered.

Noise from Propellers with Symmetrical Sections at Zero Blade Angle

NASA Technical Reports Server (NTRS)

Deming, A F

1937-01-01

A theory has been deduced for the "rotation noise" from a propeller with blades of symmetrical section about the chord line and set at zero blade angle. Owing to the limitation of the theory, the equations give without appreciable error only the sound pressure for cases where the wave lengths are large compared with the blade lengths. With the aid of experimental data obtained from a two-blade arrangement, an empirical relation was introduced that permitted calculation of higher harmonics. The generality of the final relation given is indicated by the fundamental and second harmonic of a four-blade arrangement.

Wind tunnel tests of rotor blade sections with replications of ice formations accreted in hover

NASA Technical Reports Server (NTRS)

Lee, J. D.; Berger, J. H.; Mcdonald, T. J.

1986-01-01

Full scale reproductions of ice accretions molded during the documentation of a hover test program were fabricated by means of epoxy castings and used for a wind tunnel test program. Surface static pressure distributions were recorded and used to evaluate lift and pitching moment increments while drag was determined by wake surveys. Through the range of the tests, corresponding to those conditions encountered in hover and in flat pitch, integration of the pressure distributions showed negligible changes in lift and in pitching moment, but the drag was significantly increased.

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.)

Unsteady inflow effects on the wake shed from a high-lift LPT blade subjected to boundary layer laminar separation

NASA Astrophysics Data System (ADS)

Satta, Francesca; Ubaldi, Marina; Zunino, Pietro

2012-04-01

An experimental investigation on the near and far wake of a cascade of high-lift low-pressure turbine blades subjected to boundary layer separation over the suction side surface has been carried out, under steady and unsteady inflows. Two Reynolds number conditions, representative of take-off/landing and cruise operating conditions of the real engine, have been tested. The effect of upstream wake-boundary layer interaction on the wake shed from the profile has been investigated in a three-blade large-scale linear turbine cascade. The comparison between the wakes shed under steady and unsteady inflows has been performed through the analysis of mean velocity and Reynolds stress components measured at midspan of the central blade by means of a two-component crossed miniature hot-wire probe. The wake development has been analyzed in the region between 2% and 100% of the blade chord from the central blade trailing edge, aligned with the blade exit direction. Wake integral parameters, half-width and maximum velocity defects have been evaluated from the mean velocity distributions to quantify the modifications induced on the vane wake by the upstream wake. Moreover the thicknesses of the two wake shear layers have been considered separately in order to identify the effects of Reynolds number and incoming flow on the wake shape. The self-preserving state of the wake has been looked at, taking into account the different thicknesses of the two shear layers. The evaluation of the power density spectra of the velocity fluctuations allowed the study of the wake unsteady behavior, and the detection of the effects induced by the different operating conditions on the trailing edge vortex shedding.

Characterizing Operational Performance of Rotary Subwoofer Loudspeaker

DTIC Science & Technology

2017-10-01

rotation rate of the fan blades and the amplitude of the articulation of the fan blades to maximum the amplitude on the resulting acoustic pressure...on a stationary cart relatively close to the subwoofer. The speaker was attached to a piece of plywood with a circular hole for the fan blade that...supporting technology Device Specifications Eminent Technology Rotary Subwoofer TRW-17 Blade Number: 5 (300 rpm) Amplifier Requirement: 150 watts

Low noise propeller design using numerical analysis

NASA Astrophysics Data System (ADS)

Humpert, Bryce

The purpose of this study is to explore methods for reducing aircraft propeller noise with minimal losses in performance. Using numerical analysis, a standard two blade propeller configuration was taken from experiments conducted by Patrick, Finn, and Stich, and implemented into the numerical code XROTOR. The blade design modifications that were investigated to achieve the proposed goals include: increasing the number of blades, adjusting the chord length, beta distribution, radius of the blade, airfoil shape, and operating RPM. In order to determine the optimal blade design, a baseline case is first developed and the parameters listed earlier are then varied to create a new propeller design that reduces the sound pressure level (SPL) while maintaining performance levels within a predetermined range of the original specifications. From the analysis, the most significant improvements observed in lowering the acoustic signature are dominated by operating rpm and blade radius. A three-, four-, and five-blade configuration was developed that reduced the SPL generated by the propeller during cruise flight conditions. The optimum configuration that produced the greatest SPL reduction was the five-blade configuration. The resulting sound pressure level was reduced from the original 77 dB at 1000' ft above ground level (AGL), to 54 dB at 1000' AGL while remaining within 1.4% of the original thrust and efficiency.

Steady State Structural Analysis of High Pressure Gas Turbine Blade using Finite Element Analysis

NASA Astrophysics Data System (ADS)

Mazarbhuiya, Hussain Mahamed Sahed Mostafa; Murari Pandey, Krishna

2017-08-01

In gas turbines the major portion of performance dependency lies upon turbine blade design. Turbine blades experience very high centrifugal, axial and tangential force during power generation. While withstanding these forces blades undergo elongation. Different methods have proposed for better enhancement of the mechanical properties of blade to withstand in extreme condition. Present paper describes the stress and elongation for blades having properties of different materials. Steady state structural analysis have performed in the present work for different materials (In 625, In 718, In 738, In 738 LC, MAR M246, Ni-Cr, Ti-alloy, Ti-Al, Ti-T6, U500). Remarkable finding is that the root of the blade is subjected to maximum stress for all blade materials and the blade made of MAR M246 has less stress and deformation among all other blade materials which can be selected as a suitable material for gas turbine blade.

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.

Research and development of asymmetrical heat transfer augmentation method in radial channels of blades for high temperature gas turbines

NASA Astrophysics Data System (ADS)

Shevchenko, I. V.; Rogalev, A. N.; Garanin, I. V.; Vegera, A. N.; Kindra, V. O.

2017-11-01

The serpentine-like one and half-pass cooling channel systems are primarily used in blades fabricated by the lost-wax casting process. The heat transfer turbulators like cross-sectional or angled ribs used in channels of the midchord region failed to eliminate the temperature irregularity from the suction and pressure sides, which is reaching 200°C for a first stage blade of the high-pressure turbine for an aircraft engine. This paper presents the results of a numerical and experimental test of an advanced heat transfer augmentation system in radial channels developed for alignment of the temperature field from the suction and pressure sides. A numerical simulation of three-dimensional coolant flow for a wide range of Reynolds numbers was carried out using ANSYS CFX software. Effect of geometrical parameters on the heat removal asymmetry was determined. The test results of a blade with the proposed intensification system conducted in a liquid-metal thermostat confirmed the accuracy of calculations. Based on the experimental data, the dependencies for calculation of heat transfer coefficients to the cooling air in the blade studied were obtained.

Aero-thermal investigation of a highly loaded transonic linear turbine guide vane cascade. A test case for inviscid and viscous flow computations

NASA Astrophysics Data System (ADS)

Arts, T.; Lambertderouvroit, M.; Rutherford, A. W.

1990-09-01

An experimental aerothermal investigation of a highly loaded transonic turbine nozzle guide vane mounted in a linear cascade arrangement is presented. The measurements were performed in a short duration isentropic light piston compression tube facility, allowing a correct simulation of Mach and Reynolds numbers as well as of the gas to wall temperature ratio compared to the values currently observed in modern aeroengines. The experimental program consisted of the following: (1) flow periodicity checks by means of wall static pressure measurements and Schlieren flow visualizations; (2) blade velocity distribution measurements by means of static pressure tappings; (3) blade convective heat transfer measurements by means of static pressure tappings; (4) blade convective heat transfer measurements by means of platinium thin films; (5) downstream loss coefficient and exit flow angle determinations by using a new fast traversing mechanism; and (6) free stream turbulence intensity and spectrum measurements. These different measurements were performed for several combinations of the free stream flow parameters looking at the relative effects on the aerodynamic blade performance and blade convective heat transfer of Mach number, Reynolds number, and freestream turbulence intensity.

Identifying Major Transitions in the Evolution of Lithic Cutting Edge Production Rates

PubMed Central

Clarkson, Chris

2016-01-01

The notion that the evolution of core reduction strategies involved increasing efficiency in cutting edge production is prevalent in narratives of hominin technological evolution. Yet a number of studies comparing two different knapping technologies have found no significant differences in edge production. Using digital analysis methods we present an investigation of raw material efficiency in eight core technologies broadly representative of the long-term evolution of lithic technology. These are bipolar, multiplatform, discoidal, biface, Levallois, prismatic blade, punch blade and pressure blade production. Raw material efficiency is assessed by the ratio of cutting edge length to original core mass. We also examine which flake attributes contribute to maximising raw material efficiency, as well as compare the difference between expert and intermediate knappers in terms of cutting edge produced per gram of core. We identify a gradual increase in raw material efficiency over the broad sweep of lithic technological evolution. The results indicate that the most significant transition in efficiency likely took place with the introduction of small foliate biface, Levallois and prismatic blade knapping, all introduced in the Middle Stone Age / Middle Palaeolithic among early Homo sapiens and Neanderthals. This suggests that no difference in raw material efficiency existed between these species. With prismatic blade technology securely dated to the Middle Palaeolithic, by including the more recent punch and pressure blade technology our results dispel the notion that the transition to the Upper Palaeolithic was accompanied by an increase in efficiency. However, further increases in cutting edge efficiency are evident, with pressure blades possessing the highest efficiency in this study, indicating that late/epi-Palaeolithic and Neolithic blade technologies further increased efficiency. PMID:27936135

Identifying Major Transitions in the Evolution of Lithic Cutting Edge Production Rates.

PubMed

Muller, Antoine; Clarkson, Chris

2016-01-01

The notion that the evolution of core reduction strategies involved increasing efficiency in cutting edge production is prevalent in narratives of hominin technological evolution. Yet a number of studies comparing two different knapping technologies have found no significant differences in edge production. Using digital analysis methods we present an investigation of raw material efficiency in eight core technologies broadly representative of the long-term evolution of lithic technology. These are bipolar, multiplatform, discoidal, biface, Levallois, prismatic blade, punch blade and pressure blade production. Raw material efficiency is assessed by the ratio of cutting edge length to original core mass. We also examine which flake attributes contribute to maximising raw material efficiency, as well as compare the difference between expert and intermediate knappers in terms of cutting edge produced per gram of core. We identify a gradual increase in raw material efficiency over the broad sweep of lithic technological evolution. The results indicate that the most significant transition in efficiency likely took place with the introduction of small foliate biface, Levallois and prismatic blade knapping, all introduced in the Middle Stone Age / Middle Palaeolithic among early Homo sapiens and Neanderthals. This suggests that no difference in raw material efficiency existed between these species. With prismatic blade technology securely dated to the Middle Palaeolithic, by including the more recent punch and pressure blade technology our results dispel the notion that the transition to the Upper Palaeolithic was accompanied by an increase in efficiency. However, further increases in cutting edge efficiency are evident, with pressure blades possessing the highest efficiency in this study, indicating that late/epi-Palaeolithic and Neolithic blade technologies further increased efficiency.

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.

Streamwise Vortices on the Convex Surfaces of Circular Cylinders and Turbomachinery Blading

NASA Technical Reports Server (NTRS)

Gostelow, Paul

2010-01-01

In assessing the results please recall that the Mach number regimes and model geometries differ considerably. Selection of the radius of curvature at the 10% chord location is consistent but arbitrary, although it does seem representative for most blades and gives a good fit for the results. Measured spanwise wavelengths of the periodic vortex arrays on blading are predicted well by the Kestin and Wood theory. If this behavior is at all common it could have implications for turbine aerodynamic and blade cooling design. The outcome is to establish that organized streamwise vorticity may occur more frequently on convex surfaces, such as turbine blade suction surfaces, than hitherto appreciated. Investigations and predictions of flow behavior should be extended to encompass that possibility.

Turbine blade and vane heat flux sensor development, phase 2

NASA Technical Reports Server (NTRS)

Atkinson, W. H.; Cyr, M. A.; Strange, R. R.

1985-01-01

The development of heat flux sensors for gas turbine blades and vanes and the demonstration of heat transfer measurement methods are reported. The performance of the heat flux sensors was evaluated in a cylinder in cross flow experiment and compared with two other heat flux measurement methods, the slug calorimeter and a dynamic method based on fluctuating gas and surface temperature. Two cylinders, each instrumented with an embedded thermocouple sensor, a Gardon gauge, and a slug calorimeter, were fabricated. Each sensor type was calibrated using a quartz lamp bank facility. The instrumented cylinders were then tested in an atmospheric pressure combustor rig at conditions up to gas stream temperatures of 1700K and velocities to Mach 0.74. The test data are compared to other measurements and analytical prediction.

Global Search of a Three-dimensional Low Solidity Circular Cascade Diffuser for Centrifugal Blowers by Meta-model Assisted Optimization

NASA Astrophysics Data System (ADS)

Sakaguchi, Daisaku; Sakue, Daiki; Tun, Min Thaw

2018-04-01

A three-dimensional blade of a low solidity circular cascade diffuser in centrifugal blowers is designed by means of a multi-point optimization technique. The optimization aims at improving static pressure coefficient at a design point and at a small flow rate condition. Moreover, a clear definition of secondary flow expressed by positive radial velocity at hub side is taken into consideration in constraints. The number of design parameters for three-dimensional blade reaches to 10 in this study, such as a radial gap, a radial chord length and mean camber angle distribution of the LSD blade with five control points, control point between hub and shroud with two design freedom. Optimization results show clear Pareto front and selected optimum design shows good improvement of pressure rise in diffuser at small flow rate conditions. It is found that three-dimensional blade has advantage to stabilize the secondary flow effect with improving pressure recovery of the low solidity circular cascade diffuser.

SSME single crystal turbine blade dynamics

NASA Technical Reports Server (NTRS)

Moss, Larry A.; Smith, Todd E.

1987-01-01

A study was performed to determine the dynamic characteristics of the Space Shuttle main engine high pressure fuel turbopump (HPFTP) blades made of single crystal (SC) material. The first and second stage drive turbine blades of HPFTP were examined. The nonrotating natural frequencies were determined experimentally and analytically. The experimental results of the SC second stage blade were used to verify the analytical procedures. The analytical study examined the SC first stage blade natural frequencies with respect to crystal orientation at typical operating conditions. The SC blade dynamic response was predicted to be less than the directionally solidified blade. Crystal axis orientation optimization indicated the third mode interference will exist in any SC orientation.

High speed turboprop aeroacoustic study (single rotation). Volume 1: Model development

NASA Technical Reports Server (NTRS)

Whitfield, C. E.; Gliebe, P. R.; Mani, R.; Mungur, P.

1989-01-01

A frequency-domain noncompact-source theory for the steady loading and volume-displacement (thickness) noise of high speed propellers has been developed and programmed. Both near field and far field effects have been considered. The code utilizes blade surface pressure distributions obtained from three-dimensional nonlinear aerodynamic flow field analysis programs as input for evaluating the steady loading noise. Simplified mathematical models of the velocity fields induced at the propeller disk by nearby wing and fuselage surfaces and by angle-of-attack operation have been developed to provide estimates of the unsteady loading imposed on the propeller by these potential field type interactions. These unsteady blade loadings have been coupled to a chordwise compact propeller unsteady loading noise model to provide predictions of unsteady loading noise caused by these installation effects. Finally, an analysis to estimate the corrections to be applied to the free-field noise predictions in order to arrive at the measurable fuselage sound pressure levels has been formulated and programmed. This analysis considers the effects of fuselage surface reflection and diffraction together with surface boundary layer refraction. The steady loading and thickness model and the unsteady loading model have been verified using NASA-supplied data for the SR-2 and SR-3 model propfans. In addition, the steady loading and thickness model has been compared with data from the SR-6 model propfan. These theoretical models have been employed in the evaluation of the SR-7 powered Gulfstream aircraft in terms of noise characteristics at representative takeoff, cruise, and approach operating conditions. In all cases, agreement between theory and experiment is encouraging.

A Neural Network Aero Design System for Advanced Turbo-Engines

NASA Technical Reports Server (NTRS)

Sanz, Jose M.

1999-01-01

An inverse design method calculates the blade shape that produces a prescribed input pressure distribution. By controlling this input pressure distribution the aerodynamic design objectives can easily be met. Because of the intrinsic relationship between pressure distribution and airfoil physical properties, a neural network can be trained to choose the optimal pressure distribution that would meet a set of physical requirements. The neural network technique works well not only as an interpolating device but also as an extrapolating device to achieve blade designs from a given database. Two validating test cases are discussed.

Statistical analysis of 59 inspected SSME HPFTP turbine blades (uncracked and cracked)

NASA Technical Reports Server (NTRS)

Wheeler, John T.

1987-01-01

The numerical results of statistical analysis of the test data of Space Shuttle Main Engine high pressure fuel turbopump second-stage turbine blades, including some with cracks are presented. Several statistical methods use the test data to determine the application of differences in frequency variations between the uncracked and cracked blades.

78 FR 40063 - Airworthiness Directives; Erickson Air-Crane Incorporated Helicopters (Type Certificate...

Federal Register 2010, 2011, 2012, 2013, 2014

2013-07-03

... Sikorsky Model S-64E helicopters. The AD requires repetitive checks of the Blade Inspection Method (BIM... and check procedures for BIM blades installed on the Model S-64F helicopters. Several blade spars with a crack emanating from corrosion pits and other damage have been found because of BIM pressure...

Numerical Investigation of Compressor Non-Synchronous Vibration with Full Annulus Rotor-Stator Interaction

NASA Astrophysics Data System (ADS)

Espinal, Daniel

The objective of this research is to investigate and confirm the periodicity of the Non-Synchronous Vibration (NSV) mechanism of a GE axial compressor with a full-annulus simulation. A second objective is to develop a high fidelity single-passage tool with time-accurate unsteady capabilities able to capture rotor-stator interactions and NSV excitation response. A high fidelity methodology for axial turbomachinery simulation is developed using the low diffusion shock-capturing Riemann solver with high order schemes, the Spalart-Allmaras turbulence closure model, the fully conservative unsteady sliding BC for rotor-stator interaction with extension to full-annulus and single-passage configurations, and the phase lag boundary conditions applied to rotor-stator interface and circumferential BC. A URANS solver is used and captures the NSV flow excitation frequency of 2439 Hz, which agrees reasonably well with the measured NSV frequency of 2600 Hz from strain gage test data. It is observed that the circumferentially traveling vortex formed in the vicinity of the rotor tip propagates at the speed of a non-engine order frequency and causes the NSV. The vortex travels along the suction surface of the blade and crosses the passage outlet near blade trailing edge. Such a vortex motion trajectory repeats in each blade passage and generates two low pressure regions due to the vortex core positions, one at the leading edge and one at the trailing edge, both are oscillating due to the vortex coming and leaving. These two low pressure regions create a pair of coupling forces that generates a torsion moment causing NSV. The full-annulus simulation shows that the circumferentially traveling vortex has fairly periodical behavior and is a full annulus structure. Also, frequencies below the NSV excitation frequency of 2439 Hz with large amplitudes in response to flow-separation related phenomena are present. This behavior is consistent with experimental measurements. For circumferentially averaged parameters like total pressure ratio, NSV is observed to have an effect, particularly at radial locations above 70% span. Therefore, to achieve similar or better total pressure ratio a design with a smaller loading of the upper blade span and a higher loading of the mid blade spans should be considered. A fully-conservative sliding interface boundary condition (BC) is implemented with phase-lag capabilities using the Direct Store method for single-passage simulations. Also Direct Store phase-lag was applied to the circumferential BCs to enforce longer disturbance wavelengths. The unsteady simulation using single-blade-passage with periodic BC for an inlet guide vane (IGV)-rotor configuration captures a 2291 Hz NSV excitation frequency and an IGV-rotor-stator configuration predicts a 2365 Hz NSV excitation frequency with a significantly higher amplitude above 90% span. This correlates closely to the predicted NSV excitation frequency of 2439 Hz for the full-annulus configuration. The two-blade-row configuration exhibits the same vortex structures captured in the full-annulus study. The three-blade-row configuration only captures a tip vortex shedding at the leading edge, which can be attributed to the reflective nature of the BCs causing IGV-rotor-stator interactions to be augmented, becoming dominant and shifting NSV excitation response to engine order regime. Phase-lag simulations with a Nodal Diameter (ND) of 5 is enforced for the circumferential BCs for the three-blade-row configuration, and the results exactly matched the frequency response and flow structures of the periodic simulation, illustrating the small effect that phase-lag has on strongly periodic flow disturbances. A ND of 7 is enforced at the sliding interface, however the NSV excitation completely disappears and only the wake propagation from IGV-Rotor-Stator interactions are captured. Rotor blade passage exhibits a circumferentially travelling vortex similar to those observed in the full-annulus and two-blade-row simulations. This can occur when the rotating instability responsible for the NSV no longer maintains a pressure variation with a characteristic frequency signature as it rotates relative to the rotor rotation, and now has become the beginning of a spike-type stall cell. In this scenario the travelling vortex has become evidence of part-stall of the upper spans of the rotor blade, but stalling is contained maintaining stable operation. In conclusion, an efficient method of capturing NSV excitation has been proposed in a high-fidelity manner, where only 2% of the computational resources used in a full-annulus simulation are required for an accurate single-blade-passage multi-stage simulation.

Performance of a low-pressure fan stage with reverse flow

NASA Technical Reports Server (NTRS)

Moore, R. D.; Lewis, G. W., Jr.; Tysl, E. R.

1976-01-01

The reverse flow aerodynamic performance of a 51-centimeter-diameter fan stage is presented. The stage was tested with the variable pitch rotor blades set through feather at -75 deg, -80 deg, and -85 deg from design setting angle. Of the three tested the stage with the rotor blades set at -75 deg exhibited the highest pressure ratio and highest flow. For all three configurations, there was little or no flow in the inner third of the exit passage due to the rotor blade being almost perpendicular to the axial direction in the hub region.

The Influence of Blade Profile and Slots on the Performance of a Centrifugal Impeller (Influence du Profil de Pales et Fentes sur la Performance d’un Ventilateur Centrifuge),

DTIC Science & Technology

1980-01-01

instability at low flows was investigated. The major cause was found to be flow detachment from the impeller vanes . ] Slotted blades were found to be...Characteristics (70 RPM), Variant (D- Aerofoil Blades , Sets VB1, 2 & 3...12 7 Pressure Rise - Flow Characteristics (70 RPM), Variant () - Slotted Aerofoil Blades , Sets VB1, 2 & 3

Probabilistic structural analysis of aerospace components using NESSUS

NASA Technical Reports Server (NTRS)

Shiao, Michael C.; Nagpal, Vinod K.; Chamis, Christos C.

1988-01-01

Probabilistic structural analysis of a Space Shuttle main engine turbopump blade is conducted using the computer code NESSUS (numerical evaluation of stochastic structures under stress). The goal of the analysis is to derive probabilistic characteristics of blade response given probabilistic descriptions of uncertainties in blade geometry, material properties, and temperature and pressure distributions. Probability densities are derived for critical blade responses. Risk assessment and failure life analysis is conducted assuming different failure models.

Integrated Heat Exchange For Recuperation In Gas Turbine Engines

DTIC Science & Technology

2016-12-01

exchange system within the engine using existing blade surfaces to extract and insert heat. Due to the highly turbulent and transient flow, heat...transfer coefficients in turbomachinery are extremely high, making this possible. Heat transfer between the turbine and compressor blade surfaces could be...exchange system within the engine using existing blade surfaces to extract and insert heat. Due to the highly turbulent and transient flow, heat transfer

Study on temperature measurement of gas turbine blade based on analysis of error caused by the reflected radiation and emission angle

NASA Astrophysics Data System (ADS)

Li, Dong; Feng, Chi; Gao, Shan; Chen, Liwei; Daniel, Ketui

2018-06-01

Accurate measurement of gas turbine blade temperature is of great significance as far as blade health monitoring is concerned. An important method for measuring this temperature is the use of a radiation pyrometer. In this research, error of the pyrometer caused by reflected radiation from the surfaces surrounding the target and the emission angle of the target was analyzed. Important parameters for this analysis were the view factor between interacting surfaces, spectral directional emissivity, pyrometer operating wavelength and the surface temperature distribution on the blades and the vanes. The interacting surface of the rotor blade and the vane models used were discretized using triangular surface elements from which contour integral was used to calculate the view factor between the surface elements. Spectral directional emissivities were obtained from an experimental setup of Ni based alloy samples. A pyrometer operating wavelength of 1.6 μm was chosen. Computational fluid dynamics software was used to simulate the temperature distribution of the rotor blade and the guide vane based on the actual gas turbine input parameters. Results obtained in this analysis show that temperature error introduced by reflected radiation and emission angle ranges from  ‑23 K to 49 K.

High speed propeller performance and noise predictions at takeoff/landing conditions

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Woodward, R. P.; Groeneweg, J. F.

1988-01-01

The performance and noise of a high speed SR-7A model propeller under takeoff/landing conditions are considered. The blade loading distributions are obtained by solving the three-dimensional Euler equations and the sound pressure levels are computed using a time domain approach. At the nominal takeoff operating point, the blade sections near the hub are lightly or negatively loaded. The chordwise loading distributions are distinctly different from those of cruise conditions. The noise of the SR-7A model propeller at takeoff is dominated by the loading noise, similar to that at cruise conditions. The waveforms of the acoustic pressure signature are nearly sinusoidal in the plane of the propeller. The computed directivity of the blade passing frequency tone agrees fairly well with the data at nominal takeoff blade angle.

High speed propeller performance and noise predictions at takeoff/landing conditions

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Woodward, R. P.; Groeneweg, J. F.

1987-01-01

The performance and noise of a high speed SR-7A model propeller under takeoff/landing conditions are considered. The blade loading distributions are obtained by solving the three-dimensional Euler equations and the sound pressure levels are computed using a time domain approach. At the nominal takeoff operating point, the blade sections near the hub are lightly or negatively loaded. The chordwise loading distributions are distinctly different from those of cruise conditions. The noise of the SR-7A model propeller at takeoff is dominated by the loading noise, similar to that at cruise conditions. The waveforms of the acoustic pressure signature are nearly sinusoidal in the plane of the propeller. The computed directivity of the blade passing frequency tone agrees fairly well with the data at nominal takeoff blade angle.

BladeCAD: An Interactive Geometric Design Tool for Turbomachinery Blades

NASA Technical Reports Server (NTRS)

Miller, Perry L., IV; Oliver, James H.; Miller, David P.; Tweedt, Daniel L.

1996-01-01

A new metthodology for interactive design of turbomachinery blades is presented. Software implementation of the meth- ods provides a user interface that is intuitive to aero-designers while operating with standardized geometric forms. The primary contribution is that blade sections may be defined with respect to general surfaces of revolution which may be defined to represent the path of fluid flow through the turbomachine. The completed blade design is represented as a non-uniform rational B-spline (NURBS) surface and is written to a standard IGES file which is portable to most design, analysis, and manufacturing applications.

Laser hardening techniques on steam turbine blade and application

NASA Astrophysics Data System (ADS)

Yao, Jianhua; Zhang, Qunli; Kong, Fanzhi; Ding, Qingming

Different laser surface hardening techniques, such as laser alloying and laser solution strengthening were adopted to perform modification treatment on the local region of inset edge for 2Cr13 and 17-4PH steam turbine blades to prolong the life of the blades. The microstructures, microhardness and anti-cavitation properties were investigated on the blades after laser treatment. The hardening mechanism and technique adaptability were researched. Large scale installation practices confirmed that the laser surface modification techniques are safe and reliable, which can improve the properties of blades greatly with advantages of high automation, high quality, little distortion and simple procedure.

Thermal Imaging of Medical Saw Blades and Guides

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

Dinwiddie, Ralph Barton; Steffner, Thomas E

2007-01-01

Better Than New, LLC., has developed a surface treatment to reduce the friction and wear of orthopedic saw blades and guides. The medical saw blades were thermally imaged while sawing through fresh animal bone and an IR camera was used to measure the blade temperature as it exited the bone. The thermal performance of as-manufactured saw blades was compared to surface-treated blades, and a freshly used blade was used for temperature calibration purposes in order to account for any emissivity changes due to organic transfer layers. Thermal imaging indicates that the treated saw blades cut faster and cooler than untreatedmore » blades. In orthopedic surgery, saw guides are used to perfectly size the bone to accept a prosthesis. However, binding can occur between the blade and guide because of misalignment. This condition increases the saw blade temperature and may result in tissue damage. Both treated ad untreated saw guides were also studied. The treated saw guide operated at a significantly lower temperature than untreated guide. Saw blades and guides that operate at a cooler temperature are expected to reduce the amount of tissue damage (thermal necrosis) and may reduce the number of post-operative complications.« less

Investigation of the Unsteady Total Pressure Profile Corresponding to Counter-Rotating Vortices in an Internal Flow Application

NASA Astrophysics Data System (ADS)

Gordon, Kathryn; Morris, Scott; Jemcov, Aleksandar; Cameron, Joshua

2013-11-01

The interaction of components in a compressible, internal flow often results in unsteady interactions between the wakes and moving blades. A prime example in which this flow feature is of interest is the interaction between the downstream rotor blades in a transonic axial compressor with the wake vortices shed from the upstream inlet guide vane (IGV). Previous work shows that a double row of counter-rotating vortices convects downstream into the rotor passage as a result of the rotor blade bow shock impinging on the IGV. The rotor-relative time-mean total pressure distribution has a region of high total pressure corresponding to the pathline of the vortices. The present work focuses on the relationship between the magnitude of the time-mean rotor-relative total pressure profile and the axial spacing between the IGV and the rotor. A survey of different axial gap sizes is performed in a two-dimensional computational study to obtain the sensitivity of the pressure profile amplitude to IGV-rotor axial spacing.

Acoustic performance of low pressure axial fan rotors with different blade chord length and radial load distribution

NASA Astrophysics Data System (ADS)

Carolus, Thomas

The paper examines the acoustic and aerodynamic performance of low-pressure axial fan rotors with a hub/tip ratio of 0.45. Six rotors were designed for the same working point by means of the well-known airfoil theory. The condition of an equilibrium between the static pressure gradient and the centrifugal forces is maintained. All rotors have unequally spaced blades to diminish tonal noise. The rotors are tested in a short cylindrical housing without guide vanes. All rotors show very similar flux-pressure difference characteristics. The peak efficiency and the noise performance is considerably influenced by the chosen blade design. The aerodynamically and acoustically optimal rotor is the one with the reduced load at the hub and increased load in the tip region under satisfied equilibrium conditions. It runs at the highest aerodynamic efficiency, and its noise spectrum is fairly smooth. The overall sound pressure level of this rotor is up to 8 dB (A) lower compared to the other rotors under consideration.

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.

A Rapid Method to Achieve Aero-Engine Blade Form Detection

PubMed Central

Sun, Bin; Li, Bing

2015-01-01

This paper proposes a rapid method to detect aero-engine blade form, according to the characteristics of an aero-engine blade surface. This method first deduces an inclination error model in free-form surface measurements based on the non-contact laser triangulation principle. Then a four-coordinate measuring system was independently developed, a special fixture was designed according to the blade shape features, and a fast measurement of the blade features path was planned. Finally, by using the inclination error model for correction of acquired data, the measurement error that was caused by tilt form is compensated. As a result the measurement accuracy of the Laser Displacement Sensor was less than 10 μm. After the experimental verification, this method makes full use of optical non-contact measurement fast speed, high precision and wide measuring range of features. Using a standard gauge block as a measurement reference, the coordinate system conversion data is simple and practical. It not only improves the measurement accuracy of the blade surface, but also its measurement efficiency. Therefore, this method increases the value of the measurement of complex surfaces. PMID:26039420

A rapid method to achieve aero-engine blade form detection.

PubMed

Sun, Bin; Li, Bing

2015-06-01

This paper proposes a rapid method to detect aero-engine blade form, according to the characteristics of an aero-engine blade surface. This method first deduces an inclination error model in free-form surface measurements based on the non-contact laser triangulation principle. Then a four-coordinate measuring system was independently developed, a special fixture was designed according to the blade shape features, and a fast measurement of the blade features path was planned. Finally, by using the inclination error model for correction of acquired data, the measurement error that was caused by tilt form is compensated. As a result the measurement accuracy of the Laser Displacement Sensor was less than 10 μm. After the experimental verification, this method makes full use of optical non-contact measurement fast speed, high precision and wide measuring range of features. Using a standard gauge block as a measurement reference, the coordinate system conversion data is simple and practical. It not only improves the measurement accuracy of the blade surface, but also its measurement efficiency. Therefore, this method increases the value of the measurement of complex surfaces.

An experimental investigation of the unsteady response of a stator located downstream of a propeller ingesting broadband turbulence

NASA Astrophysics Data System (ADS)

Lynch, Denis Aloysius, III

This experimental investigation examined the unsteady response of a stator located downstream of a four- or ten-bladed propeller encountering broadband turbulence. The response is manifested in a radiated acoustic field which can be directly attributed to the unsteady surface pressure loading on the stator by the turbulent flowfield. In order to characterize the unsteady response of the stator, a thorough analysis of the turbulent flowfield downstream of the propeller was completed. The analysis of the turbulent flowfield is organized in a manner which reflects the causal relationship between influences on the flowfield and the evolution of the flowfield itself. Mathematical models for each of these contributions, including the broadband and periodic contributions of the propeller wakes and modification of the inflow turbulence by the propeller, are presented and analyzed. A further mathematical model involving the prediction of correlation length scale aids in the accurate prediction of the radiated acoustic pressure based solely on fundamental turbulent flowfield measurements. Unsteady surface pressure measurements, originally intended to provide additional information about the response of the stator as it relates to the incoming flowfield, were found to be heavily contaminated by vibrational effects. Therefore, techniques involving cross-correlation measurements are developed to mathematically isolate the unsteady pressure signal. The success of these techniques suggests the strong possibility of future application in this area. Finally, the mathematical models developed to describe the flowfield downstream of the propeller are applied to the case of a twenty-bladed propeller. This case was selected due to the anticipated increased levels of modification of the inflow turbulence. Results provide further evidence that this complex flowfield may be fully and accurately represented using simple mathematical models supported by baseline empirical information.

Aerodynamic stiffness of an unbound eccentric whirling centrifugal impeller with an infinite number of blades

NASA Technical Reports Server (NTRS)

Allaire, P. E.; Branagan, L. A.; Kocur, J. A.

1982-01-01

An unbounded eccentric centrifugal impeller with an infinite number of log spiral blades undergoing synchronous whirling in an incompressible fluid is considered. The forces acting on it due to coriolis forces, centripetal forces, changes in linear momentum, changes in pressure due to rotating and changes in pressure due to changes in linear momentum are evaluated.

Bedsores (Pressure Sores)

MedlinePlus

... over the following sites: Tailbone or buttocks Shoulder blades and spine Backs of arms and legs where ... following: Back or sides of the head Shoulder blades Hip, lower back or tailbone Heels, ankles and ...

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 ...

Aerodynamic design and analysis of small horizontal axis wind turbine blades

NASA Astrophysics Data System (ADS)

Tang, Xinzi

This work investigates the aerodynamic design and analysis of small horizontal axis wind turbine blades via the blade element momentum (BEM) based approach and the computational fluid dynamics (CFD) based approach. From this research, it is possible to draw a series of detailed guidelines on small wind turbine blade design and analysis. The research also provides a platform for further comprehensive study using these two approaches. The wake induction corrections and stall corrections of the BEM method were examined through a case study of the NREL/NASA Phase VI wind turbine. A hybrid stall correction model was proposed to analyse wind turbine power performance. The proposed model shows improvement in power prediction for the validation case, compared with the existing stall correction models. The effects of the key rotor parameters of a small wind turbine as well as the blade chord and twist angle distributions on power performance were investigated through two typical wind turbines, i.e. a fixed-pitch variable-speed (FPVS) wind turbine and a fixed-pitch fixed-speed (FPFS) wind turbine. An engineering blade design and analysis code was developed in MATLAB to accommodate aerodynamic design and analysis of the blades.. The linearisation for radial profiles of blade chord and twist angle for the FPFS wind turbine blade design was discussed. Results show that, the proposed linearisation approach leads to reduced manufacturing cost and higher annual energy production (AEP), with minimal effects on the low wind speed performance. Comparative studies of mesh and turbulence models in 2D and 3D CFD modelling were conducted. The CFD predicted lift and drag coefficients of the airfoil S809 were compared with wind tunnel test data and the 3D CFD modelling method of the NREL/NASA Phase VI wind turbine were validated against measurements. Airfoil aerodynamic characterisation and wind turbine power performance as well as 3D flow details were studied. The detailed flow characteristics from the CFD modelling are quantitatively comparable to the measurements, such as blade surface pressure distribution and integrated forces and moments. It is confirmed that the CFD approach is able to provide a more detailed qualitative and quantitative analysis for wind turbine airfoils and rotors..

Flow range enhancement by secondary flow effect in low solidity circular cascade diffusers

NASA Astrophysics Data System (ADS)

Sakaguchi, Daisaku; Tun, Min Thaw; Mizokoshi, Kanata; Kishikawa, Daiki

2014-08-01

High-pressure ratio and wide operating range are highly required for compressors and blowers. The technical issue of the design is achievement of suppression of flow separation at small flow rate without deteriorating the efficiency at design flow rate. A numerical simulation is very effective in design procedure, however, cost of the numerical simulation is generally high during the practical design process, and it is difficult to confirm the optimal design which is combined with many parameters. A multi-objective optimization technique is the idea that has been proposed for solving the problem in practical design process. In this study, a Low Solidity circular cascade Diffuser (LSD) in a centrifugal blower is successfully designed by means of multi-objective optimization technique. An optimization code with a meta-model assisted evolutionary algorithm is used with a commercial CFD code ANSYS-CFX. The optimization is aiming at improving the static pressure coefficient at design point and at low flow rate condition while constraining the slope of the lift coefficient curve. Moreover, a small tip clearance of the LSD blade was applied in order to activate and to stabilize the secondary flow effect at small flow rate condition. The optimized LSD blade has an extended operating range of 114 % towards smaller flow rate as compared to the baseline design without deteriorating the diffuser pressure recovery at design point. The diffuser pressure rise and operating flow range of the optimized LSD blade are experimentally verified by overall performance test. The detailed flow in the diffuser is also confirmed by means of a Particle Image Velocimeter. Secondary flow is clearly captured by PIV and it spreads to the whole area of LSD blade pitch. It is found that the optimized LSD blade shows good improvement of the blade loading in the whole operating range, while at small flow rate the flow separation on the LSD blade has been successfully suppressed by the secondary flow effect.

Effect of the Axial Spacing between Vanes and Blades on a Transonic Gas Turbine Performance and Blade Loading

NASA Astrophysics Data System (ADS)

Chang, Dongil; Tavoularis, Stavros

2013-03-01

Unsteady numerical simulations have been conducted to investigate the effect of axial spacing between the stator vanes and the rotor blades on the performance of a transonic, single-stage, high-pressure, axial turbine. Three cases were considered, the normal case, which is based on the geometry of a commercial jet engine and has an axial spacing at 50% blade span equal to 42% of the vane axial chord, as well as two other cases with axial spacings equal to 31 and 52% vane axial chords, respectively. Present interest has focused on the effect of axial gap size on the instantaneous and time-averaged flows as well as on the blade loading and the turbine performance. Decreasing the gap size reduced the pressure and increased the Mach number in the core flows in the gap region. However, the flows near the two endwalls did not follow monotonic trends with the gap size change; instead, the Mach numbers for both the small gap and the large gap cases were lower than that for the normal case. This Mach number decrease was attributed to increased turbulence due to the increased wake strength for the small gap case and an increased wake width for the large gap case. In all considered cases, large pressure fluctuations were observed in the front region of the blade suction side. These pressure fluctuations were strongest for the smaller spacing. The turbine efficiencies of the cases with the larger and smaller spacings were essentially the same, but both were lower than that of the normal case. The stator loss for the smaller spacing case was lower than the one for the larger spacing case, whereas the opposite was true for the rotor loss.

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.

Energy efficient engine: Low-pressure turbine subsonic cascade component development and integration program

NASA Technical Reports Server (NTRS)

Sharma, O. P.; Kopper, F. C.; Knudsen, L. K.; Yustinich, J. B.

1982-01-01

A subsonic cascade test program was conducted to provide technical data for optimizing the blade and vane airfoil designs for the Energy Efficient Engine Low-Pressure Turbine component. The program consisted of three parts. The first involved an evaluation of the low-chamber inlet guide vane. The second, was an evaluation of two candidate aerodynamic loading philosophies for the fourth blade root section. The third part consisted of an evaluation of three candidate airfoil geometries for the fourth blade mean section. The performance of each candidate airfoil was evaluated in a linear cascade configuration. The overall results of this study indicate that the aft-loaded airfoil designs resulted in lower losses which substantiated Pratt & Whitney Aircraft's design philosophy for the Energy Efficient Engine low-pressure turbine component.

Prediction of high-speed rotor noise with a Kirchhoff formula

NASA Technical Reports Server (NTRS)

Purcell, Timothy W.; Strawn, Roger C.; Yu, Yung H.

1987-01-01

A new methodology has been developed to predict the impulsive noise generated by a transonic rotor blade. The formulation uses a full-potential finite-difference method to obtain the pressure field close to the blade. A Kirchhoff integral formulation is then used to extend these finite-difference results into the far-field. This Kirchhoff formula is written in a blade-fixed coordinate system. It requires initial data across a plane at the sonic radius. This data is provided by the finite-difference solution. Acoustic pressure predictions show excellent agreement with hover experimental data for two hover cases of 0.88 and 0.90 tip Mach number, the latter of which has delocalized transonic flow. These results represent the first successful prediction technique for peak pressure amplitudes using a computational code.

Consistent modelling of wind turbine noise propagation from source to receiver.

PubMed

Barlas, Emre; Zhu, Wei Jun; Shen, Wen Zhong; Dag, Kaya O; Moriarty, Patrick

2017-11-01

The unsteady nature of wind turbine noise is a major reason for annoyance. The variation of far-field sound pressure levels is not only caused by the continuous change in wind turbine noise source levels but also by the unsteady flow field and the ground characteristics between the turbine and receiver. To take these phenomena into account, a consistent numerical technique that models the sound propagation from the source to receiver is developed. Large eddy simulation with an actuator line technique is employed for the flow modelling and the corresponding flow fields are used to simulate sound generation and propagation. The local blade relative velocity, angle of attack, and turbulence characteristics are input to the sound generation model. Time-dependent blade locations and the velocity between the noise source and receiver are considered within a quasi-3D propagation model. Long-range noise propagation of a 5 MW wind turbine is investigated. Sound pressure level time series evaluated at the source time are studied for varying wind speeds, surface roughness, and ground impedances within a 2000 m radius from the turbine.

Consistent modelling of wind turbine noise propagation from source to receiver

DOE PAGES

Barlas, Emre; Zhu, Wei Jun; Shen, Wen Zhong; ...

2017-11-28

The unsteady nature of wind turbine noise is a major reason for annoyance. The variation of far-field sound pressure levels is not only caused by the continuous change in wind turbine noise source levels but also by the unsteady flow field and the ground characteristics between the turbine and receiver. To take these phenomena into account, a consistent numerical technique that models the sound propagation from the source to receiver is developed. Large eddy simulation with an actuator line technique is employed for the flow modelling and the corresponding flow fields are used to simulate sound generation and propagation. Themore » local blade relative velocity, angle of attack, and turbulence characteristics are input to the sound generation model. Time-dependent blade locations and the velocity between the noise source and receiver are considered within a quasi-3D propagation model. Long-range noise propagation of a 5 MW wind turbine is investigated. Sound pressure level time series evaluated at the source time are studied for varying wind speeds, surface roughness, and ground impedances within a 2000 m radius from the turbine.« less

Consistent modelling of wind turbine noise propagation from source to receiver

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

Barlas, Emre; Zhu, Wei Jun; Shen, Wen Zhong

The unsteady nature of wind turbine noise is a major reason for annoyance. The variation of far-field sound pressure levels is not only caused by the continuous change in wind turbine noise source levels but also by the unsteady flow field and the ground characteristics between the turbine and receiver. To take these phenomena into account, a consistent numerical technique that models the sound propagation from the source to receiver is developed. Large eddy simulation with an actuator line technique is employed for the flow modelling and the corresponding flow fields are used to simulate sound generation and propagation. Themore » local blade relative velocity, angle of attack, and turbulence characteristics are input to the sound generation model. Time-dependent blade locations and the velocity between the noise source and receiver are considered within a quasi-3D propagation model. Long-range noise propagation of a 5 MW wind turbine is investigated. Sound pressure level time series evaluated at the source time are studied for varying wind speeds, surface roughness, and ground impedances within a 2000 m radius from the turbine.« less

Interior Noise Predictions in the Preliminary Design of the Large Civil Tiltrotor (LCTR2)

NASA Technical Reports Server (NTRS)

Grosveld, Ferdinand W.; Cabell, Randolph H.; Boyd, David D.

2013-01-01

A prediction scheme was established to compute sound pressure levels in the interior of a simplified cabin model of the second generation Large Civil Tiltrotor (LCTR2) during cruise conditions, while being excited by turbulent boundary layer flow over the fuselage, or by tiltrotor blade loading and thickness noise. Finite element models of the cabin structure, interior acoustic space, and acoustically absorbent (poro-elastic) materials in the fuselage were generated and combined into a coupled structural-acoustic model. Fluctuating power spectral densities were computed according to the Efimtsov turbulent boundary layer excitation model. Noise associated with the tiltrotor blades was predicted in the time domain as fluctuating surface pressures and converted to power spectral densities at the fuselage skin finite element nodes. A hybrid finite element (FE) approach was used to compute the low frequency acoustic cabin response over the frequency range 6-141 Hz with a 1 Hz bandwidth, and the Statistical Energy Analysis (SEA) approach was used to predict the interior noise for the 125-8000 Hz one-third octave bands.

Aero-thermal performances of leakage flows injection from the endwall slot in linear cascade of high-pressure turbine

NASA Astrophysics Data System (ADS)

Ghopa, Wan Aizon W.; Harun, Zambri; Funazaki, Ken-ichi; Miura, Takemitsu

2015-02-01

The existence of a gap between combustor and turbine endwall in the real gas turbine induces to the leakages phenomenon. However, the leakages could be used as a coolant to protect the endwall surfaces from the hot gas since it could not be completely prevented. Thus, present study investigated the potential of leakage flows as a function of film cooling. In present study, the flow field at the downstream of high-pressure turbine blade has been investigated by 5-holes pitot tube. This is to reveal the aerodynamic performances under the influenced of leakage flows while the temperature measurement was conducted by thermochromic liquid crystal (TLC). Experimental has significantly captured theaerodynamics effect of leakage flows near the blade downstream. Furthermore, TLC measurement illustrated that the film cooling effectiveness contours were strongly influenced by the secondary flows behavior on the endwall region. Aero-thermal results were validated by the numerical simulation adopted by commercial software, ANSYS CFX 13. Both experimental and numerical simulation indicated almost similar trendinaero and also thermal behavior as the amount of leakage flows increases.

Combined experiment Phase 2 data characterization

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

Miller, M.S.; Shipley, D.E.; Young, T.S.

1995-11-01

The National Renewable Energy Laboratory`s ``Combined Experiment`` has yielded a large quantity of experimental data on the operation of a downwind horizontal axis wind turbine under field conditions. To fully utilize this valuable resource and identify particular episodes of interest, a number of databases were created that characterize individual data events and rotational cycles over a wide range of parameters. Each of the 59 five-minute data episodes collected during Phase 11 of the Combined Experiment have been characterized by the mean, minimum, maximum, and standard deviation of all data channels, except the blade surface pressures. Inflow condition, aerodynamic force coefficient,more » and minimum leading edge pressure coefficient databases have also been established, characterizing each of nearly 21,000 blade rotational cycles. In addition, a number of tools have been developed for searching these databases for particular episodes of interest. Due to their extensive size, only a portion of the episode characterization databases are included in an appendix, and examples of the cycle characterization databases are given. The search tools are discussed and the FORTRAN or C code for each is included in appendices.« less

Solid Freeform Fabrication Symposium Proceedings Held in Austin, Texas on August 9-11, 1993

DTIC Science & Technology

1993-09-01

between the accuracy and the size of the geometric description. Highly non-linear surfaces, such as those that comprise turbine blades , manifolds...flange. The fan blades were modeled using different surfacing techniques. Seven blades are then combined with the rotor to make the completed fan. Figure...successfully cast in aluminum, titanium , beryllium-copper, and stainless steel, with RMS surface finish as low as 1 micrometer, without any subsequent

Small Laminated Axial Turbine Design and Test Program.

DTIC Science & Technology

1980-12-01

ILLUSTRATIONS Figure No. Title Page 1 Typical Test Results from TFE731 -3 Hot-Rig Testing. 5 2 Laminated Blade Chordwise Flow Patterns 8 3 Laminated Blade Cooling...Flow Parameter Versus Pressure Ratio 36 24 Blade Flow Distribution 37 25 TFE731 Turbofan Engine 38 26 Laminated Turbine Wheel 40 27 Selected Blade...facility, which was specifically developed to permit evaluation of cooled compo- nents for gas turbine engines. Four TFE731 -3 Laminated Turbine Wheels

Comprehensive Forced Response Analysis of J2X Turbine Bladed-Discs with 360 Degree Variation in CFD Loading

NASA Technical Reports Server (NTRS)

Elrod, David; Christensen, Eric; Brown, Andrew

2011-01-01

The temporal frequency content of the dynamic pressure predicted by a 360 degree computational fluid dynamics (CFD) analysis of a turbine flow field provides indicators of forcing function excitation frequencies (e.g., multiples of blade pass frequency) for turbine components. For the Pratt and Whitney Rocketdyne J-2X engine turbopumps, Campbell diagrams generated using these forcing function frequencies and the results of NASTRAN modal analyses show a number of components with modes in the engine operating range. As a consequence, forced response and static analyses are required for the prediction of combined stress, high cycle fatigue safety factors (HCFSF). Cyclically symmetric structural models have been used to analyze turbine vane and blade rows, not only in modal analyses, but also in forced response and static analyses. Due to the tortuous flow pattern in the turbine, dynamic pressure loading is not cyclically symmetric. Furthermore, CFD analyses predict dynamic pressure waves caused by adjacent and non-adjacent blade/vane rows upstream and downstream of the row analyzed. A MATLAB script has been written to calculate displacements due to the complex cyclically asymmetric dynamic pressure components predicted by CFD analysis, for all grids in a blade/vane row, at a chosen turbopump running speed. The MATLAB displacements are then read into NASTRAN, and dynamic stresses are calculated, including an adjustment for possible mistuning. In a cyclically symmetric NASTRAN static analysis, static stresses due to centrifugal, thermal, and pressure loading at the mode running speed are calculated. MATLAB is used to generate the HCFSF at each grid in the blade/vane row. When compared to an approach assuming cyclic symmetry in the dynamic flow field, the current approach provides better assurance that the worst case safety factor has been identified. An extended example for a J-2X turbopump component is provided.

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.

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.

Effect of wakes from moving upstream rods on boundary layer separation from a high lift airfoil

NASA Astrophysics Data System (ADS)

Volino, Ralph J.

2011-11-01

Highly loaded airfoils in turbines allow power generation using fewer airfoils. High loading, however, can cause boundary layer separation, resulting in reduced lift and increased aerodynamic loss. Separation is affected by the interaction between rotating blades and stationary vanes. Wakes from upstream vanes periodically impinge on downstream blades, and can reduce separation. The wakes include elevated turbulence, which can induce transition, and a velocity deficit, which results in an impinging flow on the blade surface known as a ``negative jet.'' In the present study, flow through a linear cascade of very high lift airfoils is studied experimentally. Wakes are produced with moving rods which cut through the flow upstream of the airfoils, simulating the effect of upstream vanes. Pressure and velocity fields are documented. Wake spacing and velocity are varied. At low Reynolds numbers without wakes, the boundary layer separates and does not reattach. At high wake passing frequencies separation is largely suppressed. At lower frequencies, ensemble averaged velocity results show intermittent separation and reattachment during the wake passing cycle. Supported by NASA.

Transition modelling implications in the CFD analysis of a turbine nozzle vane cascade tested over a range of Mach and Reynolds numbers

NASA Astrophysics Data System (ADS)

Marconcini, Michele; Pacciani, Roberto; Arnone, Andrea

2015-11-01

The aerodynamic performance of a gas turbine nozzle vane cascade was investigated over a range of Mach and Reynolds numbers. The work is part of a vast research project aimed at the analysis of fluid dynamics and heat transfer phenomena in cooled blades. In this paper computed results on the "solid vane" (without cooling devices) are presented and discussed in comparison with experimental data. Detailed measurements were provided by the University of Bergamo where the experimental campaign was carried out by means of a subsonic wind tunnel. The impact of boundary layer transition is investigated by using a novel laminar kinetic energy transport model and the widely used Langtry-Menter γ- Re θ,t model. The comparison between calculations and measurements is presented in terms of blade loading distributions, total pressure loss coefficient contours downstream of the cascade, and velocity/turbulence-intensity profiles within the boundary layer at selected blade surface locations at mid-span. It will be shown how transitional calculations compare favorably with experiments.

High speed propeller acoustics and aerodynamics - A boundary element approach

NASA Technical Reports Server (NTRS)

Farassat, F.; Myers, M. K.; Dunn, M. H.

1989-01-01

The Boundary Element Method (BEM) is applied in this paper to the problems of acoustics and aerodynamics of high speed propellers. The underlying theory is described based on the linearized Ffowcs Williams-Hawkings equation. The surface pressure on the blade is assumed unknown in the aerodynamic problem. It is obtained by solving a singular integral equation. The acoustic problem is then solved by moving the field point inside the fluid medium and evaluating some surface and line integrals. Thus the BEM provides a powerful technique in calculation of high speed propeller aerodynamics and acoustics.

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.

The change of the inlet geometry of a centrifugal compressor stage and its influence on the compressor performance

NASA Astrophysics Data System (ADS)

Wang, Leilei; Yang, Ce; Zhao, Ben; Lao, Dazhong; Ma, Chaochen; Li, Du

2013-06-01

The impact on the compressor performance is important for designing the inlet pipe of the centrifugal compressor of a vehicle turbocharger with different inlet pipes. First, an experiment was performed to determine the compressor performance from three cases: a straight inlet pipe, a long bent inlet pipe and a short bent inlet pipe. Next, dynamic sensors were installed in key positions to collect the sign of the unsteady pressure of the centrifugal compressor. Combined with the results of numerical simulations, the total pressure distortion in the pipes, the pressure distributions on the blades and the pressure variability in the diffuser are studied in detail. The results can be summarized as follows: a bent pipe results in an inlet distortion to the compressor, which leads to performance degradation, and the effect is more apparent as the mass flow rate increases. The distortion induced by the bent inlet is not only influenced by the distance between the outlet of the bent section and the leading edge of the impeller but also by the impeller rotation. The flow fields in the centrifugal impeller and the diffuser are influenced by a coupling effect produced by the upstream inlet distortion and the downstream blocking effect from the volute tongue. If the inlet geometry is changed, the distributions and the fluctuation intensities of the static pressure on the main blade surface of the centrifugal impeller and in the diffuser are changed accordingly.

Effects of nozzle-strut integrated design concepton on the subsonic turbine stage flowfield

NASA Astrophysics Data System (ADS)

Liu, Jun; Du, Qiang; Liu, Guang; Wang, Pei; Zhu, Junqiang

2014-10-01

In order to shorten aero-engine axial length, substituting the traditional long chord thick strut design accompanied with the traditional low pressure(LP) stage nozzle, LP turbine is integrated with intermediate turbine duct (ITD). In the current paper, five vanes of the first stage LP turbine nozzle is replaced with loaded struts for supporting the engine shaft, and providing oil pipes circumferentially which fulfilled the areo-engine structure requirement. However, their bulky geometric size represents a more effective obstacle to flow from high pressure (HP) turbine rotor. These five struts give obvious influence for not only the LP turbine nozzle but also the flowfield within the ITD, and hence cause higher loss. Numerical investigation has been undertaken to observe the influence of the Nozzle-Strut integrated design concept on the flowfield within the ITD and the nearby nozzle blades. According to the computational results, three main conclusions are finally obtained. Firstly, a noticeable low speed area is formed near the strut's leading edge, which is no doubt caused by the potential flow effects. Secondly, more severe radial migration of boundary layer flow adjacent to the strut's pressure side have been found near the nozzle's trailing edge. Such boundary layer migration is obvious, especially close to the shroud domain. Meanwhile, radial pressure gradient aggravates this phenomenon. Thirdly, velocity distribution along the strut's pressure side on nozzle's suction surface differs, which means loading variation of the nozzle. And it will no doubt cause nonuniform flowfield faced by the downstream rotor blade.

Aeroacoustic model of a modulation fan with pitching blades as a sound generator.

PubMed

Du, Lin; Jing, Xiaodong; Sun, Xiaofeng; Song, Weihua

2014-10-01

This paper is to develop an aeroacoustic model for a type of modulation fan termed as rotary subwoofer that is capable of radiating low-frequency sound at high sound pressure levels. The rotary subwoofer is modeled as a baffled monopole whose source strength is specified by the fluctuating mass flow rate produced by the pitching blades that rotate at constant speed. An immersed boundary method is established to simulate the detailed unsteady flow around the blades and also to estimate the source strength for the prediction of the far-field sound pressure level (SPL). The numerical simulation shows that the rotary subwoofer can output oscillating air flow that is in phase with the pitching motion of the blades. It is found that flow separation is more likely to occur on the pitching blades at higher modulation frequency, resulting in the reduction of the radiated SPL. Increasing the maximum blade excursion is one of the most effective means to enhance the sound radiation, but this effect can also be compromised by the flow separation. As the modulation frequency increases, correspondingly increasing the rotational speed or using larger blade solidity is beneficial to suppressing the flow separation and thus improving the acoustic performance of the rotary subwoofer.

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.

@NWTC Newsletter: Spring 2013 | Wind | NREL

Science.gov Websites

turbine blades. But according to a recent National Renewable Energy Laboratory (NREL) study that was appears unlikely that the pressure changes around operating wind turbine blades are large enough to cause

Application of response surface techniques to helicopter rotor blade optimization procedure

NASA Technical Reports Server (NTRS)

Henderson, Joseph Lynn; Walsh, Joanne L.; Young, Katherine C.

1995-01-01

In multidisciplinary optimization problems, response surface techniques can be used to replace the complex analyses that define the objective function and/or constraints with simple functions, typically polynomials. In this work a response surface is applied to the design optimization of a helicopter rotor blade. In previous work, this problem has been formulated with a multilevel approach. Here, the response surface takes advantage of this decomposition and is used to replace the lower level, a structural optimization of the blade. Problems that were encountered and important considerations in applying the response surface are discussed. Preliminary results are also presented that illustrate the benefits of using the response surface.

Thermal-barrier-coated turbine blade study

NASA Technical Reports Server (NTRS)

Siemers, P. A.; Hillig, W. B.

1981-01-01

The effects of coating TBC on a CF6-50 stage 2 high-pressure turbine blade were analyzed with respect to changes in the mean bulk temperature, cooling air requirements, and high-cycle fatigue. Localized spallation was found to have a possible deleterious effect on low-cycle fatigue life. New blade design concepts were developed to take optimum advantage of TBCs. Process and material development work and rig evaluations were undertaken which identified the most promising combination as ZrO2 containing 8 w/o Y2O3 applied by air plasma spray onto a Ni22Cr-10Al-1Y bond layer. The bond layer was applied by a low-pressure, high-velocity plasma spray process onto the base alloy. During the initial startup cycles the blades experienced localized leading edge spallation caused by foreign objects.

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.

Combined Experimental and Numerical Simulations of Thermal Barrier Coated Turbine Blades Erosion

NASA Technical Reports Server (NTRS)

Hamed, Awate; Tabakoff, Widen; Swar, Rohan; Shin, Dongyun; Woggon, Nthanial; Miller, Robert

2013-01-01

A combined experimental and computational study was conducted to investigate the erosion of thermal barrier coated (TBC) blade surfaces by alumina particles ingestion in a single stage turbine. In the experimental investigation, tests of particle surface interactions were performed in specially designed tunnels to determine the erosion rates and particle restitution characteristics under different impact conditions. The experimental results show that the erosion rates increase with increased impingement angle, impact velocity and temperature. In the computational simulations, an Euler-Lagrangian two stage approach is used in obtaining numerical solutions to the three-dimensional compressible Reynolds Averaged Navier-Stokes equations and the particles equations of motion in each blade passage reference frame. User defined functions (UDF) were developed to represent experimentally-based correlations for particle surface interaction models which were employed in the three-dimensional particle trajectory simulations to determine the particle rebound characteristics after each surface impact. The experimentally based erosion UDF model was used to predict the TBC erosion rates on the turbine blade surfaces based on the computed statistical data of the particles impact locations, velocities and angles relative to the blade surface. Computational results are presented for the predicted TBC blade erosion in a single stage commercial APU turbine, for a NASA designed automotive turbine, and for the NASA turbine scaled for modern rotorcraft operating conditions. The erosion patterns in the turbines are discussed for uniform particle ingestion and for particle ingestion concentrated in the inner and outer 5 percent of the stator blade span representing the flow cooling the combustor liner.

Nonlinear heat transfer and structural analyses of SSME turbine blades

NASA Technical Reports Server (NTRS)

Abdul-Aziz, A.; Kaufman, A.

1987-01-01

Three-dimensional nonlinear finite-element heat transfer and structural analyses were performed for the first stage high-pressure fuel turbopump blade of the space shuttle main engine (SSME). Directionally solidified (DS) MAR-M 246 material properties were considered for the analyses. Analytical conditions were based on a typical test stand engine cycle. Blade temperature and stress-strain histories were calculated using MARC finite-element computer code. The study was undertaken to assess the structural response of an SSME turbine blade and to gain greater understanding of blade damage mechanisms, convective cooling effects, and the thermal-mechanical effects.

Investigation of Hot Streak Migration and Film Cooling Effects on the Heat Transfer in Rotor/Stator Interacting Flows.

DTIC Science & Technology

1991-07-15

Study of Endwall and Airfoil Surface Heat Transfer in a Large Scale Turbine Blade Cascade," Journal of Engineering for Power, Vol. 102, No. 2, April...the turbine passage and along the surface of where d6 is the distance from the blade surface to a given node, the airfoil . In addition, a specified...effects on the passage flow and blade surface heat transfer for an axial flow turbine stage. These objectives are part of an overall plan to extend the

Effect of reducing rotor blade inlet diameter on the performance of a 11.66-Centimeter radial-inflow turbine

NASA Technical Reports Server (NTRS)

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

1973-01-01

The effect of increased rotor blade loading on turbine performance was investigated by reducing rotor blade inlet diameter. The reduction was made in four stages. Each modification was tested with the same stator using cold air as the working fluid. Results are presented in terms of equivalent mass flow and efficiency at equivalent design rotative speed and over a range of pressure ratios. Internal flow characteristics are shown in terms of stator exit static pressure and the radial variation of local loss and rotor-exit flow angle with radius ratio. Included are velocity diagrams calculated from the experimental results.

Calculation of gas turbine characteristic

NASA Astrophysics Data System (ADS)

Mamaev, B. I.; Murashko, V. L.

2016-04-01

The reasons and regularities of vapor flow and turbine parameter variation depending on the total pressure drop rate π* and rotor rotation frequency n are studied, as exemplified by a two-stage compressor turbine of a power-generating gas turbine installation. The turbine characteristic is calculated in a wide range of mode parameters using the method in which analytical dependences provide high accuracy for the calculated flow output angle and different types of gas dynamic losses are determined with account of the influence of blade row geometry, blade surface roughness, angles, compressibility, Reynolds number, and flow turbulence. The method provides satisfactory agreement of results of calculation and turbine testing. In the design mode, the operation conditions for the blade rows are favorable, the flow output velocities are close to the optimal ones, the angles of incidence are small, and the flow "choking" modes (with respect to consumption) in the rows are absent. High performance and a nearly axial flow behind the turbine are obtained. Reduction of the rotor rotation frequency and variation of the pressure drop change the flow parameters, the parameters of the stages and the turbine, as well as the form of the characteristic. In particular, for decreased n, nonmonotonic variation of the second stage reactivity with increasing π* is observed. It is demonstrated that the turbine characteristic is mainly determined by the influence of the angles of incidence and the velocity at the output of the rows on the losses and the flow output angle. The account of the growing flow output angle due to the positive angle of incidence for decreased rotation frequencies results in a considerable change of the characteristic: poorer performance, redistribution of the pressure drop at the stages, and change of reactivities, growth of the turbine capacity, and change of the angle and flow velocity behind the turbine.

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.

SSME single-crystal turbine blade dynamics

NASA Technical Reports Server (NTRS)

Moss, Larry A.

1988-01-01

A study was performrd to determine the dynamic characteristics of the Space Shuttle Main Engine high pressure fuel turbopump (HPFTP) blades made of single crystal (SC) material. The first and second stage drive turbine blades of HPFTP were examined. The nonrotating natural frequencies were determined experimentally and analytically. The experimental results of the SC second stage blade were used to verify the analytical procedures. The study examined the SC first stage blade natural frequencies with respect to crystal orientation at typical operating conditions. The SC blade dynamic response was predicted to be less than the directionally solidified base. Crystal axis orientation optimization indicated that the third mode interference will exist in any SC orientation.

Numerical Investigations of Active Flow Control for Low-Pressure Turbine Blades

DTIC Science & Technology

2008-03-01

points were clustered near the wall, in the separated flow region on the suction side of the 7 blade , and in the wake . Table 1 summarizes the block...Perspective view of blade (computational domain was repeated once in spanwise direction), side view of wake , and top down view of wake . Distributions...to the wake region. A second observation is that the wake turbulence appears to be concentrated in "lumps". In analogy to other wake flows, the blade

Experimental study of low aspect ratio compressor blading

NASA Technical Reports Server (NTRS)

Reid, L.; Moore, R. D.

1979-01-01

The effects of low aspect ratio blading on aerodynamic performance were examined. Four individual transonic compressor stages, representative of the inlet stage of an advanced high pressure ratio core compressor, are discussed. The flow phenomena for the four stages are investigated. Comparisons of blade element parameters are presented for the two different aspect ratio configurations. Blade loading levels are compared for the near stall conditions and comparisons are made of loss and diffusion factors over the operating range of incidence angles.

Structural changes and damage of single-crystal turbine blades during life tests of an aviation gas turbine engine

NASA Astrophysics Data System (ADS)

Ospennikova, O. G.; Orlov, M. R.; Kolodochkina, V. G.; Nazarkin, R. M.

2015-04-01

The irreversible structural changes of the single-crystal ZhS32-VI nickel superalloy blades of a high-pressure turbine that occur during life tests of a gas turbine engine are studied. The main operation damages in the hottest section of the blade airfoil are found to be the fracture of the heat-resistant coating in the leading edge and the formation of thermomechanical fatigue cracks. The possibility of reconditioning repair of the blades is considered.

Predictions for the Effects of Free Stream Turbulence on Turbine Blade Heat Transfer

NASA Technical Reports Server (NTRS)

Boyle, Robert J.; Giel, Paul W.; Ames, Forrest E.

2004-01-01

An approach to predicting the effects of free stream turbulence on turbine vane and blade heat transfer is described. Four models for predicting the effects of free stream turbulence were in incorporated into a Navier-Stokes CFD analysis. Predictions were compared with experimental data in order to identify an appropriate model for use across a wide range of flow conditions. The analyses were compared with data from five vane geometries and from four rotor geometries. Each of these nine geometries had data for different Reynolds numbers. Comparisons were made for twenty four cases. Steady state calculations were done because all experimental data were obtained in steady state tests. High turbulence levels often result in suction surface transition upstream of the throat, while at low to moderate Reynolds numbers the pressure surface remains laminar. A two-dimensional analysis was used because the flow is predominately two-dimensional in the regions where free stream turbulence significantly augments surface heat transfer. Because the evaluation of models for predicting turbulence effects can be affected by other factors, the paper discusses modeling for transition, relaminarization, and near wall damping. Quantitative comparisons are given between the predictions and data.

Aircraft turbofan noise

NASA Astrophysics Data System (ADS)

Groeneweg, J. F.; Rice, E. J.

1987-01-01

Turbofan noise generation and suppression in aircraft engines are reviewed. The chain of physical processes which connect unsteady flow interactions with fan blades to far field noise is addressed. Mechanism identification and description, duct propagation, radiation, and acoustic suppression are discussed. The experimental techniques of fan inflow static tests are discussed. Rotor blade surface pressure and wake velocity measurements aid in the determination of the types and strengths of the generation mechanisms. Approaches to predicting or measuring acoustic mode content, optimizing treatment impedance to maximize attenuation, translating impedance into porous wall structure, and interpreting far field directivity patterns are illustrated by comparisons of analytical and experimental results. The interdependence of source and acoustic treatment design to minimize far field noise is emphasized. Areas requiring further research are discussed, and the relevance of aircraft turbofan results to quieting other turbomachinery installation is addressed.

Aircraft turbofan noise

NASA Astrophysics Data System (ADS)

Groeneweg, J. F.; Rice, E. J.

1983-03-01

Turbofan noise generation and suppression in aircraft engines are reviewed. The chain of physical processes which connect unsteady flow interactions with fan blades to far field noise is addressed. Mechanism identification and description, duct propagation, radiation and acoustic suppression are discussed. The experimental technique of fan inflow static tests are discussed. Rotor blade surface pressure and wake velocity measurements aid in the determination of the types and strengths of the generation mechanisms. Approaches to predicting or measuring acoustic mode content, optimizing treatment impedance to maximize attenuation, translating impedance into porous wall structure and interpreting far field directivity patterns are illustrated by comparisons of analytical and experimental results. The interdependence of source and acoustic treatment design to minimize far field noise is emphasized. Area requiring further research are discussed and the relevance of aircraft turbofan results to quieting other turbomachinery installations is addressed.

Aircraft turbofan noise

NASA Technical Reports Server (NTRS)

Groeneweg, J. F.; Rice, E. J.

1983-01-01

Turbofan noise generation and suppression in aircraft engines are reviewed. The chain of physical processes which connect unsteady flow interactions with fan blades to far field noise is addressed. Mechanism identification and description, duct propagation, radiation and acoustic suppression are discussed. The experimental technique of fan inflow static tests are discussed. Rotor blade surface pressure and wake velocity measurements aid in the determination of the types and strengths of the generation mechanisms. Approaches to predicting or measuring acoustic mode content, optimizing treatment impedance to maximize attenuation, translating impedance into porous wall structure and interpreting far field directivity patterns are illustrated by comparisons of analytical and experimental results. The interdependence of source and acoustic treatment design to minimize far field noise is emphasized. Area requiring further research are discussed and the relevance of aircraft turbofan results to quieting other turbomachinery installations is addressed.

Advances in measuring techniques for turbine cooling test rigs

NASA Technical Reports Server (NTRS)

Pollack, F. G.

1972-01-01

Surface temperature distribution measurements for turbine vanes and blades were obtained by measuring the infrared energy emitted by the airfoil. The IR distribution can be related to temperature distribution by suitable calibration methods and the data presented in the form of isotherm maps. Both IR photographic and real time electro-optical methods are being investigated. The methods can be adapted to rotating as well as stationary targets, and both methods can utilize computer processing. Pressure measurements on rotating components are made with a rotating system incorporating 10 miniature transducers. A mercury wetted slip ring assembly was used to supply excitation power and as a signal transfer device. The system was successfully tested up to speeds of 9000 rpm and is now being adapted to measure rotating blade airflow quantities in a spin rig and a research engine.

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.

Hot spot detection system for vanes or blades of a combustion turbine

DOEpatents

Twerdochlib, Michael

1999-01-01

This invention includes a detection system that can determine if a turbine component, such as a turbine vane or blade, has exceeded a critical temperature, such as a melting point, along any point along the entire surface of the vane or blade. This system can be employed in a conventional combustion turbine having a compressor, a combustor and a turbine section. Included within this system is a chemical coating disposed along the entire interior surface of a vane or blade and a closed loop cooling system that circulates a coolant through the interior of the vane or blade. If the temperature of the vane or blade exceeds a critical temperature, the chemical coating will be expelled from the vane or blade into the coolant. Since while traversing the closed loop cooling system the coolant passes through a detector, the presence of the chemical coating in the coolant will be sensed by the system. If the chemical coating is detected, this indicates that the vane or blade has exceeded a critical temperature.

Investigation on centrifugal impeller in an axial-radial combined compressor with inlet distortion

NASA Astrophysics Data System (ADS)

Li, Du; Yang, Ce; Zhao, Ben; Zhou, Mi; Qi, Mingxu; Zhang, Jizhong

2011-12-01

Assembling an axial rotor and a stator at centrifugal compressor upstream to build an axial-radial combined compressor could achieve high pressure ratio and efficiency by appropriate size augment. Then upstream potential flow and wake effect appear at centrifugal impeller inlet. In this paper, the axial-radial compressor is unsteadily simulated by three-dimensional Reynolds averaged Navier-Stokes equations with uniform and circumferential distorted total pressure inlet condition to investigate upstream effect on radial rotor. The results show that span-wise nonuniform total pressure distribution is generated and radial and circumferential combined distortion is formed at centrifugal rotor inlet. The upstream stator wake deflects to rotor rotation direction and decreases with blade span increases. Circumferential distortion causes different separated flow formations at different pitch positions. The tip leakage vortex is suppressed in centrifugal blade passages. Under distorted inlet condition, flow direction of centrifugal impeller leading edge upstream varies evidently near hub and shroud but varies slightly at mid-span. In addition, compressor stage inlet distortion produces remarkable effect on blade loading of centrifugal blade both along chordwise and pitchwise.

Articulated limiter blade for a tokamak fusion reactor

DOEpatents

Doll, D.W.

1982-10-21

A limiter blade for a large tokomak fusion reactor includes three articulated blade sections for enabling the limiter blade to be adjusted for plasmas of different sizes. Each blade section is formed of a rigid backing plate carrying graphite tiles coated with titanium carbide, and the limiter blade forms a generally elliptic contour in both the poloidal and toroidal directions to uniformly distribute the heat flow to the blade. The limiter blade includes a central blade section movable along the major radius of the vacuum vessel, and upper and lower pivotal blade sections which may be pivoted by linear actuators having rollers held to the back surface of the pivotal blade sections.