Unducted, counterrotating gearless front fan engine

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

Taylor, J.B.

This patent describes a high bypass ratio gas turbine engine. It comprises a core engine effective for generating combustion gases passing through a main flow path; a power turbine aft of the core engine and including first and second counter rotatable interdigitated turbine blade rows, effective for counterrotating first and second drive shafts, respectively; an unducted fan section forward of the core engine including a first fan blade row connected to the first drive shaft and a second fan blade row axially spaced aftward from the first fan blade row and connected to the second drive shaft; and a boostermore » compressor axially positioned between the first and second fan blade rows and including first compressor blade rows connected to the first drive shaft and second compressor blade rows connected to the second drive shaft.« less

A Numerical Analysis on the Effects of Self-Excited Tip Flow Unsteadiness and Upstream Blade Row Interactions on the Performance Predictions of a Transonic Compressor

NASA Astrophysics Data System (ADS)

Heberling, Brian

Computational fluid dynamics (CFD) simulations can offer a detailed view of the complex flow fields within an axial compressor and greatly aid the design process. However, the desire for quick turnaround times raises the question of how exact the model must be. At design conditions, steady CFD simulating an isolated blade row can accurately predict the performance of a rotor. However, as a compressor is throttled and mass flow rate decreased, axial flow becomes weaker making the capturing of unsteadiness, wakes, or other flow features more important. The unsteadiness of the tip clearance flow and upstream blade wake can have a significant impact on a rotor. At off-design conditions, time-accurate simulations or modeling multiple blade rows can become necessary in order to receive accurate performance predictions. Unsteady and multi- bladerow simulations are computationally expensive, especially when used in conjunction. It is important to understand which features are important to model in order to accurately capture a compressor's performance. CFD simulations of a transonic axial compressor throttling from the design point to stall are presented. The importance of capturing the unsteadiness of the rotor tip clearance flow versus capturing upstream blade-row interactions is examined through steady and unsteady, single- and multi-bladerow computations. It is shown that there are significant differences at near stall conditions between the different types of simulations.

User's Guide for MSAP2D: A Program for Unsteady Aerodynamic and Aeroelastic (Flutter and Forced Response) Analysis of Multistage Compressors and Turbines. 1.0

NASA Technical Reports Server (NTRS)

Reddy, T. S. R.; Srivastava, R.

1996-01-01

This guide describes the input data required for using MSAP2D (Multi Stage Aeroelastic analysis Program - Two Dimensional) computer code. MSAP2D can be used for steady, unsteady aerodynamic, and aeroelastic (flutter and forced response) analysis of bladed disks arranged in multiple blade rows such as those found in compressors, turbines, counter rotating propellers or propfans. The code can also be run for single blade row. MSAP2D code is an extension of the original NPHASE code for multiblade row aerodynamic and aeroelastic analysis. Euler equations are used to obtain aerodynamic forces. The structural dynamic equations are written for a rigid typical section undergoing pitching (torsion) and plunging (bending) motion. The aeroelastic equations are solved in time domain. For single blade row analysis, frequency domain analysis is also provided to obtain unsteady aerodynamic coefficients required in an eigen analysis for flutter. In this manual, sample input and output are provided for a single blade row example, two blade row example with equal and unequal number of blades in the blade rows.

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.

Computation of flow in radial- and mixed-flow cascades by an inviscid-viscous interaction method

NASA Technical Reports Server (NTRS)

Serovy, G. K.; Hansen, E. C.

1980-01-01

The use of inviscid-viscous interaction methods for the case of radial or mixed-flow cascade diffusers is discussed. A literature review of investigations considering cascade flow-field prediction by inviscid-viscous iterative computation is given. Cascade aerodynamics in the third blade row of a multiple-row radial cascade diffuser are specifically investigated.

Coupled 2-dimensional cascade theory for noise and unsteady aerodynamics of blade row interaction in turbofans. Volume 1: Theory development and parametric studies

NASA Technical Reports Server (NTRS)

Hanson, Donald B.

1994-01-01

Typical analytical models for interaction between rotor and stator in a turbofan analyze the effect of wakes from the rotor impinging on the stator, producing unsteady loading, and thereby generating noise. Reflection/transmission characteristics of the rotor are sometimes added in a separate calculation. In those models, there is a one-to-one relationship between wake harmonics and noise harmonics; that is, the BPF (blade passing frequency) wake harmonic causes only the BPF noise harmonic, etc. This report presents a more complete model in which flow tangency boundary conditions are satisfied on two cascades in relative motion for several harmonics simultaneously. By an extension of S.N. Smith's code for two dimensional flat plate cascades, the noise generation/frequency scattering/blade row reflection problem is solved in a single matrix inversion. It is found that the BPF harmonic excitation of the stator scatters considerable energy in the higher BPF harmonics due to relative motion between the blade rows. Furthermore, when swirl between the rotor and stator is modeled, a 'mode trapping' effect occurs which explains observations on fans operating at rotational speeds below BFP cuton: the BPF mode amplifies between blade rows by multiple reflections but cannot escape to the inlet and exit ducts. However, energy scattered into higher harmonics does propagate and dominates the spectrum at two and three times BPF. This report presents the complete derivation of the theory, comparison with a previous (more limited) coupled rotor/stator interaction theory due to Kaji and Okazaki, exploration of the mode trapping phenomenon, and parametric studies showing the effects of vane/blade ratio and rotor/stator interaction. For generality, the analysis applies to stages where the rotor is either upstream or downstream of the stator and to counter rotation stages. The theory has been coded in a FORTRAN program called CUP2D, documented in Volume 2 of this report. It is concluded that the new features of this analysis - unsteady coupling, frequency scattering, and flow turning between rotor and stator - have a profound effect on noise generation caused by rotor/stator interaction. Treating rotors and stators as isolated cascades is not adequate for noise analysis and prediction.

Core Engine Noise Program. Volume III. Prediction Methods -- Supplement I. - Extension of Prediction Methods

DTIC Science & Technology

1976-03-01

frequency noise transmission through turbine blade rows and addition of engine and component data to the prediction method for core noise. " Phase VI...lower turbine blade row attenuation for this low bypass engine . When the blade row attenuation is accounted for by means of a turbine work extrac...component and engine data. Currently, an in-depth program to investigate turbine blade row attenuation is underway (NAS3-19435 and DOT-FA75WA-3688). The

Deterministic blade row interactions in a centrifugal compressor stage

NASA Technical Reports Server (NTRS)

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

1991-01-01

The three-dimensional viscous flow in a low speed centrifugal compressor stage is simulated using an average passage Navier-Stokes analysis. The impeller discharge flow is of the jet/wake type with low momentum fluid in the shroud-pressure side corner coincident with the tip leakage vortex. This nonuniformity introduces periodic unsteadiness in the vane frame of reference. The effect of such deterministic unsteadiness on the time-mean is included in the analysis through the average passage stress, which allows the analysis of blade row interactions. The magnitude of the divergence of the deterministic unsteady stress is of the order of the divergence of the Reynolds stress over most of the span, from the impeller trailing edge to the vane throat. Although the potential effects on the blade trailing edge from the diffuser vane are small, strong secondary flows generated by the impeller degrade the performance of the diffuser vanes.

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

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.

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.

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

NASA Technical Reports Server (NTRS)

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

1991-01-01

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

Coupled-Flow Simulation of HP-LP Turbines Has Resulted in Significant Fuel Savings

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2001-01-01

Our objective was to create a high-fidelity Navier-Stokes computer simulation of the flow through the turbines of a modern high-bypass-ratio turbofan engine. The simulation would have to capture the aerodynamic interactions between closely coupled high- and low-pressure turbines. A computer simulation of the flow in the GE90 turbofan engine's high-pressure (HP) and low-pressure (LP) turbines was created at GE Aircraft Engines under contract with the NASA Glenn Research Center. The three-dimensional steady-state computer simulation was performed using Glenn's average-passage approach named APNASA. The areas upstream and downstream of each blade row mutually interact with each other during engine operation. The embedded blade row operating conditions are modeled since the average passage equations in APNASA actively include the effects of the adjacent blade rows. The turbine airfoils, platforms, and casing are actively cooled by compressor bleed air. Hot gas leaks around the tips of rotors through labyrinth seals. The flow exiting the high work HP turbines is partially transonic and, therefore, has a strong shock system in the transition region. The simulation was done using 121 processors of a Silicon Graphics Origin 2000 (NAS 02K) cluster at the NASA Ames Research Center, with a parallel efficiency of 87 percent in 15 hr. The typical average-passage analysis mesh size per blade row was 280 by 45 by 55, or approx.700,000 grid points. The total number of blade rows was 18 for a combined HP and LP turbine system including the struts in the transition duct and exit guide vane, which contain 12.6 million grid points. Design cycle turnaround time requirements ran typically from 24 to 48 hr of wall clock time. The number of iterations for convergence was 10,000 at 8.03x10(exp -5) sec/iteration/grid point (NAS O2K). Parallel processing by up to 40 processors is required to meet the design cycle time constraints. This is the first-ever flow simulation of an HP and LP turbine. In addition, it includes the struts in the transition duct and exit guide vanes.

Numerical Investigations of Two Typical Unsteady Flows in Turbomachinery Using the Multi-Passage Model

NASA Astrophysics Data System (ADS)

Zhou, Di; Lu, Zhiliang; Guo, Tongqing; Shen, Ennan

2016-06-01

In this paper, the research on two types of unsteady flow problems in turbomachinery including blade flutter and rotor-stator interaction is made by means of numerical simulation. For the former, the energy method is often used to predict the aeroelastic stability by calculating the aerodynamic work per vibration cycle. The inter-blade phase angle (IBPA) is an important parameter in computation and may have significant effects on aeroelastic behavior. For the latter, the numbers of blades in each row are usually not equal and the unsteady rotor-stator interactions could be strong. An effective way to perform multi-row calculations is the domain scaling method (DSM). These two cases share a common point that the computational domain has to be extended to multi passages (MP) considering their respective features. The present work is aimed at modeling these two issues with the developed MP model. Computational fluid dynamics (CFD) technique is applied to resolve the unsteady Reynolds-averaged Navier-Stokes (RANS) equations and simulate the flow fields. With the parallel technique, the additional time cost due to modeling more passages can be largely decreased. Results are presented on two test cases including a vibrating rotor blade and a turbine stage.

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

NASA Technical Reports Server (NTRS)

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

1950-01-01

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

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.

Probabilistic Aeroelastic Analysis of Turbomachinery Components

NASA Technical Reports Server (NTRS)

Reddy, T. S. R.; Mital, S. K.; Stefko, G. L.

2004-01-01

A probabilistic approach is described for aeroelastic analysis of turbomachinery blade rows. Blade rows with subsonic flow and blade rows with supersonic flow with subsonic leading edge are considered. To demonstrate the probabilistic approach, the flutter frequency, damping and forced response of a blade row representing a compressor geometry is considered. The analysis accounts for uncertainties in structural and aerodynamic design variables. The results are presented in the form of probabilistic density function (PDF) and sensitivity factors. For subsonic flow cascade, comparisons are also made with different probabilistic distributions, probabilistic methods, and Monte-Carlo simulation. The approach shows that the probabilistic approach provides a more realistic and systematic way to assess the effect of uncertainties in design variables on the aeroelastic instabilities and response.

Study of Low Reynolds Number Effects on the Losses in Low-Pressure Turbine Blade Rows

NASA Technical Reports Server (NTRS)

Ashpis, David E.; Dorney, Daniel J.

1998-01-01

Experimental data from jet-engine tests have indicated that unsteady blade row interactions and separation can have a significant impact on the efficiency of low-pressure turbine stages. Measured turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Several recent studies have revealed that Reynolds number effects may contribute to the lower efficiencies at cruise conditions. In the current study numerical experiments have been performed to study the models available for low Reynolds number flows, and to quantify the Reynolds number dependence of low-pressure turbine cascades and stages. The predicted aerodynamic results exhibit good agreement with design data.

Estimation of the energy loss at the blades in rowing: common assumptions revisited.

PubMed

Hofmijster, Mathijs; De Koning, Jos; Van Soest, A J

2010-08-01

In rowing, power is inevitably lost as kinetic energy is imparted to the water during push-off with the blades. Power loss is estimated from reconstructed blade kinetics and kinematics. Traditionally, it is assumed that the oar is completely rigid and that force acts strictly perpendicular to the blade. The aim of the present study was to evaluate how reconstructed blade kinematics, kinetics, and average power loss are affected by these assumptions. A calibration experiment with instrumented oars and oarlocks was performed to establish relations between measured signals and oar deformation and blade force. Next, an on-water experiment was performed with a single female world-class rower rowing at constant racing pace in an instrumented scull. Blade kinematics, kinetics, and power loss under different assumptions (rigid versus deformable oars; absence or presence of a blade force component parallel to the oar) were reconstructed. Estimated power losses at the blades are 18% higher when parallel blade force is incorporated. Incorporating oar deformation affects reconstructed blade kinematics and instantaneous power loss, but has no effect on estimation of power losses at the blades. Assumptions on oar deformation and blade force direction have implications for the reconstructed blade kinetics and kinematics. Neglecting parallel blade forces leads to a substantial underestimation of power losses at the blades.

Simulation of 3-D viscous compressible flow in multistage turbomachinery by finite element methods

NASA Astrophysics Data System (ADS)

Sleiman, Mohamad

1999-11-01

The flow in a multistage turbomachinery blade row is compressible, viscous, and unsteady. Complex flow features such as boundary layers, wake migration from upstream blade rows, shocks, tip leakage jets, and vortices interact together as the flow convects through the stages. These interactions contribute significantly to the aerodynamic losses of the system and degrade the performance of the machine. The unsteadiness also leads to blade vibration and a shortening of its life. It is therefore difficult to optimize the design of a blade row, whether aerodynamically or structurally, in isolation, without accounting for the effects of the upstream and downstream rows. The effects of axial spacing, blade count, clocking (relative position of follow-up rotors with respect to wakes shed by upstream ones), and levels of unsteadiness may have a significance on performance and durability. In this Thesis, finite element formulations for the simulation of multistage turbomachinery are presented in terms of the Reynolds-averaged Navier-Stokes equations for three-dimensional steady or unsteady, viscous, compressible, turbulent flows. Three methodologies are presented and compared. First, a steady multistage analysis using a a-mixing- plane model has been implemented and has been validated against engine data. For axial machines, it has been found that the mixing plane simulation methods match very well the experimental data. However, the results for a centrifugal stage, consisting of an impeller followed by a vane diffuser of equal pitch, show flagrant inconsistency with engine performance data, indicating that the mixing plane method has been found to be inappropriate for centrifugal machines. Following these findings, a more complete unsteady multistage model has been devised for a configuration with equal number of rotor and stator blades (equal pitches). Non-matching grids are used at the rotor-stator interface and an implicit interpolation procedure devised to ensure continuity of fluxes across. This permits the rotor and stator equations to be solved in a fully- coupled manner, allowing larger time steps in attaining a time-periodic solution. This equal pitch approach has been validated on the complex geometry of a centrifugal stage. Finally, for a stage configuration with unequal pitches, the time-inclined method, developed by Giles (1991) for 2-D viscous compressible flow, has been extended to 3-D and formulated in terms of the physical solution vector U, rather than Q, a non-physical one. The method has been evaluated for unsteady flow through a rotor blade passage of the power turbine of a turboprop.

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.

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.

An analytical study of the effect of coolant flow variables on the kinetic energy output of a cooled turbine blade row.

NASA Technical Reports Server (NTRS)

Prust, H. W., Jr.

1972-01-01

Demonstration that the change in output of a cooled turbine blade row relative to the specific output of the uncooled blade row can be positive, negative, or zero, depending on the velocity, injection location, injection angle, and temperature of the coolant. Comparisons between the analytical results and experimental results for four different cases of coolant discharge, all at a coolant temperature ratio of unity, show good agreement for three cases, and rather poor agreement for the other.

Casing for a gas turbine engine

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

Wiebe, David J.; Little, David A.; Charron, Richard C.

2016-07-12

A casing for a can annular gas turbine engine, including: a compressed air section (40) spanning between a last row of compressor blades (26) and a first row of turbine blades (28), the compressed air section (40) having a plurality of openings (50) there through, wherein a single combustor/advanced duct assembly (64) extends through each opening (50); and one top hat (68) associated with each opening (50) configured to enclose the associated combustor/advanced duct assembly (64) and seal the opening (50). A volume enclosed by the compressed air section (40) is not greater than a volume of a frustum (54)more » defined at an upstream end (56) by an inner diameter of the casing at the last row of compressor blades (26) and at a downstream end (60) by an inner diameter of the casing at the first row of turbine blades (28).« less

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.

Analysis of hydrodynamic force acting on commercialized rowing blades using computational fluid dynamics

NASA Astrophysics Data System (ADS)

Aziz, A. M. Y.; Harun, M. N.; Syahrom, Ardiyansyah; Omar, A. H.

2017-04-01

This paper presents a study of the hydrodynamics of several rowing blade designs. The study was done using Computational Fluid Dynamics (CFD) which enabled the investigation to be done similar to the experimental study, but with additional hydrodynamic visualization for further analysis and understanding. The CFD method was validated using quasi-static experimental data from Caplan (2007). Besides that, the proposed CFD analyses have improved the precious CFD results with the percentage of error of 6.58 percent of lift and 0.69 percent of drag force compared to 33.65 and 18.75 percent obtained by Coppel (2010). Consequent to the successful validation, the study then proceeded with the real size of Macon, Big balde and Fat blade. It was found that the hydrodynamic performance of the Fat blade was the highest due to the area, aspect ratio and the shape of the blade. Besides that, distribution of pressure for all models were also investigated which deepened the understanding of the blade fluid mechanics of rowing.

AERODYNAMIC AND BLADING DESIGN OF MULTISTAGE AXIAL FLOW COMPRESSORS

NASA Technical Reports Server (NTRS)

Crouse, J. E.

1994-01-01

The axial-flow compressor is used for aircraft engines because it has distinct configuration and performance advantages over other compressor types. However, good potential performance is not easily obtained. The designer must be able to model the actual flows well enough to adequately predict aerodynamic performance. This computer program has been developed for computing the aerodynamic design of a multistage axial-flow compressor and, if desired, the associated blading geometry input for internal flow analysis. The aerodynamic solution gives velocity diagrams on selected streamlines of revolution at the blade row edges. The program yields aerodynamic and blading design results that can be directly used by flow and mechanical analysis codes. Two such codes are TSONIC, a blade-to-blade channel flow analysis code (COSMIC program LEW-10977), and MERIDL, a more detailed hub-to-shroud flow analysis code (COSMIC program LEW-12966). The aerodynamic and blading design program can reduce the time and effort required to obtain acceptable multistage axial-flow compressor configurations by generating good initial solutions and by being compatible with available analysis codes. The aerodynamic solution assumes steady, axisymmetric flow so that the problem is reduced to solving the two-dimensional flow field in the meridional plane. The streamline curvature method is used for the iterative aerodynamic solution at stations outside of the blade rows. If a blade design is desired, the blade elements are defined and stacked within the aerodynamic solution iteration. The blade element inlet and outlet angles are established by empirical incidence and deviation angles to the relative flow angles of the velocity diagrams. The blade element centerline is composed of two segments tangentially joined at a transition point. The local blade angle variation of each element can be specified as a fourth-degree polynomial function of path distance. Blade element thickness can also be specified with fourth-degree polynomial functions of path distance from the maximum thickness point. Input to the aerodynamic and blading design program includes the annulus profile, the overall compressor mass flow, the pressure ratio, and the rotative speed. A number of input parameters are also used to specify and control the blade row aerodynamics and geometry. The output from the aerodynamic solution has an overall blade row and compressor performance summary followed by blade element parameters for the individual blade rows. If desired, the blade coordinates in the streamwise direction for internal flow analysis codes and the coordinates on plane sections through blades for fabrication drawings may be stored and printed. The aerodynamic and blading design program for multistage axial-flow compressors is written in FORTRAN IV for batch execution and has been implemented on an IBM 360 series computer with a central memory requirement of approximately 470K of 8 bit bytes. This program was developed in 1981.

Fundamental mechanisms that influence the estimate of heat transfer to gas turbine blades

NASA Technical Reports Server (NTRS)

Graham, R. W.

1979-01-01

Estimates of the heat transfer from the gas to stationary (vanes) or rotating blades poses a major uncertainty due to the complexity of the heat transfer processes. The gas flow through these blade rows is three dimensional with complex secondary viscous flow patterns that interact with the endwalls and blade surfaces. In addition, upstream disturbances, stagnation flow, curvature effects, and flow acceleration complicate the thermal transport mechanisms in the boundary layers. Some of these fundamental heat transfer effects are discussed. The chief purpose of the discussion is to acquaint those in the heat transfer community, not directly involved in gas turbines, of the seriousness of the problem and to recommend some basic research that would improve the capability for predicting gas-side heat transfer on turbine blades and vanes.

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.

Task III: Development of an Effective Computational Methodology for Body Force Representation of High-speed Rotor 37

NASA Technical Reports Server (NTRS)

Tan, Choon-Sooi; Suder, Kenneth (Technical Monitor)

2003-01-01

A framework for an effective computational methodology for characterizing the stability and the impact of distortion in high-speed multi-stage compressor is being developed. The methodology consists of using a few isolated-blade row Navier-Stokes solutions for each blade row to construct a body force database. The purpose of the body force database is to replace each blade row in a multi-stage compressor by a body force distribution to produce same pressure rise and flow turning. To do this, each body force database is generated in such a way that it can respond to the changes in local flow conditions. Once the database is generated, no hrther Navier-Stokes computations are necessary. The process is repeated for every blade row in the multi-stage compressor. The body forces are then embedded as source terms in an Euler solver. The method is developed to have the capability to compute the performance in a flow that has radial as well as circumferential non-uniformity with a length scale larger than a blade pitch; thus it can potentially be used to characterize the stability of a compressor under design. It is these two latter features as well as the accompanying procedure to obtain the body force representation that distinguish the present methodology from the streamline curvature method. The overall computational procedures have been developed. A dimensional analysis was carried out to determine the local flow conditions for parameterizing the magnitudes of the local body force representation of blade rows. An Euler solver was modified to embed the body forces as source terms. The results from the dimensional analysis show that the body forces can be parameterized in terms of the two relative flow angles, the relative Mach number, and the Reynolds number. For flow in a high-speed transonic blade row, they can be parameterized in terms of the local relative Mach number alone.

An analytical study of the effect of coolant flow variables on the kinetic energy output of a cooled turbine blade flow

NASA Technical Reports Server (NTRS)

Prust, H. W., Jr.

1971-01-01

The results of an analytical study to determine the effect of changes in the amount, velocity, injection location, injection angle, and temperature of coolant flow on blade row performance are presented. The results show that the change in output of a cooled turbine blade row relative to the specific output of the uncooled blade row can be positive, negative, or zero. Comparisons between the analytical results and experimental results for four different cases of coolant discharge, all at a coolant temperature ratio of unity, show good agreement for three cases and rather poor agreement for the other. To further test the validity of the method, more experimental data is needed, particularly at different coolant temperature ratios.

An Evaluation of a Phase-Lag Boundary Condition for Francis Hydroturbine Simulations Using a Pressure-Based Solver

NASA Astrophysics Data System (ADS)

Wouden, Alex; Cimbala, John; Lewis, Bryan

2014-11-01

While the periodic boundary condition is useful for handling rotational symmetry in many axisymmetric geometries, its application fails for analysis of rotor-stator interaction (RSI) in multi-stage turbomachinery flow. The inadequacy arises from the underlying geometry where the blade counts per row differ, since the blade counts are crafted to deter the destructive harmonic forces of synchronous blade passing. Therefore, to achieve the computational advantage of modeling a single blade passage per row while preserving the integrity of the RSI, a phase-lag boundary condition is adapted to OpenFOAM® software's incompressible pressure-based solver. The phase-lag construct is accomplished through restating the implicit periodic boundary condition as a constant boundary condition that is updated at each time step with phase-shifted data from the coupled cells adjacent to the boundary. Its effectiveness is demonstrated using a typical Francis hydroturbine modeled as single- and double-passages with phase-lag boundary conditions. The evaluation of the phase-lag condition is based on the correspondence of the overall computational performance and the calculated flow parameters of the phase-lag simulations with those of a baseline full-wheel simulation. Funded in part by DOE Award Number: DE-EE0002667.

Parallel 3D Multi-Stage Simulation of a Turbofan Engine

NASA Technical Reports Server (NTRS)

Turner, Mark G.; Topp, David A.

1998-01-01

A 3D multistage simulation of each component of a modern GE Turbofan engine has been made. An axisymmetric view of this engine is presented in the document. This includes a fan, booster rig, high pressure compressor rig, high pressure turbine rig and a low pressure turbine rig. In the near future, all components will be run in a single calculation for a solution of 49 blade rows. The simulation exploits the use of parallel computations by using two levels of parallelism. Each blade row is run in parallel and each blade row grid is decomposed into several domains and run in parallel. 20 processors are used for the 4 blade row analysis. The average passage approach developed by John Adamczyk at NASA Lewis Research Center has been further developed and parallelized. This is APNASA Version A. It is a Navier-Stokes solver using a 4-stage explicit Runge-Kutta time marching scheme with variable time steps and residual smoothing for convergence acceleration. It has an implicit K-E turbulence model which uses an ADI solver to factor the matrix. Between 50 and 100 explicit time steps are solved before a blade row body force is calculated and exchanged with the other blade rows. This outer iteration has been coined a "flip." Efforts have been made to make the solver linearly scaleable with the number of blade rows. Enough flips are run (between 50 and 200) so the solution in the entire machine is not changing. The K-E equations are generally solved every other explicit time step. One of the key requirements in the development of the parallel code was to make the parallel solution exactly (bit for bit) match the serial solution. This has helped isolate many small parallel bugs and guarantee the parallelization was done correctly. The domain decomposition is done only in the axial direction since the number of points axially is much larger than the other two directions. This code uses MPI for message passing. The parallel speed up of the solver portion (no 1/0 or body force calculation) for a grid which has 227 points axially.

Investigation of advanced counterrotation blade configuration concepts for high speed turboprop systems. Task 4: Advanced fan section aerodynamic analysis computer program user's manual

NASA Technical Reports Server (NTRS)

Crook, Andrew J.; Delaney, Robert A.

1992-01-01

The computer program user's manual for the ADPACAPES (Advanced Ducted Propfan Analysis Code-Average Passage Engine Simulation) program is included. The objective of the computer program is development of a three-dimensional Euler/Navier-Stokes flow analysis for fan section/engine geometries containing multiple blade rows and multiple spanwise flow splitters. An existing procedure developed by Dr. J. J. Adamczyk and associates at the NASA Lewis Research Center was modified to accept multiple spanwise splitter geometries and simulate engine core conditions. The numerical solution is based upon a finite volume technique with a four stage Runge-Kutta time marching procedure. Multiple blade row solutions are based upon the average-passage system of equations. The numerical solutions are performed on an H-type grid system, with meshes meeting the requirement of maintaining a common axisymmetric mesh for each blade row grid. The analysis was run on several geometry configurations ranging from one to five blade rows and from one to four radial flow splitters. The efficiency of the solution procedure was shown to be the same as the original analysis.

Gas turbine sealing apparatus

DOEpatents

Wiebe, David J; Wessell, Brian J; Ebert, Todd; Beeck, Alexander; Liang, George; Marussich, Walter H

2013-02-19

A gas turbine includes forward and aft rows of rotatable blades, a row of stationary vanes between the forward and aft rows of rotatable blades, an annular intermediate disc, and a seal housing apparatus. The forward and aft rows of rotatable blades are coupled to respective first and second portions of a disc/rotor assembly. The annular intermediate disc is coupled to the disc/rotor assembly so as to be rotatable with the disc/rotor assembly during operation of the gas turbine. The annular intermediate disc includes a forward side coupled to the first portion of the disc/rotor assembly and an aft side coupled to the second portion of the disc/rotor assembly. The seal housing apparatus is coupled to the annular intermediate disc so as to be rotatable with the annular intermediate disc and the disc/rotor assembly during operation of the gas turbine.

Optimization of a Centrifugal Impeller Design Through CFD Analysis

NASA Technical Reports Server (NTRS)

Chen, W. C.; Eastland, A. H.; Chan, D. C.; Garcia, Roberto

1993-01-01

This paper discusses the procedure, approach and Rocketdyne CFD results for the optimization of the NASA consortium impeller design. Two different approaches have been investigated. The first one is to use a tandem blade arrangement, the main impeller blade is split into two separate rows with the second blade row offset circumferentially with respect to the first row. The second approach is to control the high losses related to secondary flows within the impeller passage. Many key parameters have been identified and each consortium team member involved will optimize a specific parameter using 3-D CFD analysis. Rocketdyne has provided a series of CFD grids for the consortium team members. SECA will complete the tandem blade study, SRA will study the effect of the splitter blade solidity change, NASA LeRC will evaluate the effect of circumferential position of the splitter blade, VPI will work on the hub to shroud blade loading distribution, NASA Ames will examine the impeller discharge leakage flow impacts and Rocketdyne will continue to work on the meridional contour and the blade leading to trailing edge work distribution. This paper will also present Rocketdyne results from the tandem blade study and from the blade loading distribution study. It is the ultimate goal of this consortium team to integrate the available CFD analysis to design an advanced technology impeller that is suitable for use in the NASA Space Transportation Main Engine (STME) fuel turbopump.

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.

Turbine blade unsteady aerodynamic loading and heat transfer

NASA Astrophysics Data System (ADS)

Johnston, David Alan

Stator indexing to minimize the unsteady aerodynamic loading of closely spaced airfoil rows in turbomachinery is a new technique for the passive control of flow-induced vibrations. This technique, along with the effects of steady blade loading, were studied by means of experiments performed in a two-stage low-speed research turbine. With the second vane row fixed, the inlet vane row was indexed to six positions over one vane-pitch cycle for a range of stage loadings. The aerodynamic forcing function to the first-stage rotor was measured in the rotating reference frame, with the resulting rotor blade unsteady aerodynamic response quantified by rotor blades instrumented with dynamic pressure transducers. Reductions in the unsteady lift magnitude were achieved at all turbine operating conditions, with attenuation ranging from 37% to 74% of the maximum unsteady lift. Additionally, in complementary experiments, the effects of stator indexing and steady blade loading on the unsteady heat transfer of the first- and second-stage rotors was studied for the design and highest blade loading conditions using platinum-film heat gages. The attenuation of unsteady heat transfer coefficient was blade-loading dependent and location dependent along the chord and span, ranging 10% to 90% of maximum. Due to the high degree of location dependence of attenuation, stator indexing is therefore best suited to minimize unsteady heat transfer in local hot spots of the blade rather than the blade as a whole.

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.

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.

Forcing function modeling for flow induced vibration

NASA Technical Reports Server (NTRS)

Fleeter, Sanford

1993-01-01

The fundamental forcing function unsteady aerodynamics for application to turbomachine blade row forced response are considered, accomplished through a series of experiments performed in a rotating annular cascade and a research axial flow turbine. In particular, the unsteady periodic flowfields downstream of rotating rows of perforated plates, airfoils and turbine blade rows are measured with a cross hot-wire and an unsteady total pressure probe. The unsteady velocity and static pressure fields were then analyzed harmonically and split into vortical and potential gusts, accomplished by developing a gust splitting analysis which includes both gust unsteady static pressure and velocity data. The perforated plate gusts closely were found to be linear theory vortical gusts, satisfying the vortical gust constraints. The airfoil and turbine blade row generated velocity perturbations did not satisfy the vortical gust constraints. However, the decomposition of the unsteady flow field separated the data into a propagating vortical component which satisfied these vortical gust constraints and a decaying potential component.

Application of holography to flow visualization within rotating compressor blade row. [to determine three dimensional shock patterns

NASA Technical Reports Server (NTRS)

Wuerker, R. F.; Kobayashi, R. J.; Heflinger, L. O.; Ware, T. C.

1974-01-01

Two holographic interblade row flow visualization systems were designed to determine the three-dimensional shock patterns and velocity distributions within the rotating blade row of a transonic fan rotor, utilizing the techniques of pulsed laser transmission holography. Both single- and double-exposure bright field holograms and dark field scattered-light holograms were successfully recorded. Two plastic windows were installed in the rotor tip casing and outer casing forward of the rotor to view the rotor blade passage. The viewing angle allowed detailed investigation of the leading edge shocks and shocks in the midspan damper area; limited details of the trailing edge shocks also were visible. A technique was devised for interpreting the reconstructed holograms by constructing three dimensional models that allowed identification of the major shock systems. The models compared favorably with theoretical predictions and results of the overall and blade element data. Most of the holograms were made using the rapid double-pulse technique.

Channel flow analysis. [velocity distribution throughout blade flow field

NASA Technical Reports Server (NTRS)

Katsanis, T.

1973-01-01

The design of a proper blade profile requires calculation of the blade row flow field in order to determine the velocities on the blade surfaces. An analysis theory is presented for several methods used for this calculation and associated computer programs that were developed are discussed.

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.

Investigation of two-stage air-cooled turbine suitable for flight at Mach number of 2.5 II : blade design

NASA Technical Reports Server (NTRS)

Miser, James W; Stewart, Warner L

1957-01-01

A blade design study is presented for a two-stage air-cooled turbine suitable for flight at a Mach number of 2.5 for which velocity diagrams have been previously obtained. The detailed procedure used in the design of the blades is given. In addition, the design blade shapes, surface velocity distributions, inner and outer wall contours, and other design data are presented. Of all the blade rows, the first-stage rotor has the highest solidity, with a value of 2.289 at the mean section. The second-stage stator also had a high mean-section solidity of 1.927, mainly because of its high inlet whirl. The second-stage rotor has the highest value of the suction-surface diffusion parameter, with a value of 0.151. All other blade rows have values for this parameter under 0.100.

Estimation of Efficiency of the Cooling Channel of the Nozzle Blade of Gas-Turbine Engines

NASA Astrophysics Data System (ADS)

Vikulin, A. V.; Yaroslavtsev, N. L.; Zemlyanaya, V. A.

2018-02-01

The main direction of improvement of gas-turbine plants (GTP) and gas-turbine engines (GTE) is increasing the gas temperature at the turbine inlet. For the solution of this problem, promising systems of intensification of heat exchange in cooled turbine blades are developed. With this purpose, studies of the efficiency of the cooling channel of the nozzle blade in the basic modification and of the channel after constructive measures for improvement of the cooling system by the method of calorimetry in a liquid-metal thermostat were conducted. The combined system of heat-exchange intensification with the complicated scheme of branched channels is developed; it consists of a vortex matrix and three rows of inclined intermittent trip strips. The maximum value of hydraulic resistance ξ is observed at the first row of the trip strips, which is connected with the effect of dynamic impact of airflow on the channel walls, its turbulence, and rotation by 117° at the inlet to the channels formed by the trip strips. These factors explain the high value of hydraulic resistance equal to 3.7-3.4 for the first row of the trip strips. The obtained effect was also confirmed by the results of thermal tests, i.e., the unevenness of heat transfer on the back and on the trough of the blade is observed at the first row of the trip strips, which amounts 8-12%. This unevenness has a fading character; at the second row of the trip strips, it amounts to 3-7%, and it is almost absent at the third row. At the area of vortex matrix, the intensity of heat exchange on the blade back is higher as compared to the trough, which is explained by the different height of the matrix ribs on its opposite sides. The design changes in the nozzle blade of basic modification made it possible to increase the intensity of heat exchange by 20-50% in the area of the vortex matrix and by 15-30% on the section of inclined intermittent trip strips. As a result of research, new criteria dependences for the complicated systems of heat exchange intensification were obtained. The design of nozzle blades can be used when developing the promising high-temperature gas turbines.

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.

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.

Computer aided design and manufacturing of composite propfan blades for a cruise missile wind tunnel model

NASA Technical Reports Server (NTRS)

Thorp, Scott A.; Downey, Kevin M.

1992-01-01

One of the propulsion concepts being investigated for future cruise missiles is advanced unducted propfans. To support the evaluation of this technology applied to the cruise missile, a joint DOD and NASA test project was conducted to design and then test the characteristics of the propfans on a 0.55-scale, cruise missile model in a NASA wind tunnel. The configuration selected for study is a counterrotating rearward swept propfan. The forward blade row, having six blades, rotates in a counterclockwise direction, and the aft blade row, having six blades, rotates in a clockwise direction, as viewed from aft of the test model. Figures show the overall cruise missile and propfan blade configurations. The objective of this test was to evaluate propfan performance and suitability as a viable propulsion option for next generation of cruise missiles. This paper details the concurrent computer aided design, engineering, and manufacturing of the carbon fiber/epoxy propfan blades as the NASA Lewis Research Center.

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.

Development of a linearized unsteady Euler analysis for turbomachinery blade rows

NASA Technical Reports Server (NTRS)

Verdon, Joseph M.; Montgomery, Matthew D.; Kousen, Kenneth A.

1995-01-01

A linearized unsteady aerodynamic analysis for axial-flow turbomachinery blading is described in this report. The linearization is based on the Euler equations of fluid motion and is motivated by the need for an efficient aerodynamic analysis that can be used in predicting the aeroelastic and aeroacoustic responses of blade rows. The field equations and surface conditions required for inviscid, nonlinear and linearized, unsteady aerodynamic analyses of three-dimensional flow through a single, blade row operating within a cylindrical duct, are derived. An existing numerical algorithm for determining time-accurate solutions of the nonlinear unsteady flow problem is described, and a numerical model, based upon this nonlinear flow solver, is formulated for the first-harmonic linear unsteady problem. The linearized aerodynamic and numerical models have been implemented into a first-harmonic unsteady flow code, called LINFLUX. At present this code applies only to two-dimensional flows, but an extension to three-dimensions is planned as future work. The three-dimensional aerodynamic and numerical formulations are described in this report. Numerical results for two-dimensional unsteady cascade flows, excited by prescribed blade motions and prescribed aerodynamic disturbances at inlet and exit, are also provided to illustrate the present capabilities of the LINFLUX analysis.

On the transonic aerodynamics of a compressor blade row

NASA Technical Reports Server (NTRS)

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

1971-01-01

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

Numerical study of a high-speed miniature centrifugal compressor

NASA Astrophysics Data System (ADS)

Li, Xiaoyi

A miniature centrifugal compressor is a key component of reverse Brayton cycle cryogenic cooling system. The system is commonly used to generate a low cryogenic temperature environment for electronics to increase their efficiency, or generate, store and transport cryogenic liquids, such as liquid hydrogen and oxygen, where space limit is also an issue. Because of space limitation, the compressor is composed of a radial IGV, a radial impeller and an axial-direction diffuser (which reduces the radial size because of smaller diameter). As a result of reduction in size, rotating speed of the impeller is as high as 313,000 rpm, and Helium is used as the working fluid, in order to obtain the required static pressure ratio/rise. Two main characteristics of the compressor---miniature and high-speed, make it distinct from conventional compressors. Higher compressor efficiency is required to obtain a higher COP (coefficient of performance) system. Even though miniature centrifugal compressors start to draw researchers' attention in recent years, understanding of the performance and loss mechanism is still lacking. Since current experimental techniques are not advanced enough to capture details of flow at miniature scale, numerical methods dominate miniature turbomachinery study. This work numerically studied a high speed miniature centrifugal compressor with commercial CFD code. The overall performance of the compressor was predicted with consideration of interaction between blade rows by using sliding mesh model. The law of similarity of turbomachinery was validated for small scale machines. It was found that the specific ratio effect needs to be considered when similarity law is applied. But Reynolds number effect can be neglected. The loss mechanism of each component was analyzed. Loss due to turning bend was significant in each component. Tip leakage loss of small scale turbomachines has more impact on the impeller performance than that of large scale ones. Because the splitter was located at downstream of the impeller leading edge, any incidence at the impeller leading edge could deteriorate the splitter performance. Therefore, the impeller with twenty blades had, higher isentropic efficiency than the impeller with ten blades and ten splitters. Based on numerical study, a four-row vaned diffuser replaced a two-row vaned diffuser. It was found that the four-row vaned diffuser had much higher pressure recovery coefficient than the two-row vaned diffuser. However, most of pressure numerically is found to be recovered at the first two rows of diffuser vanes. Consequently, the following suggestions were given to further improve the performance of the miniature centrifugal compressor. (1) Redesign inlet guide vane based on the numerical simulation and experimental results. (2) Add de-swirl vanes in front of the diffuser and before the bend. (3) Replace the current impeller with a twenty-blade impeller. (4) Remove the last two rows of diffuser.

Mean velocity and decay characteristics of the guidevane and stator blade wake of an axial flow compressor

NASA Technical Reports Server (NTRS)

Lakshminarayana, B.; Davino, R.

1979-01-01

Pure tone noise, blade row vibrations, and aerodynamic losses are phenomena which are influenced by stator and IGV (inlet guide vane) blade wake production, decay, and interaction in an axial-flow compressor. The objective of this investigation is to develop a better understanding of the nature of stator and IGV blade wakes that are influenced by the presence of centrifugal forces due to flow curvature. A single sensor hot wire probe was employed to determine the three mean velocity components of stator and IGV wakes of a single stage compressor. These wake profiles indicated a varying decay rate of the tangential and axial wake velocity components and a wake profile similarity. An analysis, which predicts this trend, has been developed. The radial velocities are found to be appreciable in both IGV and the stator wakes.

Core Engine Noise Control Program. Volume III. Prediction Methods

DTIC Science & Technology

1974-08-01

turbofan engines , and Method (C) is based on an analytical description of viscous wake interaction between adjoining blade rows. Turbine Tone/ Jet ...levels for turbojet , turboshaft and turbofan engines . The turbojet data correlate highest and the turbofan data correlate lowest. Turbine Noise Noise...different engines were examined for combustor, jet and fan noise. Tnree turbojet , two turboshaft and two turbofan

Experimental measurements and analytical analysis related to gas turbine heat transfer. Part 1: Time-averaged heat-flux and surface-pressure measurements on the vanes and blades of the SSME fuel-side turbine and comparison with prediction. Part 2: Phase-resolved surface-pressure and heat-flux measurements on the first blade of the SSME fuel-side turbine

NASA Technical Reports Server (NTRS)

1994-01-01

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 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 a 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.

Assembly for directing combustion gas

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

Charron, Richard C.; Little, David A.; Snyder, Gary D.

2016-04-12

An arrangement is provided for delivering gases from a plurality of combustors of a can-annular gas turbine combustion engine to a first row of turbine blades including a first row of turbine blades. The arrangement includes a gas path cylinder, a cone and an integrated exit piece (IEP) for each combustor. Each IEP comprises an inlet chamber for receiving a gas flow from a respective combustor, and includes a connection segment. The IEPs are connected together to define an annular chamber extending circumferentially and concentric to an engine longitudinal axis, for delivering the gas flow to the first row ofmore » blades. A radiused joint extends radially inward from a radially outer side of the inlet chamber to an outer boundary of the annular chamber, and a flared fillet extends radially inward from a radially inner side of the inlet chamber to an inner boundary of the annular chamber.« less

Blade row dynamic digital compressor program. Volume 1: J85 clean inlet flow and parallel compressor models

NASA Technical Reports Server (NTRS)

Tesch, W. A.; Steenken, W. G.

1976-01-01

The results are presented of a one-dimensional dynamic digital blade row compressor model study of a J85-13 engine operating with uniform and with circumferentially distorted inlet flow. Details of the geometry and the derived blade row characteristics used to simulate the clean inlet performance are given. A stability criterion based upon the self developing unsteady internal flows near surge provided an accurate determination of the clean inlet surge line. The basic model was modified to include an arbitrary extent multi-sector parallel compressor configuration for investigating 180 deg 1/rev total pressure, total temperature, and combined total pressure and total temperature distortions. The combined distortions included opposed, coincident, and 90 deg overlapped patterns. The predicted losses in surge pressure ratio matched the measured data trends at all speeds and gave accurate predictions at high corrected speeds where the slope of the speed lines approached the vertical.

Magnetic particle testing of turbine blades mounted on the turbine rotor shaft

NASA Astrophysics Data System (ADS)

Imbert, Clement; Rampersad, Krishna

1992-07-01

An outline is presented of the general technique of magnetic particle inspection (MPI) of turbine blades mounted on the turbine rotor shaft with specific reference to the placement of the magnetizing coils. In particular, this study reports on the use of MPI in the examination of martensitic stainless steel turbine blades in power plants in Trinidad and Tobago in order to establish procedures for the detection of discontinuities. The techniques described are applicable to ferromagnetic turbine blades in general. The two practical techniques mentioned are the method of placing a preformed coil over a number of blades in one row and the method of wrapping the coil around the rotor shaft across an entire row of blades. Of the two methods, the former is preferred to the latter one, because there is greater uniformity of magnetic flux induced and lower current required to induce adequate flux density with the preformed coil. However, both methods provide satisfactory magnetic flux, and either can be used.

Effects of vane/blade ratio and spacing on fan noise, volume 1

NASA Technical Reports Server (NTRS)

Gliebe, P. R.; Kantola, R. A.

1983-01-01

The noise characteristics of a high-speed fan were studied. The experimental investigation was carried out on a 50.8 cm (20 in.) diameter scale model fan stage in an anechoic chamber with an inflow turbulence control screen installed. The forty-four blade rotor was tested with forty-eight vane and eighty-six vane stator rows, over a range of aixal rotor-stator spacings from 0.5 to 2.3 rotor tip chords. A two-dimensional strip theory model of rotor-stator interaction noise was employed to predict the measured tone power level trends, and good overall agreement with measured trends was obtained.

Aerodynamics of advanced axial-flow turbomachinery

NASA Technical Reports Server (NTRS)

Serovy, G. K.; Kavanagh, P.; Kiishi, T. H.

1980-01-01

A multi-task research program on aerodynamic problems in advanced axial-flow turbomachine configurations was carried out at Iowa State University. The elements of this program were intended to contribute directly to the improvement of compressor, fan, and turbine design methods. Experimental efforts in intra-passage flow pattern measurements, unsteady blade row interaction, and control of secondary flow are included, along with computational work on inviscid-viscous interaction blade passage flow techniques. This final report summarizes the results of this program and indicates directions which might be taken in following up these results in future work. In a separate task a study was made of existing turbomachinery research programs and facilities in universities located in the United States. Some potentially significant research topics are discussed which might be successfully attacked in the university atmosphere.

Unsteady Cascade Aerodynamic Response Using a Multiphysics Simulation Code

NASA Technical Reports Server (NTRS)

Lawrence, C.; Reddy, T. S. R.; Spyropoulos, E.

2000-01-01

The multiphysics code Spectrum(TM) is applied to calculate the unsteady aerodynamic pressures of oscillating cascade of airfoils representing a blade row of a turbomachinery component. Multiphysics simulation is based on a single computational framework for the modeling of multiple interacting physical phenomena, in the present case being between fluids and structures. Interaction constraints are enforced in a fully coupled manner using the augmented-Lagrangian method. The arbitrary Lagrangian-Eulerian method is utilized to account for deformable fluid domains resulting from blade motions. Unsteady pressures are calculated for a cascade designated as the tenth standard, and undergoing plunging and pitching oscillations. The predicted unsteady pressures are compared with those obtained from an unsteady Euler co-de refer-red in the literature. The Spectrum(TM) code predictions showed good correlation for the cases considered.

Investigation of advanced counterrotation blade configuration concepts for high speed turboprop systems. Task 4: Advanced fan section aerodynamic analysis

NASA Technical Reports Server (NTRS)

Crook, Andrew J.; Delaney, Robert A.

1992-01-01

The purpose of this study is the development of a three-dimensional Euler/Navier-Stokes flow analysis for fan section/engine geometries containing multiple blade rows and multiple spanwise flow splitters. An existing procedure developed by Dr. J. J. Adamczyk and associates and the NASA Lewis Research Center was modified to accept multiple spanwise splitter geometries and simulate engine core conditions. The procedure was also modified to allow coarse parallelization of the solution algorithm. This document is a final report outlining the development and techniques used in the procedure. The numerical solution is based upon a finite volume technique with a four stage Runge-Kutta time marching procedure. Numerical dissipation is used to gain solution stability but is reduced in viscous dominated flow regions. Local time stepping and implicit residual smoothing are used to increase the rate of convergence. Multiple blade row solutions are based upon the average-passage system of equations. The numerical solutions are performed on an H-type grid system, with meshes being generated by the system (TIGG3D) developed earlier under this contract. The grid generation scheme meets the average-passage requirement of maintaining a common axisymmetric mesh for each blade row grid. The analysis was run on several geometry configurations ranging from one to five blade rows and from one to four radial flow splitters. Pure internal flow solutions were obtained as well as solutions with flow about the cowl/nacelle and various engine core flow conditions. The efficiency of the solution procedure was shown to be the same as the original analysis.

Effects of stroke resistance on rowing economy in club rowers post-season.

PubMed

Kane, D A; Mackenzie, S J; Jensen, R L; Watts, P B

2013-02-01

In the sport of rowing, increasing the impulse applied to the oar handle during the stroke can result in greater boat velocities; this may be facilitated by increasing the surface area of the oar blade and/or increasing the length of the oars. The purpose of this study was to compare the effects of different rowing resistances on the physiological response to rowing. 5 male and 7 female club rowers completed progressive, incremental exercise tests on an air-braked rowing ergometer, using either low (LO; 100) or high (HI; 150) resistance (values are according to the adjustable "drag factor" setting on the ergometer). Expired air, blood lactate concentration, heart rate, rowing cadence, and ergometer power output were monitored during the tests. LO rowing elicited significantly greater cadences (P<0.01) and heart rates (P<0.05), whereas rowing economy (J · L O(2) equivalents(-1)) was significantly greater during HI rowing (P<0.05). These results suggest that economically, rowing with a greater resistance may be advantageous for performance. Moreover, biomechanical analysis of ergometer rowing support the notion that the impulse generated during the stroke increases positively as a function of rowing resistance. We conclude that an aerobic advantage associated with greater resistance parallels the empirical trend toward larger oar blades in competitive rowing. This may be explained by a greater stroke impulse at the higher resistance. © Georg Thieme Verlag KG Stuttgart · New York.

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.

Interaction of upstream flow distortions with high Mach number cascades

NASA Technical Reports Server (NTRS)

Englert, G. W.

1981-01-01

Features of the interaction of flow distortions, such as gusts and wakes with blade rows of advance type fans and compressors having high tip Mach numbers are modeled. A typical disturbance was assumed to have harmonic time dependence and was described, at a far upstream location, in three orthogonal spatial coordinates by a double Fourier series. It was convected at supersonic relative to a linear cascade described as an unrolled annulus. Conditions were selected so that the component of this velocity parallel to the axis of the turbomachine was subsonic, permitting interaction between blades through the upstream as well as downstream flow media. A strong, nearly normal shock was considered in the blade passages which was allowed curvature and displacement. The flows before and after the shock were linearized relative to uniform mean velocities in their respective regions. Solution of the descriptive equations was by adaption of the Wiener-Hopf technique, enabling a determination of distortion patterns through and downstream of the cascade as well as pressure distributions on the blade and surfaces. Details of interaction of the disturbance with the in-passage shock were discussed. Infuences of amplitude, wave length, and phase of the disturbance on lifts and moments of cascade configurations are presented. Numerical results are clarified by reference to an especially orderly pattern of upstream vertical motion in relation to the cascade parameters.

Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines

NASA Technical Reports Server (NTRS)

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

2015-01-01

A gas turbine engine includes a spool, a turbine coupled to drive the spool, and a propulsor that is coupled to be driven by the turbine through the spool. A gear assembly is coupled between the propulsor and the spool such that rotation of the turbine drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. The row includes no more than 20 of the propulsor blades.

Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines

NASA Technical Reports Server (NTRS)

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

2016-01-01

A gas turbine engine includes a spool, a turbine coupled to drive the spool, and a propulsor that is coupled to be driven by the turbine through the spool. A gear assembly is coupled between the propulsor and the spool such that rotation of the turbine drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extends from the hub. The row includes no more than 20 of the propulsor blades.

Cutting blade dentitions in squaliform sharks form by modification of inherited alternate tooth ordering patterns

PubMed Central

Smith, Moya Meredith

2016-01-01

The squaliform sharks represent one of the most speciose shark clades. Many adult squaliforms have blade-like teeth, either on both jaws or restricted to the lower jaw, forming a continuous, serrated blade along the jaw margin. These teeth are replaced as a single unit and successor teeth lack the alternate arrangement present in other elasmobranchs. Micro-CT scans of embryos of squaliforms and a related outgroup (Pristiophoridae) revealed that the squaliform dentition pattern represents a highly modified version of tooth replacement seen in other clades. Teeth of Squalus embryos are arranged in an alternate pattern, with successive tooth rows containing additional teeth added proximally. Asynchronous timing of tooth production along the jaw and tooth loss prior to birth cause teeth to align in oblique sets containing teeth from subsequent rows; these become parallel to the jaw margin during ontogeny, so that adult Squalus has functional tooth rows comprising obliquely stacked teeth of consecutive developmental rows. In more strongly heterodont squaliforms, initial embryonic lower teeth develop into the oblique functional sets seen in adult Squalus, with no requirement to form, and subsequently lose, teeth arranged in an initial alternate pattern. PMID:28018617

Cutting blade dentitions in squaliform sharks form by modification of inherited alternate tooth ordering patterns

NASA Astrophysics Data System (ADS)

Underwood, Charlie; Johanson, Zerina; Smith, Moya Meredith

2016-11-01

The squaliform sharks represent one of the most speciose shark clades. Many adult squaliforms have blade-like teeth, either on both jaws or restricted to the lower jaw, forming a continuous, serrated blade along the jaw margin. These teeth are replaced as a single unit and successor teeth lack the alternate arrangement present in other elasmobranchs. Micro-CT scans of embryos of squaliforms and a related outgroup (Pristiophoridae) revealed that the squaliform dentition pattern represents a highly modified version of tooth replacement seen in other clades. Teeth of Squalus embryos are arranged in an alternate pattern, with successive tooth rows containing additional teeth added proximally. Asynchronous timing of tooth production along the jaw and tooth loss prior to birth cause teeth to align in oblique sets containing teeth from subsequent rows; these become parallel to the jaw margin during ontogeny, so that adult Squalus has functional tooth rows comprising obliquely stacked teeth of consecutive developmental rows. In more strongly heterodont squaliforms, initial embryonic lower teeth develop into the oblique functional sets seen in adult Squalus, with no requirement to form, and subsequently lose, teeth arranged in an initial alternate pattern.

Bending mode flutter in a transonic linear cascade

NASA Astrophysics Data System (ADS)

Govardhan, Raghuraman; Jutur, Prahallada

2017-11-01

Vibration related issues like flutter pose a serious challenge to aircraft engine designers. The phenomenon has gained relevance for modern engines that employ thin and long fan blade rows to satisfy the growing need for compact and powerful engines. The tip regions of such blade rows operate with transonic relative flow velocities, and are susceptible to bending mode flutter. In such cases, the flow field around individual blades of the cascade is dominated by shock motions generated by the blade motions. In the present work, a new transonic linear cascade facility with the ability to oscillate a blade at realistic reduced frequencies has been developed. The facility operates at a Mach number of 1.3, with the central blade being oscillated in heave corresponding to the bending mode of the rotor. The susceptibility of the blade to undergo flutter at different reduced frequencies is quantified by the cycle-averaged power transfer to the blade calculated using the measured unsteady load on the oscillating blade. These measurements show fluid excitation (flutter) at low reduced frequencies and fluid damping (no flutter) at higher reduced frequencies. Simultaneous measurements of the unsteady shock motions are done with high speed shadowgraphy to elucidate the differences in shock motions between the excitation and damping cases.

Computation of rotor-stator interaction using the Navier-Stokes equations

NASA Technical Reports Server (NTRS)

Whitfield, David L.; Chen, Jen-Ping

1995-01-01

The numerical scheme presented belongs to a family of codes known as UNCLE (UNsteady Computation of fieLd Equations) as reported by Whitfield (1995), that is being used to solve problems in a variety of areas including compressible and incompressible flows. This derivation is specifically developed for general unsteady multi-blade-row turbomachinery problems. The scheme solves the Reynolds-averaged N-S equations with the Baldwin-Lomax turbulence model.

Changes in Blade Configuration Improve Turbopump

NASA Technical Reports Server (NTRS)

Meng, S. Y.; Bache, G. E.

1987-01-01

Cavitation reduced while suction increased. Tests conducted with model liquid-oxygen turbopump using water as pumped fluid confirms performance improved by "tandem" arrangement of blades. Findings expected to apply to other pumps having two adjacent rotor rows.

A Numerical Simulator for Three-Dimensional Flows Through Vibrating Blade Rows

NASA Technical Reports Server (NTRS)

Chuang, H. Andrew; Verdon, Joseph M.

1998-01-01

The three-dimensional, multi-stage, unsteady, turbomachinery analysis, TURBO, has been extended to predict the aeroelastic and aeroacoustic response behaviors of a single blade row operating within a cylindrical annular duct. In particular, a blade vibration capability has been incorporated so that the TURBO analysis can be applied over a solution domain that deforms with a vibratory blade motion. Also, unsteady far-field conditions have been implemented to render the computational boundaries at inlet and exit transparent to outgoing unsteady disturbances. The modified TURBO analysis is applied herein to predict unsteady subsonic and transonic flows. The intent is to partially validate this nonlinear analysis for blade flutter applications, via numerical results for benchmark unsteady flows, and to demonstrate the analysis for a realistic fan rotor. For these purposes, we have considered unsteady subsonic flows through a 3D version of the 10th Standard Cascade, and unsteady transonic flows through the first stage rotor of the NASA Lewis, Rotor 67, two-stage fan.

Design of an Object-Oriented Turbomachinery Analysis Code: Initial Results

NASA Technical Reports Server (NTRS)

Jones, Scott

2015-01-01

Performance prediction of turbomachines is a significant part of aircraft propulsion design. In the conceptual design stage, there is an important need to quantify compressor and turbine aerodynamic performance and develop initial geometry parameters at the 2-D level prior to more extensive Computational Fluid Dynamics (CFD) analyses. The Object-oriented Turbomachinery Analysis Code (OTAC) is being developed to perform 2-D meridional flowthrough analysis of turbomachines using an implicit formulation of the governing equations to solve for the conditions at the exit of each blade row. OTAC is designed to perform meanline or streamline calculations; for streamline analyses simple radial equilibrium is used as a governing equation to solve for spanwise property variations. While the goal for OTAC is to allow simulation of physical effects and architectural features unavailable in other existing codes, it must first prove capable of performing calculations for conventional turbomachines.OTAC is being developed using the interpreted language features available in the Numerical Propulsion System Simulation (NPSS) code described by Claus et al (1991). Using the NPSS framework came with several distinct advantages, including access to the pre-existing NPSS thermodynamic property packages and the NPSS Newton-Raphson solver. The remaining objects necessary for OTAC were written in the NPSS framework interpreted language. These new objects form the core of OTAC and are the BladeRow, BladeSegment, TransitionSection, Expander, Reducer, and OTACstart Elements. The BladeRow and BladeSegment consumed the initial bulk of the development effort and required determining the equations applicable to flow through turbomachinery blade rows given specific assumptions about the nature of that flow. Once these objects were completed, OTAC was tested and found to agree with existing solutions from other codes; these tests included various meanline and streamline comparisons of axial compressors and turbines at design and off-design conditions.

Design of an Object-Oriented Turbomachinery Analysis Code: Initial Results

NASA Technical Reports Server (NTRS)

Jones, Scott M.

2015-01-01

Performance prediction of turbomachines is a significant part of aircraft propulsion design. In the conceptual design stage, there is an important need to quantify compressor and turbine aerodynamic performance and develop initial geometry parameters at the 2-D level prior to more extensive Computational Fluid Dynamics (CFD) analyses. The Object-oriented Turbomachinery Analysis Code (OTAC) is being developed to perform 2-D meridional flowthrough analysis of turbomachines using an implicit formulation of the governing equations to solve for the conditions at the exit of each blade row. OTAC is designed to perform meanline or streamline calculations; for streamline analyses simple radial equilibrium is used as a governing equation to solve for spanwise property variations. While the goal for OTAC is to allow simulation of physical effects and architectural features unavailable in other existing codes, it must first prove capable of performing calculations for conventional turbomachines. OTAC is being developed using the interpreted language features available in the Numerical Propulsion System Simulation (NPSS) code described by Claus et al (1991). Using the NPSS framework came with several distinct advantages, including access to the pre-existing NPSS thermodynamic property packages and the NPSS Newton-Raphson solver. The remaining objects necessary for OTAC were written in the NPSS framework interpreted language. These new objects form the core of OTAC and are the BladeRow, BladeSegment, TransitionSection, Expander, Reducer, and OTACstart Elements. The BladeRow and BladeSegment consumed the initial bulk of the development effort and required determining the equations applicable to flow through turbomachinery blade rows given specific assumptions about the nature of that flow. Once these objects were completed, OTAC was tested and found to agree with existing solutions from other codes; these tests included various meanline and streamline comparisons of axial compressors and turbines at design and off-design conditions.

Theory of low frequency noise transmission through turbines

NASA Technical Reports Server (NTRS)

Matta, R. K.; Mani, R.

1979-01-01

Improvements of the existing theory of low frequency noise transmission through turbines and development of a working prediction tool are described. The existing actuator-disk model and a new finite-chord model were utilized in an analytical study. The interactive effect of adjacent blade rows, higher order spinning modes, blade-passage shocks, and duct area variations were considered separately. The improved theory was validated using the data acquired in an earlier NASA program. Computer programs incorporating the improved theory were produced for transmission loss prediction purposes. The programs were exercised parametrically and charts constructed to define approximately the low frequency noise transfer through turbines. The loss through the exhaust nozzle and flow(s) was also considered.

Three-dimensional effects on pure tone fan noise due to inflow distortion. [rotor blade noise prediction

NASA Technical Reports Server (NTRS)

Kobayashi, H.

1978-01-01

Two dimensional, quasi three dimensional and three dimensional theories for the prediction of pure tone fan noise due to the interaction of inflow distortion with a subsonic annular blade row were studied with the aid of an unsteady three dimensional lifting surface theory. The effects of compact and noncompact source distributions on pure tone fan noise in an annular cascade were investigated. Numerical results show that the strip theory and quasi three-dimensional theory are reasonably adequate for fan noise prediction. The quasi three-dimensional method is more accurate for acoustic power and model structure prediction with an acoustic power estimation error of about plus or minus 2db.

Computational Study of the Impact of Unsteadiness on the Aerodynamic Performance of a Variable- Speed Power Turbine

NASA Technical Reports Server (NTRS)

Welch, Gerard E.

2012-01-01

The design-point and off-design performance of an embedded 1.5-stage portion of a variable-speed power turbine (VSPT) was assessed using Reynolds-Averaged Navier-Stokes (RANS) analyses with mixing-planes and sector-periodic, unsteady RANS analyses. The VSPT provides one means by which to effect the nearly 50 percent main-rotor speed change required for the NASA Large Civil Tilt-Rotor (LCTR) application. The change in VSPT shaft-speed during the LCTR mission results in blade-row incidence angle changes of as high as 55 . Negative incidence levels of this magnitude at takeoff operation give rise to a vortical flow structure in the pressure-side cove of a high-turn rotor that transports low-momentum flow toward the casing endwall. The intent of the effort was to assess the impact of unsteadiness of blade-row interaction on the time-mean flow and, specifically, to identify potential departure from the predicted trend of efficiency with shaft-speed change of meanline and 3-D RANS/mixing-plane analyses used for design.

In-service inspection of steam turbine blades without disassembly

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

Reinhart, E.R.

1987-01-01

Loss of utility plant availability as a result of failure-causing cracks in steam turbine blades makes early detection of this problem critical. An Electric Power Research Institute survey, conducted as part of project RP 1266-24, indicated that 72% of turbine blade failures in fossil power plants occur in low-pressure (LP) turbines with half of all blade failures occurring in the last two blade stages (L-0 and L-1 rows). Failures are generally associated with blade tailing edges and root areas. Project RP 1266-24 also found that 79% of the blade problems in LP turbines were cracks. A turbine design of particularmore » concern has been the Westinghouse Building Block (B.B.) 73. Reinhart and Associates has successfully inspected seven in-place B.B. 73 units for six utilities during the past 3 yr, as well as several disassembled turbines of other manufacturers and designs. These examinations consisted of visual and eddy-current examinations of the blade roots and trailing edges. The in-place inspections were performed using prototype manipulation devices to gain access to the blades through the hand holes. The only disassembly required to gain access for the examinations was the removal of the man-way covers on the main shell and the hand-hole covers on the outer cylinder covering the L-0 and L-1 blade rows.« less

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.

Design of Multistage Axial-Flow Compressors

NASA Technical Reports Server (NTRS)

Crouse, J. E.; Gorrell, W. T.

1983-01-01

Program developed for computing aerodynamic design of multistage axialflow compressor and associated blading geometry input for internal flow analysis. Aerodynamic solution gives velocity diagrams on selected streamlines of revolution at blade row edges. Program written in FORTRAN IV.

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.

Wingtip mounted, counter-rotating proprotor for tiltwing aircraft

NASA Technical Reports Server (NTRS)

Wechsler, James K. (Inventor); Rutherford, John W. (Inventor)

1995-01-01

A tiltwing aircraft, capable of in-flight conversion between a hover and forward cruise mode, employs a counter-rotating proprotor arrangement which permits a significantly increased cruise efficiency without sacrificing either the size of the conversion envelope or the wing efficiency. A benefit in hover is also provided because of the lower effective disk loading for the counter-rotating proprotor, as opposed to a single rotation proprotor of the same diameter. At least one proprotor is provided on each wing section, preferably mounted on the wingtip, with each proprotor having two counter-rotating blade rows. Each blade row has a plurality of blades which are relatively stiff-in-plane and are mounted such that cyclic pitch adjustments may be made for hover control during flight.

Solid particle dynamic behavior through twisted blade rows

NASA Technical Reports Server (NTRS)

Hamed, A.

1982-01-01

The particle trajectory calculations provide the essential information which is required for predicting the pattern and intensity of turbomachinery erosion. Consequently, the evaluation of the machine performance deterioration due to erosion is extremely sensitive to the accuracy of the flow field and blade geometry representation in the trajectory computational model. A model is presented that is simple and efficient yet versatile and general to be applicable to axial, radial and mixed flow machines, and to inlets, nozzles, return passages and separators. The results of the computations are presented for the particle trajectories through a row of twisted vanes in the inlet flow field. The effect of the particle size on their trajectories, blade impacts, and on their redistribution and separation are discussed.

Mean-line Modeling of an Axial Turbine

NASA Astrophysics Data System (ADS)

Tkachenko, A. Yu; Ostapyuk, Ya A.; Filinov, E. P.

2018-01-01

The article describes the approach for axial turbine modeling along the mean line. It bases on the developed model of an axial turbine blade row. This model is suitable for both nozzle vanes and rotor blades simulations. Consequently, it allows the simulation of the single axial turbine stage as well as a multistage turbine. The turbine stage model can take into account the cooling air flow before and after a throat of each blade row, outlet straightener vanes existence and stagger angle controlling of nozzle vanes. The axial turbine estimation method includes the loss estimation and thermogasdynamic analysis. The single stage axial turbine was calculated with the developed model. The obtained results deviation was within 3% when comparing with the results of CFD modeling.

Two stage low noise advanced technology fan. 1: Aerodynamic, structural, and acoustic design

NASA Technical Reports Server (NTRS)

Messenger, H. E.; Ruschak, J. T.; Sofrin, T. G.

1974-01-01

A two-stage fan was designed to reduce noise 20 db below current requirements. The first-stage rotor has a design tip speed of 365.8 m/sec and a hub/tip ratio of 0.4. The fan was designed to deliver a pressure ratio of 1.9 with an adiabatic efficiency of 85.3 percent at a specific inlet corrected flow of 209.2kg/sec/sq m. Noise reduction devices include acoustically treated casing walls, a flowpath exit acoustic splitter, a translating centerbody sonic inlet device, widely spaced blade rows, and the proper ratio of blades and vanes. Multiple-circular-arc rotor airfoils, resettable stators, split outer casings, and capability to go to close blade-row spacing are also included.

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.

Validated biomechanical model for efficiency and speed of rowing.

PubMed

Pelz, Peter F; Vergé, Angela

2014-10-17

The speed of a competitive rowing crew depends on the number of crew members, their body mass, sex and the type of rowing-sweep rowing or sculling. The time-averaged speed is proportional to the rower's body mass to the 1/36th power, to the number of crew members to the 1/9th power and to the physiological efficiency (accounted for by the rower's sex) to the 1/3rd power. The quality of the rowing shell and propulsion system is captured by one dimensionless parameter that takes the mechanical efficiency, the shape and drag coefficient of the shell and the Froude propulsion efficiency into account. We derive the biomechanical equation for the speed of rowing by two independent methods and further validate it by successfully predicting race times. We derive the theoretical upper limit of the Froude propulsion efficiency for low viscous flows. This upper limit is shown to be a function solely of the velocity ratio of blade to boat speed (i.e., it is completely independent of the blade shape), a result that may also be of interest for other repetitive propulsion systems. Copyright © 2014 Elsevier Ltd. All rights reserved.

Experimental analysis of the aerodynamic performance of an innovative low pressure turbine rotor

NASA Astrophysics Data System (ADS)

Infantino, Daniele; Satta, Francesca; Simoni, Daniele; Ubaldi, Marina; Zunino, Pietro; Bertini, Francesco

2016-02-01

In the present work the aerodynamic performances of an innovative rotor blade row have been experimentally investigated. Measurements have been carried out in a large scale low speed single stage cold flow facility at a Reynolds number typical of aeroengine cruise, under nominal and off-design conditions. The time-mean blade aerodynamic loadings have been measured at three radial positions along the blade height through a pressure transducer installed inside the hollow shaft, by delivering the signal to the stationary frame with a slip ring. The time mean aerodynamic flow fields upstream and downstream of the rotor have been measured by means of a five-hole probe to investigate the losses associated with the rotor. The investigations in the single stage research turbine allow the reproduction of both wake-boundary layer interaction as well as vortex-vortex interaction. The detail of the present results clearly highlights the strong dissipative effects induced by the blade tip vortex and by the momentum defect as well as the turbulence production, which is generated during the migration of the stator wake in the rotor passage. Phase-locked hot-wire investigations have been also performed to analyze the time-varying flow during the wake passing period. In particular the interaction between stator and rotor structures has been investigated also under off-design conditions to further explain the mechanisms contributing to the loss generation for the different conditions.

A Numerical Investigation of Turbine Noise Source Hierarchy and Its Acoustic Transmission Characteristics

NASA Technical Reports Server (NTRS)

VanZante, Dale; Envia, Edmane

2008-01-01

Understanding the relative importance of the various turbine noise generation mechanisms and the characteristics of the turbine acoustic transmission loss are essential ingredients in developing robust reduced-order models for predicting the turbine noise signature. A computationally based investigation has been undertaken to help guide the development of a turbine noise prediction capability that does not rely on empiricism. The investigation relies on highly detailed numerical simulations of the unsteady flowfield inside a modern high-pressure turbine (HPT). The simulations are developed using TURBO, which is an unsteady Reynolds-averaged Navier-Stokes (URANS) code capable of multi-stage simulations. The purpose of this study is twofold. First, to determine an estimate of the relative importance of the contributions to the coherent part of the acoustic signature of a turbine from the three potential sources of turbine noise generation, namely, blade-row viscous interaction, potential field interaction, and entropic source associated with the interaction of the blade rows with the temperature nonuniformities caused by the incomplete mixing of the hot fluid and the cooling flow. Second, to develop an understanding of the turbine acoustic transmission characteristics and to assess the applicability of existing empirical and analytical transmission loss models to realistic geometries and flow conditions for modern turbine designs. The investigation so far has concentrated on two simulations: (1) a single-stage HPT and (2) a two-stage HPT and the associated inter-turbine duct/strut segment. The simulations are designed to resolve up to the second harmonic of the blade passing frequency tone in accordance with accepted rules for second order solvers like TURBO. The calculations include blade and vane cooling flows and a radial profile of pressure and temperature at the turbine inlet. The calculation can be modified later to include the combustor pattern factor at the turbine inlet to include that contribution to turbine noise. We shall present preliminary analysis of the results obtained so far in order to assess the validity of such an approach and to seek feedback on improving the approach. This work addresses both Area 1 (Turbine Tone Noise) and Area 5 (Influence of the Turbine on Combustor Noise) topics.

Heat Transfer on a Film-Cooled Rotating Blade

NASA Technical Reports Server (NTRS)

Garg, Vijay K.

1999-01-01

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

Unsteady Flows in a Single-Stage Transonic Axial-Flow Fan Stator Row. Ph.D. Thesis - Iowa State Univ.

NASA Technical Reports Server (NTRS)

Hathaway, Michael D.

1986-01-01

Measurements of the unsteady velocity field within the stator row of a transonic axial-flow fan were acquired using a laser anemometer. Measurements were obtained on axisymmetric surfaces located at 10 and 50 percent span from the shroud, with the fan operating at maximum efficiency at design speed. The ensemble-average and variance of the measured velocities are used to identify rotor-wake-generated (deterministic) unsteadiness and turbulence, respectively. Correlations of both deterministic and turbulent velocity fluctuations provide information on the characteristics of unsteady interactions within the stator row. These correlations are derived from the Navier-Stokes equation in a manner similar to deriving the Reynolds stress terms, whereby various averaging operators are used to average the aperiodic, deterministic, and turbulent velocity fluctuations which are known to be present in multistage turbomachines. The correlations of deterministic and turbulent velocity fluctuations throughout the axial fan stator row are presented. In particular, amplification and attenuation of both types of unsteadiness are shown to occur within the stator blade passage.

Efficient, Low Pressure Ratio Propulsor for Gas Turbine Engines

NASA Technical Reports Server (NTRS)

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

2018-01-01

A gas turbine engine includes a bypass flow passage that has an inlet and defines a bypass ratio in a range of approximately 8.5 to 13.5. A fan is arranged within the bypass flow passage. A first turbine is a 5-stage turbine and is coupled with a first shaft, which is coupled with the fan. A first compressor is coupled with the first shaft and is a 3-stage compressor. A second turbine is coupled with a second shaft and is a 2-stage turbine. The fan includes a row of fan blades that extend from a hub. The row includes a number (N) of the fan blades, a solidity value (R) at tips of the fab blades, and a ratio of N/R that is from 14 to 16.

Extension of Useful Operating Range of Axial-Flow Compressors by Use of Adjustable Stator Blades

DTIC Science & Technology

1948-01-01

wmprcssor. AIRCRAFT ENGINE RESEARCH LABORATORY, NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS , CLEVELAND, OHIO, December%9, 1944. APPENDIX A ENTRANCE...ADVISORYCOMbl_E FOR AERONAUTICS either stator blades m rotor bIades can be prevented by the adjustment of the proper row of stator blades : in the first...section, Reynokls number, solidity, blade -a@e setting, and degree of turbulence. Experiments on- airfoil cascades with retarded flow (reference 6, p. 75, and

Reynolds-Averaged Navier-Stokes Studies of Low Reynolds Number Effects on the Losses in a Low Pressure Turbine

NASA Technical Reports Server (NTRS)

Dorney, Daniel J.

1996-01-01

Experimental data from jet-engine tests have indicated that unsteady blade-row interaction effects can have a significant impact on the efficiency of low-pressure turbine stages. Measured turbine efficiencies at takeoff can be as much as two points higher than those at cruise conditions. Preliminary studies indicate that Reynolds number effects may contribute to the lower efficiencies at cruise conditions. In the current study, numerical experiments have been performed to quantify the Reynolds number dependence of unsteady wake/separation bubble interaction on the performance of a low-pressure turbine.

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

NASA Technical Reports Server (NTRS)

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

1974-01-01

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

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.

Experimental study of complex flow and turbulence structure around a turbomachine rotor blade operating behind a row of Inlet Guide Vanes (IGVS)

NASA Astrophysics Data System (ADS)

Soranna, Francesco

The flow and turbulence around a rotor blade operating downstream of a row of Inlet Guide Vanes (IGV) are investigated experimentally in a refractive index matched turbomachinery facility that provides unobstructed view of the entire flow field. High resolution 2D and Stereoscopic PIV measurements are performed both at midspan and in the tip region of the rotor blade, focusing on effects of wake-blade, wake-boundary-layer and wake-wake interactions. We first examine the modification to the shape of an IGV-wake as well as to the spatial distribution of turbulence within it as the wake propagates along the rotor blade. Due to the spatially non-uniform velocity distribution, the IGV wake deforms through the rotor passage, expanding near the leading edge and shrinking near the trailing edge. The turbulence within this wake becomes spatially non-uniform and highly anisotropic as a result of interaction with the non-uniform strain rate field within the rotor passage. Several mechanisms, which are associated with rapid straining and highly non-uniform production rate (P), including negative production on the suction side of the blade, contribute to the observed trends. During IGV-wake impingement, the suction side boundary layer near the trailing edge becomes significantly thinner, with lower momentum thickness and more stable profile compared to other phases at the same location. Analysis of available terms in the integral momentum equation indicates that the phase-averaged unsteady term is the main contributor to the decrease in momentum thickness within the impinging wake. Thinning of the boundary/shear layer extends into the rotor near wake, making it narrower and increasing the phase averaged shear velocity gradients and associated production term just downstream of the trailing edge. Consequently, the turbulent kinetic energy (TKE) increases causing as much as 75% phase-dependent variations in peak TKE magnitude. Further away from the blade, the rotor wake is bent and contracted as a result of exposure to regions with high axial momentum ('jets') which fill the gaps between IGV-wakes. On the suction side of the rotor wake, contraction by the jet enhances the shear velocity gradients, and, with them, the shear production term, the dominant source of turbulence. Consequently, the Reynolds stresses and turbulent kinetic energy profiles become asymmetric across the rotor wake, with peak values located on the suction side, coinciding with the region of peak production. As the rotor wake propagates away from the blade, the process of bending and contraction by the jets continues, leading to formation of distinct wake-kinks containing regions of high turbulence, which we have coined turbulent 'hot spots'.

The use of optimization techniques to design controlled diffusion compressor blading

NASA Technical Reports Server (NTRS)

Sanger, N. L.

1982-01-01

A method for automating compressor blade design using numerical optimization, and applied to the design of a controlled diffusion stator blade row is presented. A general purpose optimization procedure is employed, based on conjugate directions for locally unconstrained problems and on feasible directions for locally constrained problems. Coupled to the optimizer is an analysis package consisting of three analysis programs which calculate blade geometry, inviscid flow, and blade surface boundary layers. The optimizing concepts and selection of design objective and constraints are described. The procedure for automating the design of a two dimensional blade section is discussed, and design results are presented.

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

Computer program for definition of transonic axial-flow compressor blade rows. [computer program for fabrication and aeroelastic analysis

NASA Technical Reports Server (NTRS)

Crouse, J. E.

1974-01-01

A method is presented for designing axial-flow compressor blading from blade elements defined on cones which pass through the blade-edge streamline locations. Each blade-element centerline is composed of two segments which are tangent to each other. The centerline and surfaces of each segment have constant change of angle with path distance. The stacking line for the blade elements can be leaned in both the axial and tangential directions. The output of the computer program gives coordinates for fabrication and properties for aeroelastic analysis for planar blade sections. These coordinates and properties are obtained by interpolation across conical blade elements. The program is structured to be coupled with an aerodynamic design program.

A three-dimensional Navier-Stokes stage analysis of the flow through a compact radial turbine

NASA Technical Reports Server (NTRS)

Heidmann, James D.

1991-01-01

A steady, three dimensional Navier-Stokes average passage computer code is used to analyze the flow through a compact radial turbine stage. The code is based upon the average passage set of equations for turbomachinery, whereby the flow fields for all passages in a given blade row are assumed to be identical while retaining their three-dimensionality. A stage solution is achieved by alternating between stator and rotor calculations, while coupling the two solutions by means of a set of axisymmetric body forces which model the absent blade row. Results from the stage calculation are compared with experimental data and with results from an isolated rotor solution having axisymmetric inlet flow quantities upstream of the vacated stator space. Although the mass-averaged loss through the rotor is comparable for both solutions, the details of the loss distribution differ due to stator effects. The stage calculation predicts smaller spanwise variations in efficiency, in closer agreement with the data. The results of the study indicate that stage analyses hold promise for improved prediction of loss mechanisms in multi-blade row turbomachinery, which could lead to improved designs through the reduction of these losses.

A three-dimensional Navier-Stokes stage analysis of the flow through a compact radial turbine

NASA Technical Reports Server (NTRS)

Heidmann, James D.

1991-01-01

A steady, three-dimensional Navier-Stokes average passage computer code is used to analyze the flow through a compact radial turbine stage. The code is based upon the average passage set of equations for turbomachinery, whereby the flow fields for all passages in a given blade row are assumed to be identical while retaining their three-dimensionality. A stage solution is achieved by alternating between stator and rotor calculations, while coupling the two solutions by means of a set of axisymmetric body forces which model the absent blade row. Results from the stage calculation are compared with experimental data and with results from an isolated rotor solution having axisymmetric inlet flow quantities upstream of the vacated stator space. Although the mass-averaged loss through the rotor is comparable for both solutions, the details of the loss distribution differ due to stator effects. The stage calculation predicts smaller spanwise variations in efficiency, in closer agreement with the data. The results of the study indicate that stage analyses hold promise for improved prediction of loss mechanisms in multi-blade row turbomachinery, which could lead to improved designs through the reduction of these losses.

A Three-Dimensional Unsteady CFD Model of Compressor Stability

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.

2006-01-01

A three-dimensional unsteady CFD code called CSTALL has been developed and used to investigate compressor stability. The code solved the Euler equations through the entire annulus and all blade rows. Blade row turning, losses, and deviation were modeled using body force terms which required input data at stations between blade rows. The input data was calculated using a separate Navier-Stokes turbomachinery analysis code run at one operating point near stall, and was scaled to other operating points using overall characteristic maps. No information about the stalled characteristic was used. CSTALL was run in a 2-D throughflow mode for very fast calculations of operating maps and estimation of stall points. Calculated pressure ratio characteristics for NASA stage 35 agreed well with experimental data, and results with inlet radial distortion showed the expected loss of range. CSTALL was also run in a 3-D mode to investigate inlet circumferential distortion. Calculated operating maps for stage 35 with 120 degree distortion screens showed a loss in range and pressure rise. Unsteady calculations showed rotating stall with two part-span stall cells. The paper describes the body force formulation in detail, examines the computed results, and concludes with observations about the code.

A comparison of measured wind park load histories with the WISPER and WISPERX load spectra

NASA Astrophysics Data System (ADS)

Kelley, N. D.

1995-01-01

The blade-loading histories from two adjacent Micon 65/13 wind turbines are compared with the variable-amplitude test-loading histories known as the WISPER and WISPERX spectra. These standardized loading sequences were developed from blade flapwise load histories taken from nine different horizontal-axis wind turbines operating under a wide range of conditions in Europe. The subject turbines covered a broad spectrum of rotor diameters, materials, and operating environments. The final loading sequences were developed as a joint effort of thirteen different European organizations. The goal was to develop a meaningful loading standard for horizontal-axis wind turbine blades that represents common interaction effects seen in service. In 1990, NREL made extensive load measurements on two adjacent Micon 65/13 wind turbines in simultaneous operation in the very turbulent environment of a large wind park. Further, before and during the collection of the loads data, comprehensive measurements of the statistics of the turbulent environment were obtained at both the turbines under test and at two other locations within the park. The trend to larger but lighter wind turbine structures has made an understanding of the expected lifetime loading history of paramount importance. Experience in the US has shown that the turbulence-induced loads associated with multi-row wind parks in general are much more severe than for turbines operating individually or within widely spaced environments. Multi-row wind parks are much more common in the US than in Europe. In this paper we report on our results in applying the methodology utilized to develop the WISPER and WISPERX standardized loading sequences using the available data from the Micon turbines. While the intended purpose of the WISPER sequences were not to represent a specific operating environment, we believe the exercise is useful, especially when a turbine design is likely to be installed in a multi-row wind park.

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.

Computer program for definition of transonic axial-flow compressor blade rows

NASA Technical Reports Server (NTRS)

Crouse, J. E.

1975-01-01

Particular type of blade element used has two segments which have centerlines and surfaces described by constant change of angle with path distance on cone. Program is result of rework of earlier program to give major gains in accuracy, reliability and speed. It also covers more steps of overall compressor design procedure.

Coupled 2-dimensional cascade theory for noise an d unsteady aerodynamics of blade row interaction in turbofans. Volume 2: Documentation for computer code CUP2D

NASA Technical Reports Server (NTRS)

Hanson, Donald B.

1994-01-01

A two dimensional linear aeroacoustic theory for rotor/stator interaction with unsteady coupling was derived and explored in Volume 1 of this report. Computer program CUP2D has been written in FORTRAN embodying the theoretical equations. This volume (Volume 2) describes the structure of the code, installation and running, preparation of the input file, and interpretation of the output. A sample case is provided with printouts of the input and output. The source code is included with comments linking it closely to the theoretical equations in Volume 1.

Gas flow path for a gas turbine engine

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

Montgomery, Matthew D.; Charron, Richard C.; Snyder, Gary D.

A duct arrangement in a can annular gas turbine engine. The gas turbine engine has a gas delivery structure for delivering gases from a plurality of combustors to an annular chamber that extends circumferentially and is oriented concentric to a gas turbine engine longitudinal axis for delivering the gas flow to a first row of blades A gas flow path is formed by the duct arrangement between a respective combustor and the annular chamber for conveying gases from each combustor to the first row of turbine blades The duct arrangement includes at least one straight section having a centerline thatmore » is misaligned with a centerline of the combustor.« less

Highly Loaded Low-Pressure Turbine: Design, Numerical and Experimental Analysis (Preprint)

DTIC Science & Technology

2010-06-01

can be up 30 percent of the total weight of an aircraft engine [2], and may contain as many as 1 Copyright c⃝ 2010 by ASME 2000 individual airfoils ...reduction in stage count in a gas turbine engine and a de- crease in the part count of an individual airfoil row. The test data presented here provide...the vane row having modest design goals. So, while the Zweifel coefficient of the blade row was set Figure 1. Turbine design loop used to define the

Unsteady Aerodynamic Models for Turbomachinery Aeroelastic and Aeroacoustic Applications

NASA Technical Reports Server (NTRS)

Verdon, Joseph M.; Barnett, Mark; Ayer, Timothy C.

1995-01-01

Theoretical analyses and computer codes are being developed for predicting compressible unsteady inviscid and viscous flows through blade rows of axial-flow turbomachines. Such analyses are needed to determine the impact of unsteady flow phenomena on the structural durability and noise generation characteristics of the blading. The emphasis has been placed on developing analyses based on asymptotic representations of unsteady flow phenomena. Thus, high Reynolds number flows driven by small amplitude unsteady excitations have been considered. The resulting analyses should apply in many practical situations and lead to a better understanding of the relevant flow physics. In addition, they will be efficient computationally, and therefore, appropriate for use in aeroelastic and aeroacoustic design studies. Under the present effort, inviscid interaction and linearized inviscid unsteady flow models have been formulated, and inviscid and viscid prediction capabilities for subsonic steady and unsteady cascade flows have been developed. In this report, we describe the linearized inviscid unsteady analysis, LINFLO, the steady inviscid/viscid interaction analysis, SFLOW-IVI, and the unsteady viscous layer analysis, UNSVIS. These analyses are demonstrated via application to unsteady flows through compressor and turbine cascades that are excited by prescribed vortical and acoustic excitations and by prescribed blade vibrations. Recommendations are also given for the future research needed for extending and improving the foregoing asymptotic analyses, and to meet the goal of providing efficient inviscid/viscid interaction capabilities for subsonic and transonic unsteady cascade flows.

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.

Investigation of advanced counterrotation blade configuration concepts for high speed turboprop systems. Task 5: Unsteady counterrotation ducted propfan analysis

NASA Technical Reports Server (NTRS)

Hall, Edward J.; Delaney, Robert A.

1993-01-01

The primary objective of this study was the development of a time-marching three-dimensional Euler/Navier-Stokes aerodynamic analysis to predict steady and unsteady compressible transonic flows about ducted and unducted propfan propulsion systems employing multiple blade rows. The computer codes resulting from this study are referred to as ADPAC-AOAR\\CR (Advanced Ducted Propfan Analysis Codes-Angle of Attack Coupled Row). This document is the final report describing the theoretical basis and analytical results from the ADPAC-AOACR codes developed under task 5 of NASA Contract NAS3-25270, Unsteady Counterrotating Ducted Propfan Analysis. The ADPAC-AOACR Program is based on a flexible multiple blocked grid discretization scheme permitting coupled 2-D/3-D mesh block solutions with application to a wide variety of geometries. For convenience, several standard mesh block structures are described for turbomachinery applications. Aerodynamic calculations are based on a four-stage Runge-Kutta time-marching finite volume solution technique with added numerical dissipation. Steady flow predictions are accelerated by a multigrid procedure. Numerical calculations are compared with experimental data for several test cases to demonstrate the utility of this approach for predicting the aerodynamics of modern turbomachinery configurations employing multiple blade rows.

Investigation of advanced counterrotation blade configuration concepts for high speed turboprop systems. Task 5: Unsteady counterrotation ducted propfan analysis. Computer program user's manual

NASA Technical Reports Server (NTRS)

Hall, Edward J.; Delaney, Robert A.; Adamczyk, John J.; Miller, Christopher J.; Arnone, Andrea; Swanson, Charles

1993-01-01

The primary objective of this study was the development of a time-marching three-dimensional Euler/Navier-Stokes aerodynamic analysis to predict steady and unsteady compressible transonic flows about ducted and unducted propfan propulsion systems employing multiple blade rows. The computer codes resulting from this study are referred to as ADPAC-AOACR (Advanced Ducted Propfan Analysis Codes-Angle of Attack Coupled Row). This report is intended to serve as a computer program user's manual for the ADPAC-AOACR codes developed under Task 5 of NASA Contract NAS3-25270, Unsteady Counterrotating Ducted Propfan Analysis. The ADPAC-AOACR program is based on a flexible multiple blocked grid discretization scheme permitting coupled 2-D/3-D mesh block solutions with application to a wide variety of geometries. For convenience, several standard mesh block structures are described for turbomachinery applications. Aerodynamic calculations are based on a four-stage Runge-Kutta time-marching finite volume solution technique with added numerical dissipation. Steady flow predictions are accelerated by a multigrid procedure. Numerical calculations are compared with experimental data for several test cases to demonstrate the utility of this approach for predicting the aerodynamics of modern turbomachinery configurations employing multiple blade rows.

Linearized blade row compression component model. Stability and frequency response analysis of a J85-3 compressor

NASA Technical Reports Server (NTRS)

Tesch, W. A.; Moszee, R. H.; Steenken, W. G.

1976-01-01

NASA developed stability and frequency response analysis techniques were applied to a dynamic blade row compression component stability model to provide a more economic approach to surge line and frequency response determination than that provided by time-dependent methods. This blade row model was linearized and the Jacobian matrix was formed. The clean-inlet-flow stability characteristics of the compressors of two J85-13 engines were predicted by applying the alternate Routh-Hurwitz stability criterion to the Jacobian matrix. The predicted surge line agreed with the clean-inlet-flow surge line predicted by the time-dependent method to a high degree except for one engine at 94% corrected speed. No satisfactory explanation of this discrepancy was found. The frequency response of the linearized system was determined by evaluating its Laplace transfer function. The results of the linearized-frequency-response analysis agree with the time-dependent results when the time-dependent inlet total-pressure and exit-flow function amplitude boundary conditions are less than 1 percent and 3 percent, respectively. The stability analysis technique was extended to a two-sector parallel compressor model with and without interstage crossflow and predictions were carried out for total-pressure distortion extents of 180 deg, 90 deg, 60 deg, and 30 deg.

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 balde sections than with the DCA blade sections.

Theoretical Studies of Three Dimensional Transonic Flow through a Compressor Blade Row.

DTIC Science & Technology

1980-11-30

Row", Calspan Report No. AB-5487-A-l, AFOSR-TR-76- 1082 , AD-A031234, (August 1976). 2 Rae, W.J., "Relaxation Solutions for Three-Dimensional Transonic...S487-A-1, AFOSR-TR-76- 1082 , AD-A031234, (August 1976). 2. Rae, W.J., "Relaxation Solutions for Three-Dimensional Transonic Flow Through a Compressor

Geometry modeling and multi-block grid generation for turbomachinery configurations

NASA Technical Reports Server (NTRS)

Shih, Ming H.; Soni, Bharat K.

1992-01-01

An interactive 3D grid generation code, Turbomachinery Interactive Grid genERation (TIGER), was developed for general turbomachinery configurations. TIGER features the automatic generation of multi-block structured grids around multiple blade rows for either internal, external, or internal-external turbomachinery flow fields. Utilization of the Bezier's curves achieves a smooth grid and better orthogonality. TIGER generates the algebraic grid automatically based on geometric information provided by its built-in pseudo-AI algorithm. However, due to the large variation of turbomachinery configurations, this initial grid may not always be as good as desired. TIGER therefore provides graphical user interactions during the process which allow the user to design, modify, as well as manipulate the grid, including the capability of elliptic surface grid generation.

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.

Computer program for calculating full potential transonic, quasi-three-dimensional flow through a rotating turbomachinery blade row

NASA Technical Reports Server (NTRS)

Farrell, C. A.

1982-01-01

A fast, reliable computer code is described for calculating the flow field about a cascade of arbitrary two dimensional airfoils. The method approximates the three dimensional flow in a turbomachinery blade row by correcting for stream tube convergence and radius change in the throughflow direction. A fully conservative solution of the full potential equation is combined with the finite volume technique on a body-fitted periodic mesh, with an artificial density imposed in the transonic region to insure stability and the capture of shock waves. The instructions required to set up and use the code are included. The name of the code is QSONIC. A numerical example is also given to illustrate the output of the program.

Utilizing Direct Numerical Simulations of Transition and Turbulence in Design Optimization

NASA Technical Reports Server (NTRS)

Rai, Man M.

2015-01-01

Design optimization methods that use the Reynolds-averaged Navier-Stokes equations with the associated turbulence and transition models, or other model-based forms of the governing equations, may result in aerodynamic designs with actual performance levels that are noticeably different from the expected values because of the complexity of modeling turbulence/transition accurately in certain flows. Flow phenomena such as wake-blade interaction and trailing edge vortex shedding in turbines and compressors (examples of such flows) may require a computational approach that is free of transition/turbulence models, such as direct numerical simulations (DNS), for the underlying physics to be computed accurately. Here we explore the possibility of utilizing DNS data in designing a turbine blade section. The ultimate objective is to substantially reduce differences between predicted performance metrics and those obtained in reality. The redesign of a typical low-pressure turbine blade section with the goal of reducing total pressure loss in the row is provided as an example. The basic ideas presented here are of course just as applicable elsewhere in aerodynamic shape optimization as long as the computational costs are not excessive.

Experimental analysis of flow structure in contra-rotating axial flow pump designed with different rotational speed concept

NASA Astrophysics Data System (ADS)

Cao, Linlin; Watanabe, Satoshi; Imanishi, Toshiki; Yoshimura, Hiroaki; Furukawa, Akinori

2013-08-01

As a high specific speed pump, the contra-rotating axial flow pump distinguishes itself in a rear rotor rotating in the opposite direction of the front rotor, which remarkably contributes to the energy conversion, the reduction of the pump size, better hydraulic and cavitation performances. However, with two rotors rotating reversely, the significant interaction between blade rows was observed in our prototype contra-rotating rotors, which highly affected the pump performance compared with the conventional axial flow pumps. Consequently, a new type of rear rotor was designed by the rotational speed optimization methodology with some additional considerations, aiming at better cavitation performance, the reduction of blade rows interaction and the secondary flow suppression. The new rear rotor showed a satisfactory performance at the design flow rate but an unfavorable positive slope of the head — flow rate curve in the partial flow rate range less than 40% of the design flow rate, which should be avoided for the reliability of pump-pipe systems. In the present research, to understand the internal flow field of new rear rotor and its relation to the performances at the partial flow rates, the velocity distributions at the inlets and outlets of the rotors are firstly investigated. Then, the boundary layer flows on rotor surfaces, which clearly reflect the secondary flow inside the rotors, are analyzed through the limiting streamline observations using the multi-color oil-film method. Finally, the unsteady numerical simulations are carried out to understand the complicated internal flow structures in the rotors.

Gas Dynamic Modernization of Axial Uncooled Turbine by Means of CFD and Optimization Software

NASA Astrophysics Data System (ADS)

Marchukov, E. Yu; Egorov, I. N.

2018-01-01

The results of multicriteria optimization of three-stage low-pressure turbine are described in the paper. The aim of the optimization is to improve turbine operation process by three criteria: turbine outlet flow angle, value of residual swirl at the turbine outlet, and turbine efficiency. Full reprofiling of all blade rows is carried out while solving optimization problem. Reprofiling includes a change in both shape of flat blade sections (profiles) and three-dimensional shape of the blades. The study is carried out with 3D numerical models of turbines.

Broadband Noise of Fans - With Unsteady Coupling Theory to Account for Rotor and Stator Reflection/Transmission Effects

NASA Technical Reports Server (NTRS)

Hanson, Donald B.

2001-01-01

This report examines the effects on broadband noise generation of unsteady coupling between a rotor and stator in the fan stage of a turbofan engine. Whereas previous acoustic analyses treated the blade rows as isolated cascades, the present work accounts for reflection and transmission effects at both blade rows by tracking the mode and frequency scattering of pressure and vortical waves. The fan stage is modeled in rectilinear geometry to take advantage of a previously existing unsteady cascade theory for 3D perturbation waves and thereby use a realistic 3D turbulence spectrum. In the analysis, it was found that the set of participating modes divides itself naturally into "independent mode subsets" that couple only among themselves and not to the other such subsets. This principle is the basis for the analysis and considerably reduces computational effort. It also provides a simple, accurate scheme for modal averaging for further efficiency. Computed results for a coupled fan stage are compared with calculations for isolated blade rows. It is found that coupling increases downstream noise by 2 to 4 dB. Upstream noise is lower for isolated cascades and is further reduced by including coupling effects. In comparison with test data, the increase in the upstream/downstream differential indicates that broadband noise from turbulent inflow at the stator dominates downstream noise but is not a significant contributor to upstream noise.

Solution of plane cascade flow using improved surface singularity methods

NASA Technical Reports Server (NTRS)

Mcfarland, E. R.

1981-01-01

A solution method has been developed for calculating compressible inviscid flow through a linear cascade of arbitrary blade shapes. The method uses advanced surface singularity formulations which were adapted from those found in current external flow analyses. The resulting solution technique provides a fast flexible calculation for flows through turbomachinery blade rows. The solution method and some examples of the method's capabilities are presented.

Cooling arrangement for a tapered turbine blade

DOEpatents

Liang, George

2010-07-27

A cooling arrangement (11) for a highly tapered gas turbine blade (10). The cooling arrangement (11) includes a pair of parallel triple-pass serpentine cooling circuits (80,82) formed in an inner radial portion (50) of the blade, and a respective pair of single radial channel cooling circuits (84,86) formed in an outer radial portion (52) of the blade (10), with each single radial channel receiving the cooling fluid discharged from a respective one of the triple-pass serpentine cooling circuit. The cooling arrangement advantageously provides a higher degree of cooling to the most highly stressed radially inner portion of the blade, while providing a lower degree of cooling to the less highly stressed radially outer portion of the blade. The cooling arrangement can be implemented with known casting techniques, thereby facilitating its use on highly tapered, highly twisted Row 4 industrial gas turbine blades that could not be cooled with prior art cooling arrangements.

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.

Stator Indexing in Multistage Compressors

NASA Technical Reports Server (NTRS)

Barankiewicz, Wendy S.

1997-01-01

The relative circumferential location of stator rows (stator indexing) is an aspect of multistage compressor design that has not yet been explored for its potential impact on compressor aerodynamic performance. Although the inlet stages of multistage compressors usually have differing stator blade counts, the aft stages of core compressors can often have stage blocks with equal stator blade counts in successive stages. The potential impact of stator indexing is likely greatest in these stages. To assess the performance impact of stator indexing, researchers at the NASA Lewis Research Center used the 4 ft diameter, four-stage NASA Low Speed Axial Compressor for detailed experiments. This compressor has geometrically identical stages that can circumferentially index stator rows relative to each other in a controlled manner; thus it is an ideal test rig for such investigations.

Quasi-three-dimensional flow solution by meridional plane analysis

NASA Technical Reports Server (NTRS)

Katsanis, T.; Mcnally, W. D.

1974-01-01

A computer program has been developed to obtain subsonic or shockfree transonic, nonviscous flow analysis on the hub-shroud mid-channel flow surface of a turbomachine. The analysis may be for any annular passage, with or without blades. The blades may be fixed or rotating and may be twisted and leaned. The flow may be axial, radial or mixed. Blade surface velocities over the entire blade are approximated based on the rate of change of angular momentum. This gives a 3-D flow picture based on a 2-D analysis. The paper discusses the method used for the program and shows examples of the type of passages and blade rows which can be analyzed. Also, some numerical examples are given to show how the program can be used for practical assistance in design of blading, annular passages, and annular diffusers.

End-wall boundary layer measurements in a two-stage fan

NASA Technical Reports Server (NTRS)

Ball, C. L.; Reid, L.; Schmidt, J. F.

1983-01-01

Detailed flow measurements made in the casing boundary layer of a two-stage transonic fan are summarized. These measurements were taken at a station upstream of the fan, between all blade rows, and downstream of the last row. Conventional boundary layer parameters were calculated from the measured data. A classical two dimensional casing boundary layer was measured at the fan inlet and extended inward to approximately 15 percent of span. A highly three dimensional boundary layer was measured at the exit of each blade row and extended inward to approximately 10 percent of span. The steep radial gradient of axial velocity noted at the exit of the rotors was reduced substantially as the flow passed through the stators. This reduced gradient is attributed to flow mixing. The amount of flow mixing was reflected in the radial redistribution of total temperature as the flow passed through the stators. The blockage factors calculated from the measured data show an increase in blockage across the rotors and a decrease across the stators. For this fan the calculated blockages for the second stage were essentially the same as those for the first stage.

Solution of plane cascade flow using improved surface singularity methods. [application of panel method to internal aerodynamics

NASA Technical Reports Server (NTRS)

Mcfarland, E. R.

1981-01-01

A solution method was developed for calculating compressible inviscid flow through a linear cascade of arbitrary blade shapes. The method uses advanced surface singularity formulations which were adapted from those in current external flow analyses. The resulting solution technique provides a fast flexible calculation for flows through turbomachinery blade rows. The solution method and some examples of the method's capabilities are presented.

Modeling of Unsteady Three-dimensional Flows in Multistage Machines

NASA Technical Reports Server (NTRS)

Hall, Kenneth C.; Pratt, Edmund T., Jr.; Kurkov, Anatole (Technical Monitor)

2003-01-01

Despite many years of development, the accurate and reliable prediction of unsteady aerodynamic forces acting on turbomachinery blades remains less than satisfactory, especially when viewed next to the great success investigators have had in predicting steady flows. Hall and Silkowski (1997) have proposed that one of the main reasons for the discrepancy between theory and experiment and/or industrial experience is that many of the current unsteady aerodynamic theories model a single blade row in an infinitely long duct, ignoring potentially important multistage effects. However, unsteady flows are made up of acoustic, vortical, and entropic waves. These waves provide a mechanism for the rotors and stators of multistage machines to communicate with one another. In other words, wave behavior makes unsteady flows fundamentally a multistage (and three-dimensional) phenomenon. In this research program, we have has as goals (1) the development of computationally efficient computer models of the unsteady aerodynamic response of blade rows embedded in a multistage machine (these models will ultimately be capable of analyzing three-dimensional viscous transonic flows), and (2) the use of these computer codes to study a number of important multistage phenomena.

TIGGERC: Turbomachinery Interactive Grid Generator for 2-D Grid Applications and Users Guide

NASA Technical Reports Server (NTRS)

Miller, David P.

1994-01-01

A two-dimensional multi-block grid generator has been developed for a new design and analysis system for studying multiple blade-row turbomachinery problems. TIGGERC is a mouse driven, interactive grid generation program which can be used to modify boundary coordinates and grid packing and generates surface grids using a hyperbolic tangent or algebraic distribution of grid points on the block boundaries. The interior points of each block grid are distributed using a transfinite interpolation approach. TIGGERC can generate a blocked axisymmetric H-grid, C-grid, I-grid or O-grid for studying turbomachinery flow problems. TIGGERC was developed for operation on Silicon Graphics workstations. Detailed discussion of the grid generation methodology, menu options, operational features and sample grid geometries are presented.

TIGER: A user-friendly interactive grid generation system for complicated turbomachinery and axis-symmetric configurations

NASA Technical Reports Server (NTRS)

Shih, Ming H.; Soni, Bharat K.

1993-01-01

The issue of time efficiency in grid generation is addressed by developing a user friendly graphical interface for interactive/automatic construction of structured grids around complex turbomachinery/axis-symmetric configurations. The accuracy of geometry modeling and its fidelity is accomplished by adapting the nonuniform rational b-spline (NURBS) representation. A customized interactive grid generation code, TIGER, has been developed to facilitate the grid generation process for complicated internal, external, and internal-external turbomachinery fields simulations. The FORMS Library is utilized to build user-friendly graphical interface. The algorithm allows a user to redistribute grid points interactively on curves/surfaces using NURBS formulation with accurate geometric definition. TIGER's features include multiblock, multiduct/shroud, multiblade row, uneven blade count, and patched/overlapping block interfaces. It has been applied to generate grids for various complicated turbomachinery geometries, as well as rocket and missile configurations.

Computational Work to Support FAP/SRW Variable-Speed Power-Turbine Development

NASA Technical Reports Server (NTRS)

Ameri, Ali A.

2012-01-01

The purpose of this report is to document the work done to enable a NASA CFD code to model the transition on a blade. The purpose of the present work is to down-select a transition model that would allow the flow simulation of a Variable-Speed Power-Turbine (VSPT) to be accurately performed. The modeling is to be ultimately performed to also account for the blade row interactions and effect on transition and therefore accurate accounting for losses. The present work is limited to steady flows. The low Reynolds number k-omega model of Wilcox and a modified version of same will be used for modeling of transition on experimentally measured blade pressure and heat transfer. It will be shown that the k-omega model and its modified variant fail to simulate the transition with any degree of accuracy. A case is therefore made for more accurate transition models. Three-equation models based on the work of Mayle on Laminar Kinetic Energy were explored and the Walters and Leylek model which was thought to be in a more mature state of development is introduced and implemented in the Glenn-HT code. Two-dimensional flat plate results and three-dimensional results for flow over turbine blades and the resulting heat transfer and its transitional behavior are reported. It is shown that the transition simulation is much improved over the baseline k-omega model.

NASA Subsonic Jet Transport Noise Reduction Research

DTIC Science & Technology

2000-09-01

optical and acoustical interference. Figure 7 shows the concept and data from the installation of arrays of Herschel- Quincke tubes in the duct...tube row 16 tube row Herschel- Quincke Tube Tube length 12.5cm d = 3.8cm L = 9.2cm 2250 2350 2450 2500...Blade passage frequency, Hz R el at iv e p o w er , d B JT15D Turbofan Engine 4 d B Figure 7. Application of Herschel- Quincke tubes for

Navier-Stokes turbine heat transfer predictions using two-equation turbulence closures

NASA Technical Reports Server (NTRS)

Ameri, Ali A.; Arnone, Andrea

1992-01-01

Navier-Stokes calculations were carried out in order to predict the heat-transfer rates on turbine blades. The calculations were performed using TRAF2D which is a k-epsilon, explicit, finite volume mass-averaged Navier-Stokes solver. Turbulence was modeled using Coakley's q-omega and Chien's k-epsilon two-equation models and the Baldwin-Lomax algebraic model. The model equations along with the flow equations were solved explicitly on a nonperiodic C grid. Implicit residual smoothing (IRS) or a combination of multigrid technique and IRS was applied to enhance convergence rates. Calculations were performed to predict the Stanton number distributions on the first stage vane and blade row as well as the second stage vane row of the SSME high-pressure fuel turbine. The comparison serves to highlight the weaknesses of the turbulence models for use in turbomachinery heat-transfer calculations.

Noise of a model high speed counterrotation propeller at simulated takeoff/approach conditions (F7/A7)

NASA Technical Reports Server (NTRS)

Woodward, Richard P.

1987-01-01

A high speed advanced counterrotation propeller, was tested in the NASA-Lewis 9 x 15 foot Anechoic Wind Tunnel at simulated takeoff/approach conditions of 0.2 Mach number. Acoustic measurements were taken with fixed floor microphones, an axially translating microphone probe, and with a polar microphone probe which was fixed to the propeller nacelle and could take both sideline and circumferential acoustic surveys. Aerodynamic measurements were also made to establish the propeller operating conditions. The propeller was run over a range of blade setting angles from 36.4/36.5 to 41.1/39.4 deg, tip speeds from 165 to 259 m/sec, rotor spacings from 1.56 to 3.63 based on forward rotor tip chord to aerodynamic separation, and angles of attack to + or - 16 deg. First order rotor alone tones showed highest directivity levels near the propeller plane, while interaction tone showed high levels throughout sideline directivity, especially toward the propeller rotation axis. Interaction tone levels were sensitive to propeller row spacing while rotor alone tones showed little spacing effect. There is a decreased noise level associated with higher propeller blade numbers for the same overall propeller thrust.

Supersonic Stall Flutter of High Speed Fans. [in turbofan engines

NASA Technical Reports Server (NTRS)

Adamczyk, J. J.; Stevens, W.; Jutras, R.

1981-01-01

An analytical model is developed for predicting the onset of supersonic stall bending flutter in axial flow compressors. The analysis is based on a modified two dimensional, compressible, unsteady actuator disk theory. It is applied to a rotor blade row by considering a cascade of airfoils whose geometry and dynamic response coincide with those of a rotor blade element at 85 percent of the span height (measured from the hub). The rotor blades are assumed to be unshrouded (i.e., free standing) and to vibrate in their first flexural mode. The effects of shock waves and flow separation are included in the model through quasi-steady, empirical, rotor total-pressure-loss and deviation-angle correlations. The actuator disk model predicts the unsteady aerodynamic force acting on the cascade blading as a function of the steady flow field entering the cascade and the geometry and dynamic response of the cascade. Calculations show that the present model predicts the existence of a bending flutter mode at supersonic inlet Mach numbers. This flutter mode is suppressed by increasing the reduced frequency of the system or by reducing the steady state aerodynamic loading on the cascade. The validity of the model for predicting flutter is demonstrated by correlating the measured flutter boundary of a high speed fan stage with its predicted boundary. This correlation uses a level of damping for the blade row (i.e., the log decrement of the rotor system) that is estimated from the experimental flutter data. The predicted flutter boundary is shown to be in good agreement with the measured boundary.

Forced response unsteady aerodynamics in a multistage compressor

NASA Astrophysics Data System (ADS)

Capece, Vincent Ralph

The fundamental flow physics of the unsteady aerodynamics associated with forced vibrations in turbomachinery are investigated. Unique data are obtained through a series of experiments in a three stage axial flow research compressor which quantify the unsteady harmonic gust interaction phenomena over a range of operating and geometric conditions at high values of reduced frequency. In these experiments the effects of the following on the stator vane unsteady aerodynamics were quantified: (1) the steady aerodynamic loading, (2) the detailed waveform of the aerodynamic forcing function, including the chordwise and transverse gust components, (3) multistage blade row interactions, and (4) the solidity, ranging from a design value of 1.09 to an isolated airfoil. In addition, the effect of flow separation on the unsteady aerodynamics of an isolated airfoil was also investigated.

A Three-Dimensional Linearized Unsteady Euler Analysis for Turbomachinery Blade Rows

NASA Technical Reports Server (NTRS)

Montgomery, Matthew D.; Verdon, Joseph M.

1996-01-01

A three-dimensional, linearized, Euler analysis is being developed to provide an efficient unsteady aerodynamic analysis that can be used to predict the aeroelastic and aeroacoustic response characteristics of axial-flow turbomachinery blading. The field equations and boundary conditions needed to describe nonlinear and linearized inviscid unsteady flows through a blade row operating within a cylindrical annular duct are presented. In addition, a numerical model for linearized inviscid unsteady flow, which is based upon an existing nonlinear, implicit, wave-split, finite volume analysis, is described. These aerodynamic and numerical models have been implemented into an unsteady flow code, called LINFLUX. A preliminary version of the LINFLUX code is applied herein to selected, benchmark three-dimensional, subsonic, unsteady flows, to illustrate its current capabilities and to uncover existing problems and deficiencies. The numerical results indicate that good progress has been made toward developing a reliable and useful three-dimensional prediction capability. However, some problems, associated with the implementation of an unsteady displacement field and numerical errors near solid boundaries, still exist. Also, accurate far-field conditions must be incorporated into the FINFLUX analysis, so that this analysis can be applied to unsteady flows driven be external aerodynamic excitations.

A Three-Dimensional Linearized Unsteady Euler Analysis for Turbomachinery Blade Rows

NASA Technical Reports Server (NTRS)

Montgomery, Matthew D.; Verdon, Joseph M.

1997-01-01

A three-dimensional, linearized, Euler analysis is being developed to provide an efficient unsteady aerodynamic analysis that can be used to predict the aeroelastic and aeroacoustic responses of axial-flow turbo-machinery blading.The field equations and boundary conditions needed to describe nonlinear and linearized inviscid unsteady flows through a blade row operating within a cylindrical annular duct are presented. A numerical model for linearized inviscid unsteady flows, which couples a near-field, implicit, wave-split, finite volume analysis to a far-field eigenanalysis, is also described. The linearized aerodynamic and numerical models have been implemented into a three-dimensional linearized unsteady flow code, called LINFLUX. This code has been applied to selected, benchmark, unsteady, subsonic flows to establish its accuracy and to demonstrate its current capabilities. The unsteady flows considered, have been chosen to allow convenient comparisons between the LINFLUX results and those of well-known, two-dimensional, unsteady flow codes. Detailed numerical results for a helical fan and a three-dimensional version of the 10th Standard Cascade indicate that important progress has been made towards the development of a reliable and useful, three-dimensional, prediction capability that can be used in aeroelastic and aeroacoustic design studies.

Investigation of Performance of Axial-Flow Compressor of XT-46 Turbine-Propeller Engine. II - Performance of Revised Compressor at Design Equivalent Speed. II; Performance of Revised Compressor at Design Equivalent Speed

NASA Technical Reports Server (NTRS)

Creagh, John W. R.

1950-01-01

The compressor from the XT-46 turbine-propeller engine was revised by removing the last two rows of stator blades and by eliminating the interstage leakage paths described in a previous report. With the revised compressor, the flow choking point shifted upstream into the last rotor-blade row but the maximum weight flow was not increased over that of the original compressor. The flow range of the revised compressor was reduced to about two-thirds that obtained with the original compressor. The later stages of the compressor did not produce the design static-pressure increase probably because of excessive boundary-layer build-up in this region. Measurements obtained in the ninth-stage stator showed that the performance up to this station was promising but that the last three stages of the compressor were limiting the useful operating range of the preceding stages. Some modifications in flow-passage geometry and blade settings are believed to be necessary, however, before any major improvements in over-all compressor performance can be obtained.

Effect of Coolant Temperature and Mass Flow on Film Cooling of Turbine Blades

NASA Technical Reports Server (NTRS)

Garg, Vijay K.; Gaugler, Raymond E.

1997-01-01

A three-dimensional Navier Stokes code has been used to study the effect of coolant temperature, and coolant to mainstream mass flow ratio on the adiabatic effectiveness of a film-cooled turbine blade. The blade chosen is the VKI rotor with six rows of cooling holes including three rows on the shower head. The mainstream is akin to that under real engine conditions with stagnation temperature = 1900 K and stagnation pressure = 3 MPa. Generally, the adiabatic effectiveness is lower for a higher coolant temperature due to nonlinear effects via the compressibility of air. However, over the suction side of shower-head holes, the effectiveness is higher for a higher coolant temperature than that for a lower coolant temperature when the coolant to mainstream mass flow ratio is 5% or more. For a fixed coolant temperature, the effectiveness passes through a minima on the suction side of shower-head holes as the coolant to mainstream mass flow, ratio increases, while on the pressure side of shower-head holes, the effectiveness decreases with increase in coolant mass flow due to coolant jet lift-off. In all cases, the adiabatic effectiveness is highly three-dimensional.

Holographic studies of shock waves within transonic fan rotors

NASA Technical Reports Server (NTRS)

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

1974-01-01

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

Effect of external turbulence on the efficiency of film cooling with coolant injection into a transverse trench

NASA Astrophysics Data System (ADS)

Khalatov, A. A.; Panchenko, N. A.; Severin, S. D.

2017-09-01

Film cooling is among the basic methods used for thermal protection of blades in modern high-temperature gas turbines. Results of computer simulation of film cooling with coolant injection via a row of conventional inclined holes or a row of holes in a trench are presented in this paper. The ANSYS CFX 14 commercial software package was used for CFD-modeling. The effect is studied of the mainstream turbulence on the film cooling efficiency for the blowing ratio range between 0.6 and 2.3 and three different turbulence intensities of 1, 5, and 10%. The mainstream velocity was 150 and 400 m/s, while the temperatures of the mainstream and the injected coolant were 1100 and 500°C, respectively. It is demonstrated that, for the coolant injection via one row of trenched holes, an increase in the mainstream turbulence intensity reduces the film cooling efficiency in the entire investigated range of blowing ratios. It was revealed that freestream turbulence had varied effects on the film cooling efficiency depending on the blowing ratio and mainstream velocity in a blade channel. Thus, an increase in the mainstream turbulence intensity from 1 to 10% decreases the surface-averaged film cooling efficiency by 3-10% at a high mainstream velocity (400 m/s) in the blade channel and by 12-23% at a moderate velocity (of 150 m/s). Here, lower film cooling efficiencies correspond to higher blowing ratios. The effect of mainstream turbulence intensity on the film cooling efficiency decreases with increasing the mainstream velocity in the modeled channel for both investigated configurations.

Pump CFD code validation tests

NASA Technical Reports Server (NTRS)

Brozowski, L. A.

1993-01-01

Pump CFD code validation tests were accomplished by obtaining nonintrusive flow characteristic data at key locations in generic current liquid rocket engine turbopump configurations. Data were obtained with a laser two-focus (L2F) velocimeter at scaled design flow. Three components were surveyed: a 1970's-designed impeller, a 1990's-designed impeller, and a four-bladed unshrouded inducer. Two-dimensional velocities were measured upstream and downstream of the two impellers. Three-dimensional velocities were measured upstream, downstream, and within the blade row of the unshrouded inducer.

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.

Application of local indentations for film cooling of gas turbine blade leading edge

NASA Astrophysics Data System (ADS)

Petelchyts, V. Yu.; Khalatov, A. A.; Pysmennyi, D. N.; Dashevskyy, Yu. Ya.

2016-09-01

The paper presents results of computer simulation of the film cooling on the turbine blade leading edge model where the air coolant is supplied through radial holes and row of cylindrical inclined holes placed inside hemispherical dimples or trench. The blowing factor was varied from 0.5 to 2.0. The model size and key initial parameters for simulation were taken as for a real blade of a high-pressure high-performance gas turbine. Simulation was performed using commercial software code ANSYS CFX. The simulation results were compared with reference variant (no dimples or trench) both for the leading edge area and for the flat plate downstream of the leading edge.

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.

Ingestion resistant seal assembly

DOEpatents

Little, David A [Chuluota, FL

2011-12-13

A seal assembly limits gas leakage from a hot gas path to one or more disc cavities in a gas turbine engine. The seal assembly includes a seal apparatus associated with a blade structure including a row of airfoils. The seal apparatus includes an annular inner shroud associated with adjacent stationary components, a wing member, and a first wing flange. The wing member extends axially from the blade structure toward the annular inner shroud. The first wing flange extends radially outwardly from the wing member toward the annular inner shroud. A plurality of regions including one or more recirculation zones are defined between the blade structure and the annular inner shroud that recirculate working gas therein back toward the hot gas path.

Tip cap for a turbine rotor blade

DOEpatents

Kimmel, Keith D

2014-03-25

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

Acoustic Scattering by Three-Dimensional Stators and Rotors Using the SOURCE3D Code. Volume 1; Analysis and Results

NASA Technical Reports Server (NTRS)

Meyer, Harold D.

1999-01-01

This report provides a study of rotor and stator scattering using the SOURCE3D Rotor Wake/Stator Interaction Code. SOURCE3D is a quasi-three-dimensional computer program that uses three-dimensional acoustics and two-dimensional cascade load response theory to calculate rotor and stator modal reflection and transmission (scattering) coefficients. SOURCE3D is at the core of the TFaNS (Theoretical Fan Noise Design/Prediction System), developed for NASA, which provides complete fully coupled (inlet, rotor, stator, exit) noise solutions for turbofan engines. The reason for studying scattering is that we must first understand the behavior of the individual scattering coefficients provided by SOURCE3D, before eventually understanding the more complicated predictions from TFaNS. To study scattering, we have derived a large number of scattering curves for vane and blade rows. The curves are plots of output wave power divided by input wave power (in dB units) versus vane/blade ratio. Some of these plots are shown in this report. All of the plots are provided in a separate volume. To assist in understanding the plots, formulas have been derived for special vane/blade ratios for which wavefronts are either parallel or normal to rotor or stator chords. From the plots, we have found that, for the most part, there was strong transmission and weak reflection over most of the vane/blade ratio range for the stator. For the rotor, there was little transmission loss.

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.

Analysis of internal flow of J85-13 multistage compressor

NASA Technical Reports Server (NTRS)

Hager, R. D.

1977-01-01

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

Large-Scale Simulations and Detailed Flow Field Measurements for Turbomachinery Aeroacoustics

NASA Technical Reports Server (NTRS)

VanZante, Dale

2008-01-01

The presentation is a review of recent work in highly loaded compressors, turbine aeroacoustics and cooling fan noise. The specific topics are: the importance of correct numerical modeling to capture blade row interactions in the Ultra Efficient Engine Technology Proof-of-Concept Compressor, the attenuation of a detonation pressure wave by an aircraft axial turbine stage, current work on noise sources and acoustic attenuation in turbines, and technology development work on cooling fans for spaceflight applications. The topic areas were related to each other by certain themes such as the advantage of an experimentalist s viewpoint when analyzing numerical simulations and the need to improve analysis methods for very large numerical datasets.

Anisotropy of the Reynolds stress tensor in the wakes of wind turbine arrays in Cartesian arrangements with counter-rotating rotors

NASA Astrophysics Data System (ADS)

Hamilton, Nicholas; Cal, Raúl Bayoán

2015-01-01

A 4 × 3 wind turbine array in a Cartesian arrangement was constructed in a wind tunnel setting with four configurations based on the rotational sense of the rotor blades. The fourth row of devices is considered to be in the fully developed turbine canopy for a Cartesian arrangement. Measurements of the flow field were made with stereo particle-image velocimetry immediately upstream and downstream of the selected model turbines. Rotational sense of the turbine blades is evident in the mean spanwise velocity W and the Reynolds shear stress - v w ¯ . The flux of kinetic energy is shown to be of greater magnitude following turbines in arrays where direction of rotation of the blades varies. Invariants of the normalized Reynolds stress anisotropy tensor (η and ξ) are plotted in the Lumley triangle and indicate that distinct characters of turbulence exist in regions of the wake following the nacelle and the rotor blade tips. Eigendecomposition of the tensor yields principle components and corresponding coordinate system transformations. Characteristic spheroids representing the balance of components in the normalized anisotropy tensor are composed with the eigenvalues yielding shapes predicted by the Lumley triangle. Rotation of the coordinate system defined by the eigenvectors demonstrates trends in the streamwise coordinate following the rotors, especially trailing the top-tip of the rotor and below the hub. Direction of rotation of rotor blades is shown by the orientation of characteristic spheroids according to principle axes. In the inflows of exit row turbines, the normalized Reynolds stress anisotropy tensor shows cumulative effects of the upstream turbines, tending toward prolate shapes for uniform rotational sense, oblate spheroids for streamwise organization of rotational senses, and a mixture of characteristic shapes when the rotation varies by row. Comparison between the invariants of the Reynolds stress anisotropy tensor and terms from the mean mechanical energy equation indicate correlation between the degree of anisotropy and the regions of the wind turbine wakes where turbulence kinetic energy is produced. The flux of kinetic energy into the momentum-deficit area of the wake from above the canopy is associated with prolate characteristic spheroids. Flux upward into the wake from below the rotor area is associated with oblate characteristic spheroids. Turbulence in the region of the flow directly following the nacelle of the wind turbines demonstrates greater isotropy than regions following the rotor blades. The power and power coefficients for wind turbines indicate that flow structures on the order of magnitude of the spanwise turbine spacing that increase turbine efficiency depending on particular array configuration.

Application of holography to the determination of flow conditions within the rotating blade row of a compressor

NASA Technical Reports Server (NTRS)

Hantman, R. G.; Burr, R. J.; Alwang, W. G.; Williams, M. C.

1973-01-01

The double-pulse, double-exposure holography technique was applied to visualize the flow field within a transonic compressor rotor with a tip speed of 1800 ft/sec. The principal objective was to visualize the shock waves created in the flow field which was supersonic relative to the rotating blade row. The upstream rotor blade bow shocks and, at high speed, the outermost portion of the leading edge passage shock were successfully observed in the holograms. Techniques were devised for locating these shocks in three dimensions, and the results were compared with theoretical predictions. Density changes between the two pulses due to motion of the shocks were large and, therefore, it was not possible to resolve the fringe systems in detail for the 100% speed conditions. However, gross features of the shocks were easily observed, and the upstream shocks were well displayed. In all cases the shock angles were somewhat larger than predicted by theory, and a distinct increase in angle near the outer wall was observed, which may be attributed to endwall boundary layer effects. The location and orientation of the observed leading edge passage shocks were in good agreement with static pressure contours obtained from measurements in the outer casing over the rotor tip.

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.

Holographic studies of shock waves within transonic fan rotors

NASA Technical Reports Server (NTRS)

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

1973-01-01

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

Forced response analysis of an aerodynamically detuned supersonic turbomachine rotor

NASA Technical Reports Server (NTRS)

Hoyniak, D.; Fleeter, S.

1985-01-01

High performance aircraft-engine fan and compressor blades are vulnerable to aerodynamically forced vibrations generated by inlet flow distortions due to wakes from upstream blade and vane rows, atmospheric gusts, and maldistributions in inlet ducts. In this report, an analysis is developed to predict the flow-induced forced response of an aerodynamically detuned rotor operating in a supersonic flow with a subsonic axial component. The aerodynamic detuning is achieved by alternating the circumferential spacing of adjacent rotor blades. The total unsteady aerodynamic loading acting on the blading, as a result of the convection of the transverse gust past the airfoil cascade and the resulting motion of the cascade, is developed in terms of influence coefficients. This analysis is used to investigate the effect of aerodynamic detuning on the forced response of a 12-blade rotor, with Verdon's Cascade B flow geometry as a uniformly spaced baseline configuration. The results of this study indicate that, for forward traveling wave gust excitations, aerodynamic detuning is very beneficial, resulting in significantly decreased maximum-amplitude blade responses for many interblade phase angles.

Holographic testing of composite propfans for a cruise missile wind tunnel model

NASA Technical Reports Server (NTRS)

Miller, Christopher J.

1994-01-01

Each of the approximately 90 composite propfan blades constructed for a 55 percent scale cruise missile wind tunnel model were holographically tested to obtain natural frequencies and mode shapes. These data were used not only for quality assurance, but also to select sets of similar blades for each blade row. Presented along with the natural frequency data is a description of a computer-based image processing system developed to supplement the photographic based system for holographic image analysis and storage. The new system is quicker and cheaper, the holograms are indexed better, and several engineers can access the data simultaneously. The only negative effect is a slight reduction in image resolution, which does not influence the end use.

Subsonic Swept Fan Blade

NASA Technical Reports Server (NTRS)

Gallagher, Edward J. (Inventor); Rogers, Thomas H. (Inventor)

2017-01-01

A gas turbine engine includes a spool, a turbine coupled to drive the spool, a propulsor coupled to be driven at a at a design speed by the turbine through the spool, and a gear assembly coupled between the propulsor and the spool. Rotation of the turbine drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extend from the hub. Each of the propulsor blades includes an airfoil body. The leading edge of the airfoil body has a swept profile such that, at the design speed, a component of a relative velocity vector of a working gas that is normal to the leading edge is subsonic along the entire radial span.

Gas dynamic improvement of the axial compressor design for reduction of the flow non-uniformity level

NASA Astrophysics Data System (ADS)

Matveev, V. N.; Baturin, O. V.; Kolmakova, D. A.; Popov, G. M.

2017-01-01

Circumferential nonuniformity of gas flow is one of the main problems in the gas turbine engine. Usually, the flow circumferential nonuniformity appears near the annular frame located in the flow passage of the engine. The presence of circumferential nonuniformity leads to the increased dynamic stresses in the blade rows and the blade damage. The goal of this research was to find the ways of the flow non-uniformity reduction, which would not require a fundamental changing of the engine design. A new method for reducing the circumferential nonuniformity of the gas flow was proposed that allows the prediction of the pressure peak values of the rotor blades without computationally expensive CFD calculations.

Design and Analysis of Bionic Cutting Blades Using Finite Element Method.

PubMed

Li, Mo; Yang, Yuwang; Guo, Li; Chen, Donghui; Sun, Hongliang; Tong, Jin

2015-01-01

Praying mantis is one of the most efficient predators in insect world, which has a pair of powerful tools, two sharp and strong forelegs. Its femur and tibia are both armed with a double row of strong spines along their posterior edges which can firmly grasp the prey, when the femur and tibia fold on each other in capturing. These spines are so sharp that they can easily and quickly cut into the prey. The geometrical characteristic of the praying mantis's foreleg, especially its tibia, has important reference value for the design of agricultural soil-cutting tools. Learning from the profile and arrangement of these spines, cutting blades with tooth profile were designed in this work. Two different sizes of tooth structure and arrangement were utilized in the design on the cutting edge. A conventional smooth-edge blade was used to compare with the bionic serrate-edge blades. To compare the working efficiency of conventional blade and bionic blades, 3D finite element simulation analysis and experimental measurement were operated in present work. Both the simulation and experimental results indicated that the bionic serrate-edge blades showed better performance in cutting efficiency.

Design and Analysis of Bionic Cutting Blades Using Finite Element Method

PubMed Central

Li, Mo; Yang, Yuwang; Guo, Li; Chen, Donghui; Sun, Hongliang; Tong, Jin

2015-01-01

Praying mantis is one of the most efficient predators in insect world, which has a pair of powerful tools, two sharp and strong forelegs. Its femur and tibia are both armed with a double row of strong spines along their posterior edges which can firmly grasp the prey, when the femur and tibia fold on each other in capturing. These spines are so sharp that they can easily and quickly cut into the prey. The geometrical characteristic of the praying mantis's foreleg, especially its tibia, has important reference value for the design of agricultural soil-cutting tools. Learning from the profile and arrangement of these spines, cutting blades with tooth profile were designed in this work. Two different sizes of tooth structure and arrangement were utilized in the design on the cutting edge. A conventional smooth-edge blade was used to compare with the bionic serrate-edge blades. To compare the working efficiency of conventional blade and bionic blades, 3D finite element simulation analysis and experimental measurement were operated in present work. Both the simulation and experimental results indicated that the bionic serrate-edge blades showed better performance in cutting efficiency. PMID:27019583

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

NASA Technical Reports Server (NTRS)

Cheatham, J. G.

1974-01-01

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

Two- and three-dimensional turbine blade row flow field simulations

NASA Technical Reports Server (NTRS)

Buggeln, R. C.; Briley, W. R.; Mcdonald, H.; Shamroth, S. J.; Weinberg, B. C.

1987-01-01

Work performed in the numerical simulation of turbine passage flows via a Navier-Stokes approach is discussed. Both laminar and turbulent simulations in both two and three dimensions are discussed. An outline of the approach, background, and an overview of the results are given.

Research on the nonintrusive measurement of the turbine blade vibration

NASA Astrophysics Data System (ADS)

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

2008-11-01

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

Research on inverse methods and optimization in Italy

NASA Technical Reports Server (NTRS)

Larocca, Francesco

1991-01-01

The research activities in Italy on inverse design and optimization are reviewed. The review is focused on aerodynamic aspects in turbomachinery and wing section design. Inverse design of blade rows and ducts of turbomachinery in subsonic and transonic regime are illustrated by the Politecnico di Torino and turbomachinery industry (FIAT AVIO).

Experimental analysis of the flow in a two stage axial compressor at off-design conditions

NASA Astrophysics Data System (ADS)

Massardo, Aristide; Satta, Antonio

1987-05-01

The experimental analysis of the flow that develops in a two-stage axial flow compressor at off-design conditions is presented. The measurements are performed upstream, between, and downstream of the four blade rows of the compressor. The analysis shows the off-design effects on the local conditions of the flow field. Low-energy flow zones are identified, and the development of annulus-boundary-layer, secondary, and tip-clearance flows is shown. The tip-clearance flows are also present in the stator rows with various outlying conditions (stationary or rotating hub).

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.

Biomechanical factors contributing to self-organization in seagrass landscapes

USGS Publications Warehouse

Fonseca, M.S.; Koehl, M.A.R.; Kopp, B.S.

2007-01-01

Field observations have revealed that when water flow is consistently from one direction, seagrass shoots align in rows perpendicular to the primary axis of flow direction. In this study, live Zostera marina shoots were arranged either randomly or in rows perpendicular to the flow direction and tested in a seawater flume under unidirectional flow and waves to determine if shoot arrangement: a) influenced flow-induced force on individual shoots, b) differentially altered water flow through the canopy, and c) influenced light interception by the canopy. In addition, blade breaking strength was compared with flow-induced force to determine if changes in shoot arrangement might reduce the potential for damage to shoots. Under unidirectional flow, both current velocity in the canopy and force on shoots were significantly decreased when shoots were arranged in rows as compared to randomly. However, force on shoots was nearly constant with downstream distance, arising from the trade-off of shoot bending and in-canopy flow reduction. The coefficient of drag was higher for randomly-arranged shoots at low velocities (< 30 cm s- 1) but converged rapidly among the two shoot arrangements at higher velocities. Shoots arranged in rows tended to intercept slightly more light than those arranged randomly. Effects of shoot arrangement under waves were less clear, potentially because we did not achieve the proper plant size?row spacing ratio. At this point, we may only suggest that water motion, as opposed to light capture, is the dominant physical mechanism responsible for these shoot arrangements. Following a computation of the Environmental Stress Factor, we concluded that even photosynthetically active blades may be damaged or broken under frequently encountered storm conditions, irrespective of shoot arrangement. We hypothesize that when flow is generally from one direction, seagrass bed patterns over multiple scales of consideration may arise as a cumulative effect of individual shoot self-organization driven by reduced force and drag on the shoots and somewhat improved light capture.

Design and performance of a 427-meter-per-second-tip-speed two-stage fan having a 2.40 pressure ratio

NASA Technical Reports Server (NTRS)

Cunnan, W. S.; Stevans, W.; Urasek, D. C.

1978-01-01

The aerodynamic design and the overall and blade-element performances are presented of a 427-meter-per-second-tip-speed two-stage fan designed with axially spaced blade rows to reduce noise transmitted upstream of the fan. At design speed the highest recorded adiabatic efficiency was 0.796 at a pressure of 2.30. Peak efficiency was not established at design speed because of a damper failure which terminated testing prematurely. The overall efficiencies, at 60 and 80 percent of design speed, peaked at approximately 0.83.

FPCAS3D User's guide: A three dimensional full potential aeroelastic program, version 1

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.

1995-01-01

The FPCAS3D computer code has been developed for aeroelastic stability analysis of bladed disks such as those in fans, compressors, turbines, propellers, or propfans. The aerodynamic analysis used in this code is based on the unsteady three-dimensional full potential equation which is solved for a blade row. The structural analysis is based on a finite-element model for each blade. Detailed explanations of the aerodynamic analysis, the numerical algorithms, and the aeroelastic analysis are not given in this report. This guide can be used to assist in the preparation of the input data required by the FPCAS3D code. A complete description of the input data is provided in this report. In addition, six examples, including inputs and outputs, are provided.

Contribution of heat transfer to turbine blades and vanes for high temperature industrial gas turbines. Part 2: Heat transfer on serpentine flow passage.

PubMed

Takeishi, K; Aoki, S

2001-05-01

The improvement of the heat transfer coefficient of the 1st row blades in high temperature industrial gas turbines is one of the most important issues to ensure reliable performance of these components and to attain high thermal efficiency of the facility. This paper deals with the contribution of heat transfer to increase the turbine inlet temperature of such gas turbines in order to attain efficient and environmentally benign engines. Following the experiments described in Part 1, a set of trials was conducted to clarify the influence of the blade's rotating motion on the heat transfer coefficient for internal serpentine flow passages with turbulence promoters. Test results are shown and discussed in this second part of the contribution.

Fan Stagger Angle for Dirt Rejection

NASA Technical Reports Server (NTRS)

Gallagher, Edward J. (Inventor); Rose, Becky E. (Inventor); Brilliant, Lisa I. (Inventor)

2015-01-01

A gas turbine engine includes a spool, a turbine coupled to drive the spool, a propulsor coupled to be rotated about an axis by the turbine through the spool, and a gear assembly coupled between the propulsor and the spool such that rotation of the turbine drives the propulsor at a different speed than the spool. The propulsor includes a hub and a row of propulsor blades that extend from the hub. Each of the propulsor blades has a span between a root at the hub and a tip, and a chord between a leading edge and a trailing edge. The chord forms a stagger angle alpha with the axis, and the stagger angle alpha is less than 15 deg. at a position along the propulsor blade that is within an inboard 20% of the span.

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.

Task I: A Computational Model for Short Wavelength Stall Inception and Development In Multi-Stage Compressors

NASA Technical Reports Server (NTRS)

Suder, Kenneth (Technical Monitor); Tan, Choon-Sooi

2003-01-01

A computational model is presented for simulating axial compressor stall inception and development via disturbances with length scales on the order of several (typically about three) blade pitches. The model was designed for multi-stage compressors in which stall is initiated by these short wavelength disturbances, also referred to as spikes. The inception process described is fundamentally nonlinear, in contrast to the essentially linear behavior seen in so-called modal stall inception . The model was able to capture the following experimentally observed phenomena: (1) development of rotating stall via short wavelength disturbances, (2) formation and evolution of localized short wavelength stall cells in the first stage of a mismatched compressor, (3) the switch from long to short wavelength stall inception resulting from the re-staggering of the inlet guide vane, (4) the occurrence of rotating stall inception on the negatively sloped portion of the compressor characteristic. Parametric investigations indicated that (1) short wavelength disturbances were supported by the rotor blade row, (2) the disturbance strength was attenuated within the stators, and (3) the reduction of inter-blade row gaps can suppress the growth of short wavelength disturbances. It is argued that each local component group (rotor plus neighboring stators) has its own instability point (i.e. conditions at which disturbances are sustained) for short wavelength disturbances, with the instability point for the compressor set by the most unstable component group.

Measurement and Analysis of the Noise Radiated by Low Mach Number Centrifugal Blowers.

NASA Astrophysics Data System (ADS)

Yeager, David Marvin

An investigation was performed of the broad band, aerodynamically generated noise in low tip-speed Mach number, centrifugal air moving devices. An interdisciplinary experimental approach was taken which involved investigation of the aerodynamic and acoustic fields, and their mutual relationship. The noise generation process was studied using two experimental vehicles: (1) a scale model of a homologous family of centrifugal blowers typical of those used to cool computer and business equipment, and (2) a single blade from a centrifugal blower impeller placed in a known, controllable flow field. The radiation characteristics of the model blower were investigated by measuring the acoustic intensity distribution near the blower inlet and comparing it with the intensity near the inlet to an axial flow fan. Results showed that the centrifugal blower is a distributed, random noise source, unlike an axial fan which exhibited the effects of a coherent, interacting source distribution. Aerodynamic studies of the flow field in the inlet and at the discharge to the rotating impeller were used to assess the mean flow distribution through the impeller blade channels and to identify regions of excessive turbulence near the rotating blade row. Both circumferential and spanwise mean flow nonuniformities were identified along with a region of increased turbulence just downstream of the scroll cutoff. The fluid incidence angle, normally taken as an indicator of blower performance, was estimated from mean flow data as deviating considerably from an ideal impeller design. An investigation of the noise radiated from the single, isolated airfoil was performed using modern correlation and spectral analysis techniques. Radiation from the single blade in flow was characterized using newly developed expressions for the correlation area and the dipole source strength per unit area, and from the relationship between the blade surface pressure and the incident turbulent flow field. Results showed that radiation from the single blade was dominated by the effects of the incident turbulence. Normalized correlations areas of approximately 25% were measured at low frequencies. While the noise generation was more efficient at the trailing edge of the isolated blade, more noise was radiated from the region near the leading edge.

Heat Transfer on a Film-Cooled Blade - Effect of Hole Physics

NASA Technical Reports Server (NTRS)

Garg, Vijay K.; Rigby, David L.

1998-01-01

A multi-block, three-dimensional Navier-Stokes code has been used to study the within-hole and near-hole physics in relation to heat transfer on a film-cooled blade. The flow domain consists of the coolant flow through the plenum and hole-pipes for the three staggered rows of shower-head holes on the VK1 rotor, and the main flow over the blade. A multi-block grid is generated that is nearly orthogonal to the various surfaces. It may be noted that for the VK1 rotor the shower-head holes are inclined at 30 deg. to the spanwise direction, and are normal to the streamwise direction on the blade. Wilcox's k-omega turbulence model is used. The present study provides a much better comparison for the heat transfer coefficient at the blade mid-span with the experimental data than an earlier analysis wherein coolant velocity and temperature distributions were specified at the hole exits rather than extending the computational domain into the hole-pipe and plenum. Details of the distributions of coolant velocity, temperature, k and omega at the hole exits are also presented.

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.

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.

Component Performance Investigation of J71 Type II Turbines: III - Overall Performance of J71 Type IIA Turbine

NASA Technical Reports Server (NTRS)

Schum, Harold J.; Davison, Elmer H.; Petrash, Donald A.

1955-01-01

The over-all component performance characteristics of the J71 Type IIA three-stage turbine were experimentally determined over a range of speed and over-all turbine total-pressure ratio at inlet-air conditions af 35 inches of mercury absolute and 700 deg. R. The results are compared with those obtained for the J71 Type IIF turbine, which was previously investigated, the two turbines being designed for the same engine application. Geometrically the two turbines were much alike, having the same variation of annular flow area and the same number of blades for corresponding stator and rotor rows. However, the blade throat areas downstream of the first stator of the IIA turbine were smaller than those of the IIF; and the IIA blade profiles were curve-backed, whereas those of the IIF were straight-backed. The IIA turbine passed the equivalent design weight flow and had a brake internal efficiency of 0.880 at design equivalent speed and work output. A maximum efficiency of 0.896 occurred at 130 percent of design equivalent speed and a pressure ratio of 4.0. The turbine had a wide range of efficient operation. The IIA turbine had slightly higher efficiencies than the IIF turbine at comparable operating conditions. The fact that the IIA turbine obtained the design equivalent weight flow at the design equivalent operating point was probably a result of the decrease in the blading throat areas downstream of the first stator from those of the IIF turbine, which passed 105 percent of design weight flow at the corresponding operating point. The third stator row of blades of the IIA turbine choked at the design equivalent speed and at an over-all pressure ratio of 4.2; the third rotor choked at a pressure ratio of approximately 4.9

NASA contributions to radial turbine aerodynamic analyses

NASA Technical Reports Server (NTRS)

Glassman, A. J.

1980-01-01

A brief description of the radial turbine and its analysis needs is followed by discussions of five analytical areas; design geometry and performance, off design performance, blade row flow, scroll flow, and duct flow. The functions of the programs, areas of applicability, and limitations and uncertainties are emphasized. Both past contributions and current activities are discussed.

Vitamin D Supplementation for Prevention of Post-Traumatic Osteoarthritis: Evaluation in Animal and Clinical Models

DTIC Science & Technology

2017-10-01

June 16, 2000, Gotebörg, Sweden. (podium) 6. Wolf JM; Weiss APC; Akelman E: Mini-open carpal tunnel release using a new protective guide and blade ...Hartford, Connecticut. PERSONAL Married to Douglas S. Wolf 2 children Hobbies: rowing, hiking, running , travel, exploring restaurants Volunteer

Nereis alacranensis, a new species of polychaete (Annelida, Nereididae) from Alacranes Reef, southern Gulf of Mexico, with a key to Nereis from the Grand Caribbean.

PubMed

Ramírez Hernández, Adriana; Hernández-Alcántara, Pablo; Solís-Weiss, Vivianne

2015-09-02

A new species of polychaete, Nereis alacranensis n. sp., was found in dead coral rocks in the intertidal zone of Alacranes reef, southern Gulf of Mexico. N. alacranensis n. sp. can be included in a group of nereidids characterized by the absence of paragnaths in areas I and V of the pharynx, the presence of cones in a single row or absent in areas VII-VIII, and short blades in notopodial homogomph falcigers. The new species can be separated from the other species of the group by the presence of 3-7 cones in area VI and 7 cones arranged in a row in areas VII-VIII, finely dentate blades in notopodial homogomph falcigers, but most of all, by the presence of an unusual brown coarse arc shaped plate on the external ventral region of the peristomium. This structure has not yet been reported, at least in this genus. A taxonomic key of the species of Nereis recorded from the Grand Caribbean region is included.

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 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.

Development of an Aeroelastic Analysis Including a Viscous Flow Model

NASA Technical Reports Server (NTRS)

Keith, Theo G., Jr.; Bakhle, Milind A.

2001-01-01

Under this grant, Version 4 of the three-dimensional Navier-Stokes aeroelastic code (TURBO-AE) has been developed and verified. The TURBO-AE Version 4 aeroelastic code allows flutter calculations for a fan, compressor, or turbine blade row. This code models a vibrating three-dimensional bladed disk configuration and the associated unsteady flow (including shocks, and viscous effects) to calculate the aeroelastic instability using a work-per-cycle approach. Phase-lagged (time-shift) periodic boundary conditions are used to model the phase lag between adjacent vibrating blades. The direct-store approach is used for this purpose to reduce the computational domain to a single interblade passage. A disk storage option, implemented using direct access files, is available to reduce the large memory requirements of the direct-store approach. Other researchers have implemented 3D inlet/exit boundary conditions based on eigen-analysis. Appendix A: Aeroelastic calculations based on three-dimensional euler analysis. Appendix B: Unsteady aerodynamic modeling of blade vibration using the turbo-V3.1 code.

Incompressible lifting-surface aerodynamics for a rotor-stator combination

NASA Technical Reports Server (NTRS)

Ramachandra, S. M.

1984-01-01

Current literature on the three dimensional flow through compressor cascades deals with a row of rotor blades in isolation. Since the distance between the rotor and stator is usually 10 to 20 percent of the blade chord, the aerodynamic interference between them has to be considered for a proper evaluation of the aerothermodynamic performance of the stage. A unified approach to the aerodynamics of the incompressible flow through a stage is presented that uses the lifting surface theory for a compressor cascade of arbitrary camber and thickness distribution. The effects of rotor stator interference are represented as a linear function of the rotor and stator flows separately. The loading distribution on the rotor and stator flows separately. The loading distribution on the rotor and stator blades and the interference factor are determined concurrently through a matrix iteration process.

NWTC Aerodynamics Studies Improve Energy Capture and Lower Costs of Wind-Generated Electricity

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

2015-08-01

Researchers at the National Wind Technology Center (NWTC) at the National Renewable Energy Laboratory (NREL) have expanded wind turbine aerodynamic research from blade and rotor aerodynamics to wind plant and atmospheric inflow effects. The energy capture from wind plants is dependent on all of these aerodynamic interactions. Research at the NWTC is crucial to understanding how wind turbines function in large, multiple-row wind plants. These conditions impact the cumulative fatigue damage of turbine structural components that ultimately effect the useful lifetime of wind turbines. This work also is essential for understanding and maximizing turbine and wind plant energy production. Bothmore » turbine lifetime and wind plant energy production are key determinants of the cost of wind-generated electricity.« less

Effects of Shrouded Stator Cavity Flows on Multistage Axial Compressor Aerodynamic Performance

NASA Technical Reports Server (NTRS)

Wellborn, Steven R.; Okiishi, Theodore H.

1996-01-01

Experiments were performed on a low-speed multistage axial-flow compressor to assess the effects of shrouded stator cavity flows on aerodynamic performance. Five configurations, which involved changes in seal-tooth leakage rates and/or elimination of the shrouded stator cavities, were tested. Data collected enabled differences in overall individual stage and the third stage blade element performance parameters to be compared. The results show conclusively that seal-tooth leakage ran have a large impact on compressor aerodynamic performance while the presence of the shrouded stator cavities alone seemed to have little influence. Overall performance data revealed that for every 1% increase in the seal-tooth clearance to blade-height ratio the pressure rise dropped up to 3% while efficiency was reduced by 1 to 1.5 points. These observed efficiency penalty slopes are comparable to those commonly reported for rotor and cantilevered stator tip clearance variations. Therefore, it appears that in order to correctly predict overall performance it is equally important to account for the effects of seal-tooth leakage as it is to include the influence of tip clearance flows. Third stage blade element performance data suggested that the performance degradation observed when leakage was increased was brought about in two distinct ways. First, increasing seal-tooth leakage directly spoiled the near hub performance of the stator row in which leakage occurred. Second, the altered stator exit now conditions caused by increased leakage impaired the performance of the next downstream stage by decreasing the work input of the downstream rotor and increasing total pressure loss of the downstream stator. These trends caused downstream stages to progressively perform worse. Other measurements were acquired to determine spatial and temporal flow field variations within the up-and-downstream shrouded stator cavities. Flow within the cavities involved low momentum fluid traveling primarily in the circumferential direction at about 40% of the hub wheel speed. Measurements indicated that the flow within both cavities was much more complex than first envisioned. A vortical flow structure in the meridional plane, similar to a driven cavity, existed within the upstream cavity Furthermore, other spatial and temporal variations in Row properties existed. the most prominent being caused by the upstream potential influence of the downstream blade. This influence caused the fluid within cavities near the leading edges of either stator blades in space or rotor blades in time to be driven radially inward relative to fluid near blade mid-pitch. This influence also produced large unsteady velocity fluctuations in the downstream cavity because of the passing of the downstream rotor blade.

Investigation of Performance of Axial-Flow Compressor of XT-46 Turbine-Propeller Engine. I - Preliminary Investigation at 50-,70-, and 100-Percent Design Equivalent Speed

NASA Technical Reports Server (NTRS)

Creagh, John W.R.; Sandercrock, Donald M.

1950-01-01

An investigation is being conducted to determine the performance of the 12-stage axial-flow compressor of the XT-46 turbine-propeller engine. This compressor was designed to produce a pressure ratio of 9 at an adiabatic efficiency of 0.86. The design pressure ratios per stage were considerably greater than any employed in current aircraft gas-turbine engines using this type of compressor. The compressor performance was evaluated at two stations. The station near the entrance section of the combustors indicated a peak pressure ratio of 6.3 at an adiabatic efficiency of 0.63 for a corrected weight flow of 23.1 pounds per second. The other, located one blade-chord downstream of the last stator row, indicated a peak pressure ratio of 6.97 at an adiabatic efficiency of 0.81 for a corrected weight flow of 30.4 pounds per second. The difference in performance obtained at the two stations is attributed to shock waves in the vicinity of the last stator row. These shock waves and the accompanying flow choking, together with interstage circulatory flows, shift the compressor operating curves into the region where surge would normally occur. The inability of the compressor to meet design pressure ratio is probably due to boundary-layer buildup in the last stages, which cause axial velocities greater than design values that, in turn, adversely affect the angles of attack and turning angles in these blade rows.

Method of making an aero-derivative gas turbine engine

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

Wiebe, David J.

A method of making an aero-derivative gas turbine engine (100) is provided. A combustor outer casing (68) is removed from an existing aero gas turbine engine (60). An annular combustor (84) is removed from the existing aero gas turbine engine. A first row of turbine vanes (38) is removed from the existing aero gas turbine engine. A can annular combustor assembly (122) is installed within the existing aero gas turbine engine. The can annular combustor assembly is configured to accelerate and orient combustion gasses directly onto a first row of turbine blades of the existing aero gas turbine engine. Amore » can annular combustor assembly outer casing (108) is installed to produce the aero-derivative gas turbine engine (100). The can annular combustor assembly is installed within an axial span (85) of the existing aero gas turbine engine vacated by the annular combustor and the first row of turbine vanes.« less

Acoustic Reflection and Transmission of 2-Dimensional Rotors and Stators, Including Mode and Frequency Scattering Effects

NASA Technical Reports Server (NTRS)

Hanson, Donald B.

1999-01-01

A reduced order modeling scheme has been developed for the unsteady acoustic and vortical coupling between blade rows of a turbomachine. The essential behavior of the system is governed by modal scattering coefficients (i.e., reflection and transmission coefficients) of the rotor, stator, inlet and nozzle, which are calculated as if they were connected to non-reflecting ducts. The objective of this report is to identify fundamental behavior of these scattering coefficients for a better understanding of the role of blade row reflection and transmission in noise generation. A 2D flat plate unsteady cascade model is used for the analysis with the expectation that the general behavior presented herein will carry over to models that include more realistic flow and geometry. It is shown that stators scatter input waves into many modes at the same frequency whereas rotors scatter on frequency, or harmonic order. Important cases are shown here the rotor reflection coefficient is greater than unity; a mode at blade passing frequency (BPF) traveling from the stator with unit sound power is reflected by the rotor with more than unit power at 2xBPF and 3xBPE Analysis is presented to explain this unexpected phenomenon. Scattering curves are presented in a format chosen for design use and for physical interpretation. To aid in interpretation of the curves, formulas are derived for special condition where waveforms are parallel to perpendicular to the rotor.

Three-dimensional flow in radial turbomachinery and its impact on design

NASA Technical Reports Server (NTRS)

Tan, Choon S.; Hawthorne, William

1993-01-01

In the two papers on the 'Theory of Blade Design for Large Deflections' published in 1984, a new inverse design technique was presented for designing the shape of turbomachinery blades in three-dimensional flow. The technique involves the determination of the blade profile from the specification of a distribution of the product of the radius and the pitched averaged tangential velocity (i.e., r bar-V(sub theta), the mean swirl schedule) within the bladed region. This is in contrast to the conventional inverse design technique for turbomachinery blading in two dimensional flow in which the blade surface pressure or velocity distribution is specified and the blade profile determined as a result; this is feasible in two-dimensional flow because the streamlines along the blade surfaces are known a priori. However, in three-dimensional flow, the stream surface is free to deform within the blade passage so that the streamlines on the blade surfaces are not known a priori; thus it is difficult and not so useful to prescribe the blade surface pressure or velocity distribution and determine the resulting blade profile. It therefore seems logical to prescribe the swirl schedule within the bladed region for designing a turbomachinery blade profile in three-dimensional flow. Furthermore, specifying r bar-V(sub theta) has the following advantages: (1) it is related to the circulation around the blade (i.e., it is an aerodynamic quantity); (2) the work done or extracted is approximately proportional to the overall change in r bar-V(sub theta) across a given blade row (Euler turbine equation); and (3) the rate of change of r bar-V(sub theta) along the mean streamline at the blade is related to the pressure jump across the blade and therefore the blade loading. Since the publications of those two papers, the technique has been applied to the design of a low speed as well as a high speed radial inflow turbine (for turbocharger applications) both of which showed definite improvements in performance over that of wheels of conventional designs, the design study of a high pressure ratio radial inflow turbine with and without splitter blades.

Effect of Film-Hole Shape on Turbine Blade Film Cooling Performance

NASA Technical Reports Server (NTRS)

Han, J. C.; Teng, S.

2000-01-01

The detailed heat transfer coefficient and film cooling effectiveness distributions as well as tile detailed coolant jet temperature profiles on the suction side of a gas turbine blade A,ere measured using a transient liquid crystal image method and a traversing cold wire and a traversing 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 10(exp 5). Upstream unsteady wakes were simulated using a spoke-wheel type wake generator. The wake Strouhal number was kept at 0 or 0.1. 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.

Modeling Film-Coolant Flow Characteristics at the Exit of Shower-Head Holes

NASA Technical Reports Server (NTRS)

Garg, Vijay K.; Gaugler, R. E. (Technical Monitor)

2000-01-01

The coolant flow characteristics at the hole exits of a film-cooled blade are derived from an earlier analysis where the hole pipes and coolant plenum were also discretized. The blade chosen is the VKI rotor with three staggered rows of shower-head holes. The present analysis applies these flow characteristics at the shower-head hole exits. A multi-block three-dimensional Navier-Stokes code with Wilcox's k-omega model is used to compute the heat transfer coefficient on the film-cooled turbine blade. A reasonably good comparison with the experimental data as well as with the more complete earlier analysis where the hole pipes and coolant plenum were also gridded is obtained. If the 1/7th power law is assumed for the coolant flow characteristics at the hole exits, considerable differences in the heat transfer coefficient on the blade surface, specially in the leading-edge region, are observed even though the span-averaged values of h (heat transfer coefficient based on T(sub o)-T(sub w)) match well with the experimental data. This calls for span-resolved experimental data near film-cooling holes on a blade for better validation of the code.

Analyzing Aeroelasticity in Turbomachines

NASA Technical Reports Server (NTRS)

Reddy, T. S. R.; Srivastava, R.

2003-01-01

ASTROP2-LE is a computer program that predicts flutter and forced responses of blades, vanes, and other components of such turbomachines as fans, compressors, and turbines. ASTROP2-LE is based on the ASTROP2 program, developed previously for analysis of stability of turbomachinery components. In developing ASTROP2- LE, ASTROP2 was modified to include a capability for modeling forced responses. The program was also modified to add a capability for analysis of aeroelasticity with mistuning and unsteady aerodynamic solutions from another program, LINFLX2D, that solves the linearized Euler equations of unsteady two-dimensional flow. Using LINFLX2D to calculate unsteady aerodynamic loads, it is possible to analyze effects of transonic flow on flutter and forced response. ASTROP2-LE can be used to analyze subsonic, transonic, and supersonic aerodynamics and structural mistuning for rotors with blades of differing structural properties. It calculates the aerodynamic damping of a blade system operating in airflow so that stability can be assessed. The code also predicts the magnitudes and frequencies of the unsteady aerodynamic forces on the airfoils of a blade row from incoming wakes. This information can be used in high-cycle fatigue analysis to predict the fatigue lives of the blades.

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.

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.

Thermal shields for gas turbine rotor

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

Ross, Christopher W.; Acar, Bulent

A turbomachine including a rotor having an axis and a plurality of disks positioned adjacent to each other in the axial direction, each disk including opposing axially facing surfaces and a circumferentially extending radially facing surface located between the axially facing surfaces. At least one row of blades is positioned on each of the disks, and the blades include an airfoil extending radially outward from the disk A non-segmented circumferentially continuous ring structure includes an outer rim defining a thermal barrier extending axially in overlapping relation over a portion of the radially facing surface of at least one disk, andmore » extending to a location adjacent to a blade on the disk A compliant element is located between a radially inner circumferential portion of the ring structure and a flange structure that extends axially from an axially facing surface of the disk.« less

Response of a thin airfoil encountering strong density discontinuity

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

Marble, F.E.

1993-12-01

Airfoil theory for unsteady motion has been developed extensively assuming the undisturbed medium to be of uniform density, a restriction accurate for motion in the atmosphere. In some instances, notably for airfoil comprising fan, compressor and turbine blade rows, the undisturbed medium may carry density variations or ``spots``, resulting from non-uniformities in temperature or composition, of a size comparable to the blade chord. This condition exists for turbine blades, immediately downstream of the main burner of a gas turbine engine where the density fluctuations of the order of 50 percent may occur. Disturbances of a somewhat smaller magnitude arise frommore » the ingestion of hot boundary layers into fans, and exhaust into hovercraft. Because these regions of non-uniform density convect with the moving medium, the airfoil experiences a time varying load and moment which the authors calculate.« less

TIGER: Turbomachinery interactive grid generation

NASA Technical Reports Server (NTRS)

Soni, Bharat K.; Shih, Ming-Hsin; Janus, J. Mark

1992-01-01

A three dimensional, interactive grid generation code, TIGER, is being developed for analysis of flows around ducted or unducted propellers. TIGER is a customized grid generator that combines new technology with methods from general grid generation codes. The code generates multiple block, structured grids around multiple blade rows with a hub and shroud for either C grid or H grid topologies. The code is intended for use with a Euler/Navier-Stokes solver also being developed, but is general enough for use with other flow solvers. TIGER features a silicon graphics interactive graphics environment that displays a pop-up window, graphics window, and text window. The geometry is read as a discrete set of points with options for several industrial standard formats and NASA standard formats. Various splines are available for defining the surface geometries. Grid generation is done either interactively or through a batch mode operation using history files from a previously generated grid. The batch mode operation can be done either with a graphical display of the interactive session or with no graphics so that the code can be run on another computer system. Run time can be significantly reduced by running on a Cray-YMP.

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.

Achieving more efficient operation of the nozzle vane and rotor blade rows of gas turbines through using nonaxisymmetric end wall surfaces of interblade channels

NASA Astrophysics Data System (ADS)

Inozemtsev, A. A.; Samokhvalov, N. Yu.; Tikhonov, A. S.

2012-09-01

Results from a numerical study of three versions of the end-wall generatrix of the interblade channel used in the second-stage nozzle vanes of a prospective engine's turbine are presented. Recommendations for designing nonaxisymmetric end-wall surfaces are suggested based on the obtained data.

Survey of inlet noise reduction concepts for gas turbine engines

NASA Technical Reports Server (NTRS)

Lansing, D. L.; Chestnutt, D.

1976-01-01

This paper presents an overview of advanced concepts for the suppression of noise in the inlets of gas turbine engines. Noise suppression concepts are described, the directions of current research are reviewed, and problem areas requiring further work are indicated. The discussion focuses on acoustic liners, high Mach number inlets, active acoustic absorption, water vapor injection, and blade row reflection.

Open Rotor Aeroacoustic Installation Effects for Conventional and Unconventional Airframes

NASA Technical Reports Server (NTRS)

Czech, Michael J.; Thomas, Russell H.

2013-01-01

As extensive experimental campaign was performed to study the aeroacoustic installation effects of an open rotor with respect to both a conventional tube and wing type airframe and an unconventional hybrid wing body airframe. The open rotor rig had two counter rotating rows of blades each with eight blades of a design originally flight tested in the 1980s. The aeroacoustic installation effects measured in an aeroacoustic wind tunnel included those from flow effects due to inflow distortion or wake interaction and acoustic propagation effects such as shielding and reflection. The objective of the test campaign was to quantify the installation effects for a wide range of parameters and configurations derived from the two airframe types. For the conventional airframe, the open rotor was positioned in increments in front of and then over the main wing and then in positions representative of tail mounted aircraft with a conventional tail, a T-tail and a U-tail. The interaction of the wake of the open rotor as well as acoustic scattering results in an increase of about 10 dB when the rotor is positioned in front of the main wing. When positioned over the main wing a substantial amount of noise reduction is obtained and this is also observed for tail-mounted installations with a large U-tail. For the hybrid wing body airframe, the open rotor was positioned over the airframe along the centerline as well as off-center representing a twin engine location. A primary result was the documentation of the noise reduction from shielding as a function of the location of the open rotor upstream of the trailing edge of the hybrid wing body. The effects from vertical surfaces and elevon deflection were also measured. Acoustic lining was specially designed and inserted flush with the elevon and airframe surface, the result was an additional reduction in open rotor noise propagating to the far field microphones. Even with the older blade design used, the experiment provided quantification of the aeroacoustic installation effects for a wide range of open rotor and airframe configurations and can be used with data processing methods to evaluate the aeroacoustic installation effects for open rotors with modern blade designs.

Comparison of steady and unsteady secondary flows in a turbine stator cascade

NASA Technical Reports Server (NTRS)

Hebert, Gregory J.; Tiederman, William G.

1989-01-01

The effect of periodic rotor wakes on the secondary flow structure in a turbine stator cascade was investigated. A mechanism simulated the wakes shed from rotor blades by passing cylindrical rods across the inlet to a linear cascade installed in a recirculating water flow loop. Velocity measurements showed a passage vortex, similar to that seen in steady flow, during the time associated with undisturbed fluid. However, as the rotor wake passed through the blade row, a large crossflow toward the suction surface was observed in the midspan region. This caused the development of two large areas of circulation between the midspan and endwall regions, significantly distorting and weakening the passage vortices.

Multiple piece turbine blade

DOEpatents

Kimmel, Keith D [Jupiter, FL

2012-05-29

A turbine rotor blade with a spar and shell construction, the spar including an internal cooling supply channel extending from an inlet end on a root section and ending near the tip end, and a plurality of external cooling channels formed on both side of the spar, where a middle external cooling channel is connected to the internal cooling supply channels through a row of holes located at a middle section of the channels. The spar and the shell are held together by hooks that define serpentine flow passages for the cooling air and include an upper serpentine flow circuit and a lower serpentine flow circuit. the serpentine flow circuits all discharge into a leading edge passage or a trailing edge passage.

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.

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

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

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.

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.

Effect of RANS-Type Turbulence Models on Adiabatic Film Cooling Effectiveness over a Scaled Up Gas Turbine Blade Leading Edge Surface

NASA Astrophysics Data System (ADS)

Yepuri, Giridhara Babu; Talanki Puttarangasetty, Ashok Babu; Kolke, Deepak Kumar; Jesuraj, Felix

2016-06-01

Increasing the gas turbine inlet temperature is one of the key technologies in raising gas turbine engine power output. Film cooling is one of the efficient cooling techniques to cool the hot section components of a gas turbine engines in turn the turbine inlet temperature can be increased. This study aims at investigating the effect of RANS-type turbulence models on adiabatic film cooling effectiveness over a scaled up gas turbine blade leading edge surfaces. For the evaluation, five different two equation RANS-type turbulent models have been taken in consideration, which are available in the ANSYS-Fluent. For this analysis, the gas turbine blade leading edge configuration is generated using Solid Works. The meshing is done using ANSYS-Workbench Mesh and ANSYS-Fluent is used as a solver to solve the flow field. The considered gas turbine blade leading edge model is having five rows of film cooling circular holes, one at stagnation line and the two each on either side of stagnation line at 30° and 60° respectively. Each row has the five holes with the hole diameter of 4 mm, pitch of 21 mm arranged in staggered manner and has the hole injection angle of 30° in span wise direction. The experiments are carried in a subsonic cascade tunnel facility at heat transfer lab of CSIR-National Aerospace Laboratory with a Reynolds number of 1,00,000 based on leading edge diameter. From the Computational Fluid Dynamics (CFD) evaluation it is found that K-ɛ Realizable model gives more acceptable results with the experimental values, compared to the other considered turbulence models for this type of geometries. Further the CFD evaluated results, using K-ɛ Realizable model at different blowing ratios are compared with the experimental results.

A Model to Assess the Risk of Ice Accretion Due to Ice Crystal Ingestion in a Turbofan Engine and its Effects on Performance

NASA Technical Reports Server (NTRS)

Jorgenson, Philip C. E.; Veres, Joseph P.; Wright, William B.; Struk, Peter M.

2013-01-01

The occurrence of ice accretion within commercial high bypass aircraft turbine engines has been reported under certain atmospheric conditions. Engine anomalies have taken place at high altitudes that were attributed to ice crystal ingestion, partially melting, and ice accretion on the compression system components. The result was one or more of the following anomalies: degraded engine performance, engine roll back, compressor surge and stall, and flameout of the combustor. The main focus of this research is the development of a computational tool that can estimate whether there is a risk of ice accretion by tracking key parameters through the compression system blade rows at all engine operating points within the flight trajectory. The tool has an engine system thermodynamic cycle code, coupled with a compressor flow analysis code, and an ice particle melt code that has the capability of determining the rate of sublimation, melting, and evaporation through the compressor blade rows. Assumptions are made to predict the complex physics involved in engine icing. Specifically, the code does not directly estimate ice accretion and does not have models for particle breakup or erosion. Two key parameters have been suggested as conditions that must be met at the same location for ice accretion to occur: the local wet-bulb temperature to be near freezing or below and the local melt ratio must be above 10%. These parameters were deduced from analyzing laboratory icing test data and are the criteria used to predict the possibility of ice accretion within an engine including the specific blade row where it could occur. Once the possibility of accretion is determined from these parameters, the degree of blockage due to ice accretion on the local stator vane can be estimated from an empirical model of ice growth rate and time spent at that operating point in the flight trajectory. The computational tool can be used to assess specific turbine engines to their susceptibility to ice accretion in an ice crystal environment.

Computation of leading edge film cooling from a CONSOLE geometry (CONverging Slot hOLE)

NASA Astrophysics Data System (ADS)

Guelailia, A.; Khorsi, A.; Hamidou, M. K.

2016-01-01

The aim of this study is to investigate the effect of mass flow rate on film cooling effectiveness and heat transfer over a gas turbine rotor blade with three staggered rows of shower-head holes which are inclined at 30° to the spanwise direction, and are normal to the streamwise direction on the blade. To improve film cooling effectiveness, the standard cylindrical holes, located on the leading edge region, are replaced with the converging slot holes (console). The ANSYS CFX has been used for this computational simulation. The turbulence is approximated by a k-ɛ model. Detailed film effectiveness distributions are presented for different mass flow rate. The numerical results are compared with experimental data.

Aerodynamic performances of three fan stator designs operating with rotor having tip speed of 337 meters per second and pressure ratio of 1.54. 1: Experimental performance

NASA Technical Reports Server (NTRS)

Gelder, T. F.

1980-01-01

The aerodynamic performances of four stator-blade rows are presented and evaluated. The aerodynamic designs of two of these stators were compromised to reduce noise, a third design was not. On a calculated operating line passing through the design point pressure ratio, the best stator had overall pressure-ratio and efficiency decrements of 0.031 and 0.044, respectively, providing a stage pressure ratio of 1.483 and efficiency of 0.865. The other stators showed some correctable deficiencies due partly to the design compromises for noise. In the end-wall regions blade-element losses were significantly less for the shortest chord studied.

Reduction of Helicopter Blade-Vortex Interaction Noise by Active Rotor Control Technology

NASA Technical Reports Server (NTRS)

Yu, Yung H.; Gmelin, Bernd; Splettstoesser, Wolf; Brooks, Thomas F.; Philippe, Jean J.; Prieur, Jean

1997-01-01

Helicopter blade-vortex interaction noise is one of the most severe noise sources and is very important both in community annoyance and military detection. Research over the decades has substantially improved basic physical understanding of the mechanisms generating rotor blade-vortex interaction noise and also of controlling techniques, particularly using active rotor control technology. This paper reviews active rotor control techniques currently available for rotor blade vortex interaction noise reduction, including higher harmonic pitch control, individual blade control, and on-blade control technologies. Basic physical mechanisms of each active control technique are reviewed in terms of noise reduction mechanism and controlling aerodynamic or structural parameters of a blade. Active rotor control techniques using smart structures/materials are discussed, including distributed smart actuators to induce local torsional or flapping deformations, Published by Elsevier Science Ltd.

Investigation of Advanced Counterrotation Blade Configuration Concepts for High Speed Turboprop Systems. Task 3: Advanced Fan Section Grid Generator Final Report and Computer Program User's Manual

NASA Technical Reports Server (NTRS)

Crook, Andrew J.; Delaney, Robert A.

1991-01-01

A procedure is studied for generating three-dimensional grids for advanced turbofan engine fan section geometries. The procedure constructs a discrete mesh about engine sections containing the fan stage, an arbitrary number of axisymmetric radial flow splitters, a booster stage, and a bifurcated core/bypass flow duct with guide vanes. The mesh is an h-type grid system, the points being distributed with a transfinite interpolation scheme with axial and radial spacing being user specified. Elliptic smoothing of the grid in the meridional plane is a post-process option. The grid generation scheme is consistent with aerodynamic analyses utilizing the average-passage equation system developed by Dr. John Adamczyk of NASA Lewis. This flow solution scheme requires a series of blade specific grids each having a common axisymmetric mesh, but varying in the circumferential direction according to the geometry of the specific blade row.

The Supersonic Axial-Flow Compressor

NASA Technical Reports Server (NTRS)

Kantrowitz, Arthur

1950-01-01

An investigation has been made to explore the possibilities of axial-flow compressors operating with supersonic velocities into the blade rows. Preliminary calculations showed that very high pressure ratios across a stage, together with somewhat increased mass flows, were apparently possible with compressors which decelerated air through the speed of sound in their blading. The first phase of the investigation was the development of efficient supersonic diffusers to decelerate air through the speed of sound. The present report is largely a general discussion of some of the essential aerodynamics of single-stage supersonic axial-flow compressors. As an approach to the study of supersonic compressors, three possible velocity diagrams are discussed briefly. Because of the encouraging results of this study, an experimental single-stage supersonic compressor has been constructed and tested in Freon-12. In this compressor, air decelerates through the speed of sound in the rotor blading and enters the stators at subsonic speeds. A pressure ratio of about 1.8 at an efficiency of about 80 percent has been obtained.

Simulation of multistage turbine flows

NASA Technical Reports Server (NTRS)

Adamczyk, John J.; Mulac, Richard A.

1987-01-01

A flow model has been developed for analyzing multistage turbomachinery flows. This model, referred to as the average passage flow model, describes the time-averaged flow field with a typical passage of a blade row embedded within a multistage configuration. Computer resource requirements, supporting empirical modeling, formulation code development, and multitasking and storage are discussed. Illustrations from simulations of the space shuttle main engine (SSME) fuel turbine performed to date are given.

A Numerical Investigation of Turbine Noise Source Hierarchy and Its Acoustic Transmission Characteristics: Proof-of-Concept Progress

NASA Technical Reports Server (NTRS)

VanZante, Dale; Envia, Edmane

2008-01-01

A CFD-based simulation of single-stage turbine was done using the TURBO code to assess its viability for determining acoustic transmission through blade rows. Temporal and spectral analysis of the unsteady pressure data from the numerical simulations showed the allowable Tyler-Sofrin modes that are consistent with expectations. This indicated that high-fidelity acoustic transmission calculations are feasible with TURBO.

Parallel computation of three-dimensional aeroelastic fluid-structure interaction

NASA Astrophysics Data System (ADS)

Sadeghi, Mani

This dissertation presents a numerical method for the parallel computation of aeroelasticity (ParCAE). A flow solver is coupled to a structural solver by use of a fluid-structure interface method. The integration of the three-dimensional unsteady Navier-Stokes equations is performed in the time domain, simultaneously to the integration of a modal three-dimensional structural model. The flow solution is accelerated by using a multigrid method and a parallel multiblock approach. Fluid-structure coupling is achieved by subiteration. A grid-deformation algorithm is developed to interpolate the deformation of the structural boundaries onto the flow grid. The code is formulated to allow application to general, three-dimensional, complex configurations with multiple independent structures. Computational results are presented for various configurations, such as turbomachinery blade rows and aircraft wings. Investigations are performed on vortex-induced vibrations, effects of cascade mistuning on flutter, and cases of nonlinear cascade and wing flutter.

Transonic flow visualization using holographic interferometry

NASA Technical Reports Server (NTRS)

Bryanston-Cross, Peter J.

1987-01-01

An account is made of some of the applications of holographic interferometry to the visualization of transonic flows. In the case of the compressor shock visualization, the method is used regularly and has moved from being a research department invention to a design test tool. With the implementation of automatic processing and simple digitization systems, holographic vibrational analysis has also moved into routine nondestructive testing. The code verification interferograms were instructive, but the main turbomachinery interest is now in 3 dimensional flows. A major data interpretation effort will be required to compute tomographically the 3 dimensional flow around the leading or the trailing edges of a rotating blade row. The bolt on approach shows the potential application to current unsteady flows of interest. In particular that of the rotor passing and vortex interaction effects is experienced by the new generation of unducted fans. The turbocharger tests presents a new area for the application of holography.

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.

Rotor blade system with reduced blade-vortex interaction noise

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

Stagnation region gas film cooling for turbine blade leading edge applications

NASA Technical Reports Server (NTRS)

Luckey, D. W.; Winstanley, D. K.; Hanus, G. J.; Lecuyer, M. R.

1976-01-01

An experimental investigation was conducted to model the film-cooling performance for a turbine-vane leading edge using the stagnation region of a cylinder in cross flow. Experiments were conducted with a single row of spanwise-angled coolant holes for a range of the coolant blowing ratio with a freestream-to-wall temperature ratio of about 2.1 and a Reynolds number of 170,000, characteristic of the gas-turbine environment. Data from local heat-flux measurements are presented for coolant-hole injection angles of 25, 35, and 45 deg with the row of holes located at three positions relative to the stagnation line on the cylinder. Results show the spanwise (hole-to-hole) variation of heat-flux reduction due to film cooling and indicate conditions for the optimum film-cooling performance.

Effect of hole geometry and Electric-Discharge Machining (EDM) on airflow rates through small diameter holes in turbine blade material

NASA Technical Reports Server (NTRS)

Hippensteele, S. A.; Cochran, R. P.

1980-01-01

The effects of two design parameters, electrode diameter and hole angle, and two machine parameters, electrode current and current-on time, on air flow rates through small-diameter (0.257 to 0.462 mm) electric-discharge-machined holes were measured. The holes were machined individually in rows of 14 each through 1.6 mm thick IN-100 strips. The data showed linear increase in air flow rate with increases in electrode cross sectional area and current-on time and little change with changes in hole angle and electrode current. The average flow-rate deviation (from the mean flow rate for a given row) decreased linearly with electrode diameter and increased with hole angle. Burn time and finished hole diameter were also measured.

Efficient Fourier-based algorithms for time-periodic unsteady problems

NASA Astrophysics Data System (ADS)

Gopinath, Arathi Kamath

2007-12-01

This dissertation work proposes two algorithms for the simulation of time-periodic unsteady problems via the solution of Unsteady Reynolds-Averaged Navier-Stokes (URANS) equations. These algorithms use a Fourier representation in time and hence solve for the periodic state directly without resolving transients (which consume most of the resources in a time-accurate scheme). In contrast to conventional Fourier-based techniques which solve the governing equations in frequency space, the new algorithms perform all the calculations in the time domain, and hence require minimal modifications to an existing solver. The complete space-time solution is obtained by iterating in a fifth pseudo-time dimension. Various time-periodic problems such as helicopter rotors, wind turbines, turbomachinery and flapping-wings can be simulated using the Time Spectral method. The algorithm is first validated using pitching airfoil/wing test cases. The method is further extended to turbomachinery problems, and computational results verified by comparison with a time-accurate calculation. The technique can be very memory intensive for large problems, since the solution is computed (and hence stored) simultaneously at all time levels. Often, the blade counts of a turbomachine are rescaled such that a periodic fraction of the annulus can be solved. This approximation enables the solution to be obtained at a fraction of the cost of a full-scale time-accurate solution. For a viscous computation over a three-dimensional single-stage rescaled compressor, an order of magnitude savings is achieved. The second algorithm, the reduced-order Harmonic Balance method is applicable only to turbomachinery flows, and offers even larger computational savings than the Time Spectral method. It simulates the true geometry of the turbomachine using only one blade passage per blade row as the computational domain. In each blade row of the turbomachine, only the dominant frequencies are resolved, namely, combinations of neighbor's blade passing. An appropriate set of frequencies can be chosen by the analyst/designer based on a trade-off between accuracy and computational resources available. A cost comparison with a time-accurate computation for an Euler calculation on a two-dimensional multi-stage compressor obtained an order of magnitude savings, and a RANS calculation on a three-dimensional single-stage compressor achieved two orders of magnitude savings, with comparable accuracy.

Unobstructed particle image velocimetry measurements within an axial turbo-pump using liquid and blades with matched refractive indices

NASA Astrophysics Data System (ADS)

Uzol, O.; Chow, Y.-C.; Katz, J.; Meneveau, C.

2002-08-01

Performing PIV measurements within complex turbomachinery with multiple blade rows is difficult due to the optical obstruction to the illuminating sheet and to the camera caused by the blades. This paper introduces a refractive index matched facility that overcomes this problem. The rotor and stator blades are made of transparent acrylic, and the working fluid has the same optical refractive index as the blades. A 64% by weight solution of sodium iodide in water is used for this purpose. This liquid has a kinematic viscosity of about 1.1×10-6 m2/s, which is almost the same as that of water enabling operation at high Reynolds numbers. Issues related to operating with this fluid such as chemical stability, variations in transmittance and solutions to these problems are discussed. This setup allows full optical access to the entire rotor and stator passages both to the laser sheet and the camera. The experiments are conducted at different streamwise locations covering the entire flow fields around the rotor, the stator, the gap between them, and the wakes behind. Vector maps of the instantaneous and phase-averaged flow fields as well as the distribution of turbulent kinetic energy are obtained. Measurements at different magnifications enable us to obtain an overview of the flow structure, as well as detailed velocity distributions in the boundary layers and in the wakes.

Solution of the Average-Passage Equations for the Incompressible Flow through Multiple-Blade-Row Turbomachinery

DTIC Science & Technology

1994-02-01

numerical treatment. An explicit numerical procedure based on Runqe-Kutta time stepping for cell-centered, hexahedral finite volumes is...An explicit numerical procedure based on Runge-Kutta time stepping for cell-centered, hexahedral finite volumes is outlined for the approximate...Discretization 16 3.1 Cell-Centered Finite -Volume Discretization in Space 16 3.2 Artificial Dissipation 17 3.3 Time Integration 21 3.4 Convergence

Fundamental Characterization of Spanwise Loading and Trailed Wake Vortices

DTIC Science & Technology

2016-07-01

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

Mach number scaling of helicopter rotor blade/vortex interaction noise

NASA Technical Reports Server (NTRS)

Leighton, Kenneth P.; Harris, Wesley L.

1985-01-01

A parametric study of model helicopter rotor blade slap due to blade vortex interaction (BVI) was conducted in a 5 by 7.5-foot anechoic wind tunnel using model helicopter rotors with two, three, and four blades. The results were compared with a previously developed Mach number scaling theory. Three- and four-bladed rotor configurations were found to show very good agreement with the Mach number to the sixth power law for all conditions tested. A reduction of conditions for which BVI blade slap is detected was observed for three-bladed rotors when compared to the two-bladed baseline. The advance ratio boundaries of the four-bladed rotor exhibited an angular dependence not present for the two-bladed configuration. The upper limits for the advance ratio boundaries of the four-bladed rotors increased with increasing rotational speed.

Ducted-Fan Engine Acoustic Predictions using a Navier-Stokes Code

NASA Technical Reports Server (NTRS)

Rumsey, C. L.; Biedron, R. T.; Farassat, F.; Spence, P. L.

1998-01-01

A Navier-Stokes computer code is used to predict one of the ducted-fan engine acoustic modes that results from rotor-wake/stator-blade interaction. A patched sliding-zone interface is employed to pass information between the moving rotor row and the stationary stator row. The code produces averaged aerodynamic results downstream of the rotor that agree well with a widely used average-passage code. The acoustic mode of interest is generated successfully by the code and is propagated well upstream of the rotor; temporal and spatial numerical resolution are fine enough such that attenuation of the signal is small. Two acoustic codes are used to find the far-field noise. Near-field propagation is computed by using Eversman's wave envelope code, which is based on a finite-element model. Propagation to the far field is accomplished by using the Kirchhoff formula for moving surfaces with the results of the wave envelope code as input data. Comparison of measured and computed far-field noise levels show fair agreement in the range of directivity angles where the peak radiation lobes from the inlet are observed. Although only a single acoustic mode is targeted in this study, the main conclusion is a proof-of-concept: Navier-Stokes codes can be used both to generate and propagate rotor/stator acoustic modes forward through an engine, where the results can be coupled to other far-field noise prediction codes.

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

NASA Technical Reports Server (NTRS)

Davino, R.; Lakshminarayana, B.

1982-01-01

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

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.

MPT Prediction of Aircraft-Engine Fan Noise

NASA Technical Reports Server (NTRS)

Connell, Stuart D.

2004-01-01

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

Advancing-side directivity and retreating-side interactions of model rotor blade-vortex interaction noise

NASA Technical Reports Server (NTRS)

Martin, R. M.; Splettstoesser, W. R.; Elliott, J. W.; Schultz, K.-J.

1988-01-01

Acoustic data are presented from a 40 percent scale model of the four-bladed BO-105 helicopter main rotor, tested in a large aerodynamic wind tunnel. Rotor blade-vortex interaction (BVI) noise data in the low-speed flight range were acquired using a traversing in-flow microphone array. Acoustic results presented are used to assess the acoustic far field of BVI noise, to map the directivity and temporal characteristics of BVI impulsive noise, and to show the existence of retreating-side BVI signals. The characterics of the acoustic radiation patterns, which can often be strongly focused, are found to be very dependent on rotor operating condition. The acoustic signals exhibit multiple blade-vortex interactions per blade with broad impulsive content at lower speeds, while at higher speeds, they exhibit fewer interactions per blade, with much sharper, higher amplitude acoustic signals. Moderate-amplitude BVI acoustic signals measured under the aft retreating quadrant of the rotor are shown to originate from the retreating side of the rotor.

High fidelity simulation of non-synchronous vibration for aircraft engine fan/compressor

NASA Astrophysics Data System (ADS)

Im, Hong-Sik

The objectives of this research are to develop a high fidelity simulation methodology for turbomachinery aeromechanical problems and to investigate the mechanism of non-synchronous vibration (NSV) of an aircraft engine axial compressor. A fully conservative rotor/stator sliding technique is developed to accurately capture the unsteadiness and interaction between adjacent blade rows. Phase lag boundary conditions (BC) based on the time shift (direct store) method and the Fourier series phase lag BC are implemented to take into account the effect of phase difference for a sector of annulus simulation. To resolve the nonlinear interaction between flow and vibrating blade structure, a fully coupled fluid-structure interaction (FSI) procedure that solves the structural modal equations and time accurate Navier-Stokes equations simultaneously is adopted. An advanced mesh deformation method that generates the blade tip block mesh moving with the blade displacement is developed to ensure the mesh quality. An efficient and low diffusion E-CUSP (LDE) scheme as a Riemann solver designed to minimize numerical dissipation is used with an improved hybrid RANS/LES turbulence strategy, delayed detached eddy simulation (DDES). High order accuracy (3rd and 5th order) weighted essentially non-oscillatory (WENO) schemes for inviscid flux and a conservative 2nd and 4th order viscous flux differencing are employed. Extensive validations are conducted to demonstrate high accuracy and robustness of the high fidelity FSI simulation methodology. The validated cases include: (1) DDES of NACA 0012 airfoil at high angle of attack with massive separation. The DDES accurately predicts the drag whereas the URANS model significantly over predicts the drag. (2) The AGARD Wing 445.6 flutter boundary is accurately predicted including the point at supersonic incoming flow. (3) NASA Rotor 67 validation for steady state speed line and radial profiles at peak efficiency point and near stall point. The calculated results agree excellently with the experiment. (4) NASA Stage 35 speed line and radial profiles to validate the steady state mixing plane BC for multistage computation. Excellent agreement is obtained between the computation and experiment. (5) NASA Rotor 67 full annulus and single passage FSI simulation at near peak condition to validate phase lag BC. The time shifted phase lag BC accurately predicts blade vibration responses that agrees better with the full annulus FSI simulation. The DDES methodology is used to investigate the stall inception of NASA Rotor 67. The stall process begins with spike inception and develops to full stall. The whole process is simulated with full annulus of the rotor. The fully coupled FSI is then used to simulate the stall flutter of NASA Rotor 67. The multistage simulations of a GE aircraft engine high pressure compressor (HPC) reveal for the first time that the travelling tornado vortex formed on the rotor blade tip region is the root cause for the NSV of the compressor. The rotor blades under NSV have large torsional vibration due to the tornado vortex propagation in the opposite to the rotor rotation. The tornado vortex frequency passing the suction surface of each blade in the tip region agrees with the NSV frequency. The predicted NSV frequency based on URANS model with rigid blades agrees very well with the experimental measurement with only 3.3% under-predicted. The NSV prediction using FSI with vibrating blades also obtain the same frequency as the rigid blades. This is because that the NSV is primarily caused by the flow vortex instability and the no resonance occurs. The blade structures respond passively and the small amplitudes of the blade vibration do not have significant effect on the flow. The predicted frequency using DDES with rigid blades is more deviated from the experiment and is 14.7% lower. The reason is that the DDES tends to predict the rotor stall earlier than the URANS and the NSV can be achieved only at higher mass flow rate, which generates a lower frequency. The possible reason for the DDES to predict the rotor stall early may be because DDES is more sensitive to wave reflection and a non-reflective boundary condition may be necessary. Overall, the high fidelity FSI methodology developed in this thesis for aircraft engine fan/compressor aeromechanics simulation is demonstrated to be very successful and has advanced the forefront of the state of the art. Future work to continue to improve the accuracy and efficiency is discussed at the end of the thesis.

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

NASA Technical Reports Server (NTRS)

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

1983-01-01

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

Computational Aerodynamic Simulations of an 840 ft/sec Tip Speed Advanced Ducted Propulsor Fan System Model for Acoustic Methods Assessment and Development

NASA Technical Reports Server (NTRS)

Tweedt, Daniel L.

2014-01-01

Computational Aerodynamic simulations of an 840 ft/sec tip speed, Advanced Ducted Propulsor fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, lownoise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15- foot Low Speed Wind Tunnel at the NASA Glenn Research Center, resulting in quality, detailed aerodynamic and acoustic measurement data. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating conditions simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, excluding a long core duct section downstream of the core inlet guide vane. As a result, only fan rotational speed and system bypass ratio, set by specifying static pressure downstream of the core inlet guide vane row, were adjusted in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. The computed blade row flow fields for all five fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow fields reveals no excessive boundary layer separations or related secondary-flow problems. A few spanwise comparisons between computational and measurement data in the bypass duct show that they are in good agreement, thus providing a partial validation of the computational results.

Gas turbine sealing apparatus

DOEpatents

Marra, John Joseph; Wessell, Brian J.; Liang, George

2013-03-05

A sealing apparatus in a gas turbine. The sealing apparatus includes a seal housing apparatus coupled to a disc/rotor assembly so as to be rotatable therewith during operation of the gas turbine. The seal housing apparatus comprises a base member, a first leg portion, a second leg portion, and spanning structure. The base member extends generally axially between forward and aft rows of rotatable blades and is positioned adjacent to a row of stationary vanes. The first leg portion extends radially inwardly from the base member and is coupled to the disc/rotor assembly. The second leg portion is axially spaced from the first leg portion, extends radially inwardly from the base member, and is coupled to the disc/rotor assembly. The spanning structure extends between and is rigidly coupled to each of the base member, the first leg portion, and the second leg portion.

Studies of blade-vortex interaction noise reduction by rotor blade modification

NASA Technical Reports Server (NTRS)

Brooks, Thomas F.

1993-01-01

Blade-vortex interaction (BVI) noise is one of the most objectionable types of helicopter noise. This impulsive blade-slap noise can be particularly intense during low-speed landing approach and maneuvers. Over the years, a number of flight and model rotor tests have examined blade tip modification and other blade design changes to reduce this noise. Many times these tests have produced conflicting results. In the present paper, a number of these studies are reviewed in light of the current understanding of the BVI noise problem. Results from one study in particular are used to help establish the noise reduction potential and to shed light on the role of blade design. Current blade studies and some new concepts under development are also described.

Computational Assessment of the Aerodynamic Performance of a Variable-Speed Power Turbine for Large Civil Tilt-Rotor Application

NASA Technical Reports Server (NTRS)

Welch, Gerard E.

2011-01-01

The main rotors of the NASA Large Civil Tilt-Rotor notional vehicle operate over a wide speed-range, from 100% at take-off to 54% at cruise. The variable-speed power turbine offers one approach by which to effect this speed variation. Key aero-challenges include high work factors at cruise and wide (40 to 60 deg.) incidence variations in blade and vane rows over the speed range. The turbine design approach must optimize cruise efficiency and minimize off-design penalties at take-off. The accuracy of the off-design incidence loss model is therefore critical to the turbine design. In this effort, 3-D computational analyses are used to assess the variation of turbine efficiency with speed change. The conceptual design of a 4-stage variable-speed power turbine for the Large Civil Tilt-Rotor application is first established at the meanline level. The design of 2-D airfoil sections and resulting 3-D blade and vane rows is documented. Three-dimensional Reynolds Averaged Navier-Stokes computations are used to assess the design and off-design performance of an embedded 1.5-stage portion-Rotor 1, Stator 2, and Rotor 2-of the turbine. The 3-D computational results yield the same efficiency versus speed trends predicted by meanline analyses, supporting the design choice to execute the turbine design at the cruise operating speed.

High Fidelity Aeroelasticity Simulations of Aircraft and Turbomachinery with Fully-Coupled Fluid-Structure Interaction

NASA Astrophysics Data System (ADS)

Gan, Jiaye

The purpose of this research is to develop high fidelity numerical methods to investigate the complex aeroelasticity fluid-structural problems of aircraft and aircraft engine turbomachinery. Unsteady 3D compressible Navier-Stokes equations in generalized coordinates are solved to simulate the complex fluid dynamic problems in aeroelasticity. An efficient and low diffusion E-CUSP (LDE) scheme designed to minimize numerical dissipation is used as a Riemann solver to capture shock waves in transonic and supersonic flows. An improved hybrid turbulence modeling, delayed detached eddy simulation (DDES), is implemented to simulate shock induced separation and rotating stall flows. High order accuracy (3rd and 5th order) weighted essentially non-oscillatory (WENO) schemes for inviscid flux and a conservative 2nd and 4th order viscous flux differencing are employed. To resolve the nonlinear interaction between flow and vibrating blade structures, a fully coupled fluid-structure interaction (FSI) procedure that solves the structural modal equations and time accurate Navier-Stokes equations simultaneously is adopted. A rotor/stator sliding interpolation technique is developed to accurately capture the blade rows interaction at the interface with general grid distribution. Phase lag boundary conditions (BC) based on the time shift (direct store) method and the Fourier series phase lag BC are applied to consider the effect of phase difference for a sector of annulus simulation. Extensive validations are conducted to demonstrate high accuracy and robustness of the high fidelity FSI methodology. The accuracy and robustness of RANS, URANS and DDES turbulence models with high order schemes for predicting the lift and drag of the DLR-F6 configuration are verified. The DDES predicts the drag very well whereas the URANS model significantly over predicts the drag. DDES of a finned projectile base flows is conducted to further validate the high fidelity methods with vortical flow. The DDES is demonstrated to be superior to the URANS for the projectile flow prediction. DDES of a 3D transonic wing flutter is validated with AGARD Wing 445.6 aeroelasticity experiment at free stream Mach number varied from subsonic to supersonic. The predicted flutter boundary at different free stream Mach number including the sonic dip achieves very good agreement with the experiment. In particular, the predicted flutter boundaries at the supersonic conditions match the experiment accurately. The mechanism of sonic dip is investigated. Simulation of supersonic fluid-structural interaction of a flat panel is performed by using DDES with high order shock capturing scheme. The panel vibration induced by the shock boundary layer interaction is well resolved by the high fidelity method. The dominant panel response agrees well with the experiment in terms of the mean panel displacement and frequency. The DDES methodology is used to investigate the stall inception of NASA Stage 35 compressor. The process of rotating stall is compared between the results using both URANS and DDES with full annulus. The stall process begins with spike inception and develops to full stall. The numbers of stall cell, and the size and propagating speed of the stall cells are well captured by both URANS and DDES. Two stall cells with 42% rotor rotating speed are resolved by DDES and one stall cell with 90% rotor rotating speed by URANS. It is not conclusive which method is more accurate since there is no experimental data, but the DDES does show more realistic vortical turbulence with more small scale structures. The non-synchronous vibration (NSV) of a high speed 1-1/2 stage axial compressor is investigated by using rigid blade and vibrating blade with fluid-structural interaction. An interpolation sliding boundary condition is used for the rotor-stator interaction. The URANS simulation with rigid blades shows that the leading edge(LE) circumferentially traveling vortices, roughly above 80% rotor span, travel backwards relative to the rotor rotation and cause an excitation with the frequency agreeing with the measured NSV frequency. The predicted excitation frequency of the traveling vortices in the rigid blade simulation is a non-engine order frequency of 2603 Hz, which agrees very well with the rig measured frequency of 2600 Hz. For the FSI simulation, the results show that there exist two dominant frequencies in the spectrum of the blade vibration. The lower dominant frequency is close to the first bending mode. The higher dominant frequency close to the first torsional mode agrees very well with the measured NSV frequency. To investigate whether the NSV is caused by flow excitation or by flow-structure locked-in phenomenon, the rotating speed is varied within a small RPM range, in which the rig test detected the NSV. The unsteady flows with rigid blades are simulated first at several RPMs. A dominant excitation NSV frequency caused by the circumferentially traveling tip vortices are captured. The simulation then switches to fluid structure interaction that allows the blades to vibrate freely. (Abstract shortened by ProQuest.).

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.

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.

Time-Shifted Boundary Conditions Used for Navier-Stokes Aeroelastic Solver

NASA Technical Reports Server (NTRS)

Srivastava, Rakesh

1999-01-01

Under the Advanced Subsonic Technology (AST) Program, an aeroelastic analysis code (TURBO-AE) based on Navier-Stokes equations is currently under development at NASA Lewis Research Center s Machine Dynamics Branch. For a blade row, aeroelastic instability can occur in any of the possible interblade phase angles (IBPA s). Analyzing small IBPA s is very computationally expensive because a large number of blade passages must be simulated. To reduce the computational cost of these analyses, we used time shifted, or phase-lagged, boundary conditions in the TURBO-AE code. These conditions can be used to reduce the computational domain to a single blade passage by requiring the boundary conditions across the passage to be lagged depending on the IBPA being analyzed. The time-shifted boundary conditions currently implemented are based on the direct-store method. This method requires large amounts of data to be stored over a period of the oscillation cycle. On CRAY computers this is not a major problem because solid-state devices can be used for fast input and output to read and write the data onto a disk instead of storing it in core memory.

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

NASA Technical Reports Server (NTRS)

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

1977-01-01

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

Low pressure cooling seal system for a gas turbine engine

DOEpatents

Marra, John J

2014-04-01

A low pressure cooling system for a turbine engine for directing cooling fluids at low pressure, such as at ambient pressure, through at least one cooling fluid supply channel and into a cooling fluid mixing chamber positioned immediately downstream from a row of turbine blades extending radially outward from a rotor assembly to prevent ingestion of hot gases into internal aspects of the rotor assembly. The low pressure cooling system may also include at least one bleed channel that may extend through the rotor assembly and exhaust cooling fluids into the cooling fluid mixing chamber to seal a gap between rotational turbine blades and a downstream, stationary turbine component. Use of ambient pressure cooling fluids by the low pressure cooling system results in tremendous efficiencies by eliminating the need for pressurized cooling fluids for sealing this gap.

Numerical and experimental investigation of turbine blade film cooling

NASA Astrophysics Data System (ADS)

Berkache, Amar; Dizene, Rabah

2017-12-01

The blades in a gas turbine engine are exposed to extreme temperature levels that exceed the melting temperature of the material. Therefore, efficient cooling is a requirement for high performance of the gas turbine engine. The present study investigates film cooling by means of 3D numerical simulations using a commercial code: Fluent. Three numerical models, namely k-ɛ, RSM and SST turbulence models; are applied and then prediction results are compared to experimental measurements conducted by PIV technique. The experimental model realized in the ENSEMA laboratory uses a flat plate with several rows of staggered holes. The performance of the injected flow into the mainstream is analyzed. The comparison shows that the RANS closure models improve the over-predictions of center-line film cooling velocities that is caused by the limitations of the RANS method due to its isotropy eddy diffusivity.

Reduction of turbomachinery noise

NASA Technical Reports Server (NTRS)

Waitz, Ian A. (Inventor); Hayden, Belva J. (Inventor); Ingard, K. Uno (Inventor); Brookfield, John M. (Inventor); Sell, Julian (Inventor)

1999-01-01

In the invention, propagating broad band and tonal acoustic components of noise characteristic of interaction of a turbomachine blade wake, produced by a turbomachine blade as the blade rotates, with a turbomachine component downstream of the rotating blade, are reduced. This is accomplished by injection of fluid into the blade wake through a port in the rotor blade. The mass flow rate of the fluid injected into the blade wake is selected to reduce the momentum deficit of the wake to correspondingly increase the time-mean velocity of the wake and decrease the turbulent velocity fluctuations of the wake. With this fluid injection, reduction of both propagating broad band and tonal acoustic components of noise produced by interaction of the blade wake with a turbomachine component downstream of the rotating blade is achieved. In a further noise reduction technique, boundary layer fluid is suctioned into the turbomachine blade through a suction port on the side of the blade that is characterized as the relatively low-pressure blade side. As with the fluid injection technique, the mass flow rate of the fluid suctioned into the blade is here selected to reduce the momentum deficit of the wake to correspondingly increase the time-mean velocity of the wake and decrease the turbulent velocity fluctuations of the wake; reduction of both propagating broad band and tonal acoustic components of noise produced by interaction of the blade wake with a turbomachine component downstream of the rotating blade is achieved with this suction technique. Blowing and suction techniques are also provided in the invention for reducing noise associated with the wake produced by fluid flow around a stationary blade upstream of a rotating turbomachine.

Perpendicular blade vortex interaction and its implications for helicopter noise prediction: Wave-number frequency spectra in a trailing vortex for BWI noise prediction

NASA Technical Reports Server (NTRS)

Devenport, William J.; Glegg, Stewart A. L.

1993-01-01

Perpendicular blade vortex interactions are a common occurrence in helicopter rotor flows. Under certain conditions they produce a substantial proportion of the acoustic noise. However, the mechanism of noise generation is not well understood. Specifically, turbulence associated with the trailing vortices shed from the blade tips appears insufficient to account for the noise generated. The hypothesis that the first perpendicular interaction experienced by a trailing vortex alters its turbulence structure in such a way as to increase the acoustic noise generated by subsequent interactions is examined. To investigate this hypothesis a two-part investigation was carried out. In the first part, experiments were performed to examine the behavior of a streamwise vortex as it passed over and downstream of a spanwise blade in incompressible flow. Blade vortex separations between +/- one eighth chord were studied for at a chord Reynolds number of 200,000. Three-component velocity and turbulence measurements were made in the flow from 4 chord lengths upstream to 15 chordlengths downstream of the blade using miniature 4-sensor hot wire probes. These measurements show that the interaction of the vortex with the blade and its wake causes the vortex core to loose circulation and diffuse much more rapidly than it otherwise would. Core radius increases and peak tangential velocity decreases with distance downstream of the blade. True turbulence levels within the core are much larger downstream than upstream of the blade. The net result is a much larger and more intense region of turbulent flow than that presented by the original vortex and thus, by implication, a greater potential for generating acoustic noise. In the second part, the turbulence measurements described above were used to derive the necessary inputs to a Blade Wake Interaction (BWI) noise prediction scheme. This resulted in significantly improved agreement between measurements and calculations of the BWI noise spectrum especially for the spectral peak at low frequencies, which previously was poorly predicted.

Flow structure generated by perpendicular blade-vortex interaction and implications for helicopter noise prediction. Volume 1: Measurements

NASA Technical Reports Server (NTRS)

Wittmer, Kenneth S.; Devenport, William J.

1996-01-01

The perpendicular interaction of a streamwise vortex with an infinite span helicopter blade was modeled experimentally in incompressible flow. Three-component velocity and turbulence measurements were made using a sub-miniature four sensor hot-wire probe. Vortex core parameters (radius, peak tangential velocity, circulation, and centerline axial velocity deficit) were determined as functions of blade-vortex separation, streamwise position, blade angle of attack, vortex strength, and vortex size. The downstream development of the flow shows that the interaction of the vortex with the blade wake is the primary cause of the changes in the core parameters. The blade sheds negative vorticity into its wake as a result of the induced angle of attack generated by the passing vortex. Instability in the vortex core due to its interaction with this negative vorticity region appears to be the catalyst for the magnification of the size and intensity of the turbulent flowfield downstream of the interaction. In general, the core radius increases while peak tangential velocity decreases with the effect being greater for smaller separations. These effects are largely independent of blade angle of attack; and if these parameters are normalized on their undisturbed values, then the effects of the vortex strength appear much weaker. Two theoretical models were developed to aid in extending the results to other flow conditions. An empirical model was developed for core parameter prediction which has some rudimentary physical basis, implying usefulness beyond a simple curve fit. An inviscid flow model was also created to estimate the vorticity shed by the interaction blade, and to predict the early stages of its incorporation into the interacting vortex.

Design geometry and design/off-design performance computer codes for compressors and turbines

NASA Technical Reports Server (NTRS)

Glassman, Arthur J.

1995-01-01

This report summarizes some NASA Lewis (i.e., government owned) computer codes capable of being used for airbreathing propulsion system studies to determine the design geometry and to predict the design/off-design performance of compressors and turbines. These are not CFD codes; velocity-diagram energy and continuity computations are performed fore and aft of the blade rows using meanline, spanline, or streamline analyses. Losses are provided by empirical methods. Both axial-flow and radial-flow configurations are included.

The design and instrumentation of the Purdue annular cascade facility with initial data acquisition and analysis

NASA Technical Reports Server (NTRS)

Stauter, R. C.; Fleeter, S.

1982-01-01

Three dimensional aerodynamic data, required to validate and/or indicate necessary refinements to inviscid and viscous analyses of the flow through turbomachine blade rows, are discussed. Instrumentation and capabilities for pressure measurement, probe insertion and traversing, and flow visualization are reviewed. Advanced measurement techniques including Laser Doppler Anemometers, are considered. Data processing is reviewed. Predictions were correlated with the experimental data. A flow visualization technique using helium filled soap bubbles was demonstrated.

Leaf primordium size specifies leaf width and vein number among row-type classes in barley.

PubMed

Thirulogachandar, Venkatasubbu; Alqudah, Ahmad M; Koppolu, Ravi; Rutten, Twan; Graner, Andreas; Hensel, Goetz; Kumlehn, Jochen; Bräutigam, Andrea; Sreenivasulu, Nese; Schnurbusch, Thorsten; Kuhlmann, Markus

2017-08-01

Exploring genes with impact on yield-related phenotypes is the preceding step to accomplishing crop improvements while facing a growing world population. A genome-wide association scan on leaf blade area (LA) in a worldwide spring barley collection (Hordeum vulgare L.), including 125 two- and 93 six-rowed accessions, identified a gene encoding the homeobox transcription factor, Six-rowed spike 1 (VRS1). VRS1 was previously described as a key domestication gene affecting spike development. Its mutation converts two-rowed (wild-type VRS1, only central fertile spikelets) into six-rowed spikes (mutant vrs1, fully developed fertile central and lateral spikelets). Phenotypic analyses of mutant and wild-type leaves revealed that mutants had an increased leaf width with more longitudinal veins. The observed significant increase of LA and leaf nitrogen (%) during pre-anthesis development in vrs1 mutants also implies a link between wider leaf and grain number, which was validated from the association of vrs1 locus with wider leaf and grain number. Histological and gene expression analyses indicated that VRS1 might influence the size of leaf primordia by affecting cell proliferation of leaf primordial cells. This finding was supported by the transcriptome analysis of mutant and wild-type leaf primordia where in the mutant transcriptional activation of genes related to cell proliferation was detectable. Here we show that VRS1 has an independent role on barley leaf development which might influence the grain number. © 2017 The Authors. The Plant Journal published by John Wiley & Sons Ltd and Society for Experimental Biology.

Rotorcraft Blade-Vortex Interaction Controller

NASA Technical Reports Server (NTRS)

Schmitz, Fredric H. (Inventor)

1995-01-01

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

Numerical simulation of blade-passage noise.

PubMed

Yauwenas, Yendrew; Zajamšek, Branko; Reizes, John; Timchenko, Victoria; Doolan, Con J

2017-09-01

Numerical simulations are used to investigate the noise generated by the passage of a rotor blade past a fixed object (the blade-passage effects), which was studied by simulating a three-bladed rotor that is supported by a vertical cylindrical tower. To isolate the blade-passage effects, no incoming wind was introduced in the simulation. The symmetric blade was set to zero pitch angle relative to the plane of rotation and two blade-tower distances were investigated. The sliding mesh method was used to simulate the rotation of the blades and Curle's acoustic analogy was used to predict the noise generated from the simulated flow data. Intense force fluctuations occur during the interaction on both the tower and the passing blade, and these are the primary sources of blade-passage noise. The contribution of the force fluctuations on the support tower to blade-passage noise, which previously had been ignored, was revealed to be more significant than that of the blades. The numerical model successfully predicts the noise spectra, which are validated by the very good agreement with experimental measurements. The simulations provide a framework to better understand blade-tower interaction noise in various applications.

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.

Turbofan gas turbine engine with variable fan outlet guide vanes

NASA Technical Reports Server (NTRS)

Wood, Peter John (Inventor); LaChapelle, Donald George (Inventor); Grant, Carl (Inventor); Zenon, Ruby Lasandra (Inventor); Mielke, Mark Joseph (Inventor)

2010-01-01

A turbofan gas turbine engine includes a forward fan section with a row of fan rotor blades, a core engine, and a fan bypass duct downstream of the forward fan section and radially outwardly of the core engine. The forward fan section has only a single stage of variable fan guide vanes which are variable fan outlet guide vanes downstream of the forward fan rotor blades. An exemplary embodiment of the engine includes an afterburner downstream of the fan bypass duct between the core engine and an exhaust nozzle. The variable fan outlet guide vanes are operable to pivot from a nominal OGV position at take-off to an open OGV position at a high flight Mach Number which may be in a range of between about 2.5-4+. Struts extend radially across a radially inwardly curved portion of a flowpath of the engine between the forward fan section and the core engine.

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.

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.

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.

Measurement and analysis of the noise radiated by low Mach numbers centrifugal blowers

NASA Astrophysics Data System (ADS)

Yeager, D. M.; Lauchle, G. C.

1987-11-01

The broad band, aerodynamically generated noise in low tip-speed Mach number, centrifugal air moving devices is investigated. An interdisciplinary approach was taken which involved investigation of the aerodynamic and acoustic fields, and their mutual relationship. The noise generation process was studied using two experimental vehicles: (1) a scale model of a homologous family of centrifugal blowers typical of those used to cool computer and business equipment, and (2) a single blade from a centrifugal blower impeller which was placed in a known, controllable flow field. The radiation characteristics of the model blower were investigated by measuring the acoustic intensity distribution near the blower inlet and comparing it with the intensity near the inlet to an axial flow fan. Aerodynamic studies of the flow field in the inlet and at the discharge to the rotating impeller were used to assess the mean flow distribution through the impeller blade channels and to identify regions of excessive turbulence near the rotating blade row. New frequency-domain expressions for the correlation area and dipole source strength per unit area on a surface immersed in turbulence were developed which can be used to characterize the noise generation process over a rigid surface immersed in turbulence. An investigation of the noise radiated from the single, isolated airfoil (impeller blade) was performed using modern correlation and spectral analysis techniques.

Single-stage experimental evaluation of tandem-airfoil rotor stator blading for compressors. Part 6: Data and performance for stage D

NASA Technical Reports Server (NTRS)

Clemmons, D. R.

1973-01-01

An axial flow compressor stage, having single-airfoil blading, was designed for zero rotor prewhirl, constant rotor work across the span, and axial discharge flow. The stage was designed to produce a pressure ratio of 1.265 at a rotor tip velocity of 757 ft/sec. The rotor had an inlet hub/tip ratio of 0.8. The design procedure accounted for the rotor inlet boundary layer and included the effects of axial velocity ratio and secondary flow on blade row performance. The objectives of this experimental program were: (1) to obtain performance with uniform and distorted inlet flow for comparison with the performance of a stage consisting of tandem-airfoil blading designed for the same vector diagrams; and (2) to evaluate the effectiveness of accounting for the inlet boundary layer, axial velocity ratio, and secondary flows in the stage design. With uniform inlet flow, the rotor achieved a maximum adiabatic efficiency of 90.1% at design equivalent rotor speed and a pressure ratio of 1.281. The stage maximum adiabatic efficiency at design equivalent rotor speed with uniform inlet flow was 86.1% at a pressure ratio of 1.266. Hub radial, tip radial, and circumferential distortion of the inlet flow caused reductions in surge pressure ratio of approximately 2, 10 and 5%, respectively, at design rotor speed.

Supramolecular order and structural dynamics: A STM study of 2H-tetraphenylporphycene on Cu(111)

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

Stark, Michael; Träg, Johannes; Ditze, Stefanie

2015-03-14

The adsorption of 2H-tetraphenylporphycene (2HTPPc) on Cu(111) was investigated by scanning tunneling microscopy (STM). At medium coverages, supramolecular ordered islands are observed. The individual 2HTPPc molecules appear as two pairs of intense protrusions which are separated by an elongated depression. In the islands, the molecules are organized in rows oriented along one of the close packed Cu(111) substrate rows; the structure is stabilized by T-type interactions of the phenyl substituents of neighboring molecules. Two types of rows are observed, namely, highly ordered rows in which all molecules exhibit the same orientation, and less ordered rows in which the molecules exhibitmore » two perpendicular orientations. Altogether, three different azimuthal orientations of 2HTPPc are observed within one domain, all of them rotated by 15° ± 1° relative to one closed packed Cu direction. The highly ordered rows are always separated by either one or two less ordered rows, with the latter structure being the thermodynamically more stable one. The situation in the islands is highly dynamic, such that molecules in the less ordered rows occasionally change orientation, also complete highly ordered rows can move. The supramolecular order and structural dynamics are discussed on the basis of the specific molecule-substrate and molecule-molecule interactions.« less

Flow visualizations of perpendicular blade vortex interactions

NASA Technical Reports Server (NTRS)

Rife, Michael C.; Davenport, William J.

1992-01-01

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

Weight Assessment for Fuselage Shielding on Aircraft With Open-Rotor Engines and Composite Blade Loss

NASA Technical Reports Server (NTRS)

Carney, Kelly; Pereira, Michael; Kohlman, Lee; Goldberg, Robert; Envia, Edmane; Lawrence, Charles; Roberts, Gary; Emmerling, William

2013-01-01

The Federal Aviation Administration (FAA) has been engaged in discussions with airframe and engine manufacturers concerning regulations that would apply to new technology fuel efficient "openrotor" engines. Existing regulations for the engines and airframe did not envision features of these engines that include eliminating the fan blade containment systems and including two rows of counter-rotating blades. Damage to the airframe from a failed blade could potentially be catastrophic. Therefore the feasibility of using aircraft fuselage shielding was investigated. In order to establish the feasibility of this shielding, a study was conducted to provide an estimate for the fuselage shielding weight required to provide protection from an open-rotor blade loss. This estimate was generated using a two-step procedure. First, a trajectory analysis was performed to determine the blade orientation and velocity at the point of impact with the fuselage. The trajectory analysis also showed that a blade dispersion angle of 3deg bounded the probable dispersion pattern and so was used for the weight estimate. Next, a finite element impact analysis was performed to determine the required shielding thickness to prevent fuselage penetration. The impact analysis was conducted using an FAA-provided composite blade geometry. The fuselage geometry was based on a medium-sized passenger composite airframe. In the analysis, both the blade and fuselage were assumed to be constructed from a T700S/PR520 triaxially-braided composite architecture. Sufficient test data on T700S/PR520 is available to enable reliable analysis, and also demonstrate its good impact resistance properties. This system was also used in modeling the surrogate blade. The estimated additional weight required for fuselage shielding for a wing- mounted counterrotating open-rotor blade is 236 lb per aircraft. This estimate is based on the shielding material serving the dual use of shielding and fuselage structure. If the shielding material is not used for dual purpose, and is only used for shielding, then the additional weight per aircraft is estimated to be 428 lb. This weight estimate is based upon a number of assumptions that would need to be revised when applying this concept to an actual airplane design. For example, the weight savings that will result when there is no fan blade containment system, manufacturing limitations which may increase the weight where variable thicknesses was assumed, engine placement on the wing versus aft fuselage, etc.

Blade vortex interaction noise reduction techniques for a rotorcraft

NASA Technical Reports Server (NTRS)

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

1996-01-01

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

Blade vortex interaction noise reduction techniques for a rotorcraft

NASA Technical Reports Server (NTRS)

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

1998-01-01

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

Lift distribution and velocity field measurements for a three-dimensional, steady blade/vortex interaction

NASA Technical Reports Server (NTRS)

Dunagan, Stephen E.; Norman, Thomas R.

1987-01-01

A wind tunnel experiment simulating a steady three-dimensional helicopter rotor blade/vortex interaction is reported. The experimental configuration consisted of a vertical semispan vortex-generating wing, mounted upstream of a horizontal semispan rotor blade airfoil. A three-dimensional laser velocimeter was used to measure the velocity field in the region of the blade. Sectional lift coefficients were calculated by integrating the velocity field to obtain the bound vorticity. Total lift values, obtained by using an internal strain-gauge balance, verified the laser velocimeter data. Parametric variations of vortex strength, rotor blade angle of attack, and vortex position relative to the rotor blade were explored. These data are reported (with attention to experimental limitations) to provide a dataset for the validation of analytical work.

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.

Cascade aeroacoustics including steady loading effects

NASA Astrophysics Data System (ADS)

Chiang, Hsiao-Wei D.; Fleeter, Sanford

A mathematical model is developed to analyze the effects of airfoil and cascade geometry, steady aerodynamic loading, and the characteristics of the unsteady flow field on the discrete frequency noise generation of a blade row in an incompressible flow. The unsteady lift which generates the noise is predicted with a complex first-order cascade convected gust analysis. This model was then applied to the Gostelow airfoil cascade and variations, demonstrating that steady loading, cascade solidity, and the gust direction are significant. Also, even at zero incidence, the classical flat plate cascade predictions are unacceptable.

Gas turbine engine with radial diffuser and shortened mid section

DOEpatents

Charron, Richard C.; Montgomery, Matthew D.

2015-09-08

An industrial gas turbine engine (10), including: a can annular combustion assembly (80), having a plurality of discrete flow ducts configured to receive combustion gas from respective combustors (82) and deliver the combustion gas along a straight flow path at a speed and orientation appropriate for delivery directly onto the first row (56) of turbine blades (62); and a compressor diffuser (32) having a redirecting surface (130, 140) configured to receive an axial flow of compressed air and redirect the axial flow of compressed air radially outward.

Performance of a tandem-rotor/tandem-stator conical-flow compressor designed for a pressure ratio of 3

NASA Technical Reports Server (NTRS)

Wood, J. R.; Owen, A. K.; Schumann, L. F.

1982-01-01

A conical-flow compressor stage with a large radius change through the rotor was tested at three values of rotor tip clearance. The stage had a tandem rotor and a tandem stator. Peak efficiency at design speed was 0.774 at a pressure ratio of 2.613. The rotor was tested without the stator, and detailed survey data were obtained for each rotor blade row. Overall peak rotor efficiency was 0.871 at a pressure ratio of 2.952.

A three-dimensional structured/unstructured hybrid Navier-Stokes method for turbine blade rows

NASA Technical Reports Server (NTRS)

Tsung, F.-L.; Loellbach, J.; Kwon, O.; Hah, C.

1994-01-01

A three-dimensional viscous structured/unstructured hybrid scheme has been developed for numerical computation of high Reynolds number turbomachinery flows. The procedure allows an efficient structured solver to be employed in the densely clustered, high aspect-ratio grid around the viscous regions near solid surfaces, while employing an unstructured solver elsewhere in the flow domain to add flexibility in mesh generation. Test results for an inviscid flow over an external transonic wing and a Navier-Stokes flow for an internal annular cascade are presented.

Numerical Studies of the Radiation Condition at the Inlet of a Transonic Compressor Blade Row.

DTIC Science & Technology

1980-04-01

284-296 (1977). 2 aL.) ITH~rOU7 REFLECTOAJ &RWD b)WITH PJEFL.ECr-IOA, FIGLJ/AE I NV9 PATTERNhS EXPEC TED located a chord length or so upstream of the...C) W In -0N -W e --N 0 -00 e -’N -" M-M M V -NM ""NM N NM 0 - - - - - - - - - - - - --4 WUN NWCON NMWN NN VN N-CYN N N NOJcyN N-N N-N N O~j N 000 00

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.

Repulsive Interaction of Sulfide Layers on Compressor Impeller Blades Remanufactured Through Plasma Spray Welding

NASA Astrophysics Data System (ADS)

Chang, Y.; Zhou, D.; Wang, Y. L.; Huang, H. H.

2016-12-01

This study investigated the repulsive interaction of sulfide layers on compressor impeller blades remanufactured through plasma spray welding (PSW). Sulfide layers on the blades made of FV(520)B steel were prepared through multifarious corrosion experiments, and PSW was utilized to remanufacture blade specimens. The specimens were evaluated through optical microscopy, scanning electron microscopy, energy-dispersive spectroscopy, 3D surface topography, x-ray diffraction, ImageJ software analysis, Vicker's micro-hardness test and tensile tests. Results showed a large number of sulfide inclusions in the fusion zone generated by sulfide layers embodied into the molten pool during PSW. These sulfide inclusions seriously degraded the mechanical performance of the blades remanufactured through PSW.

Interaction of Atmospheric Turbulence with Blade Boundary Layer Dynamics on a 5MW Wind Turbine using Blade-Boundary-Layer-Resolved CFD with hybrid URANS-LES.

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

Vijayakumar, Ganesh; Brasseur, James; Lavely, Adam

We describe the response of the NREL 5 MW wind turbine blade boundary layer to the passage of atmospheric turbulence using blade-boundary-layer-resolved computational fluid dynamics with hybrid URANS-LES modeling.

Transonic Axial Splittered Rotor Tandem Stator Stage

DTIC Science & Technology

2016-12-01

CODE 13. ABSTRACT (maximum 200 words) Development of a procedure to model the hot shape of a rotor blade and a comparison analysis of the transonic...fluid-structure interaction. Rotational forces as well as gas loading forces were observed as an influence on blade deformation. Utilizing the...Turbomachinery, splittered rotor, tandem stator, transonic compressor, blade deformation, fluid-structure interaction 15. NUMBER OF PAGES 87 16. PRICE

Prediction of Undsteady Flows in Turbomachinery Using the Linearized Euler Equations on Deforming Grids

NASA Technical Reports Server (NTRS)

Clark, William S.; Hall, Kenneth C.

1994-01-01

A linearized Euler solver for calculating unsteady flows in turbomachinery blade rows due to both incident gusts and blade motion is presented. The model accounts for blade loading, blade geometry, shock motion, and wake motion. Assuming that the unsteadiness in the flow is small relative to the nonlinear mean solution, the unsteady Euler equations can be linearized about the mean flow. This yields a set of linear variable coefficient equations that describe the small amplitude harmonic motion of the fluid. These linear equations are then discretized on a computational grid and solved using standard numerical techniques. For transonic flows, however, one must use a linear discretization which is a conservative linearization of the non-linear discretized Euler equations to ensure that shock impulse loads are accurately captured. Other important features of this analysis include a continuously deforming grid which eliminates extrapolation errors and hence, increases accuracy, and a new numerically exact, nonreflecting far-field boundary condition treatment based on an eigenanalysis of the discretized equations. Computational results are presented which demonstrate the computational accuracy and efficiency of the method and demonstrate the effectiveness of the deforming grid, far-field nonreflecting boundary conditions, and shock capturing techniques. A comparison of the present unsteady flow predictions to other numerical, semi-analytical, and experimental methods shows excellent agreement. In addition, the linearized Euler method presented requires one or two orders-of-magnitude less computational time than traditional time marching techniques making the present method a viable design tool for aeroelastic analyses.

Three-dimensional Aerodynamic Instability in Multi-stage Axial Compressors

NASA Technical Reports Server (NTRS)

Suder, Kenneth (Technical Monitor); Tan, Choon-Sooi

2003-01-01

Four separate tasks are reported. The first task: A Computational Model for Short Wavelength Stall Inception and Development In Multi-Stage Compressors; the second task: Three-dimensional Rotating Stall Inception and Effects of Rotating Tip Clearance Asymmetry in Axial Compressors; the third task:Development of an Effective Computational Methodology for Body Force Representation of High-speed Rotor 37; and the fourth task:Development of Circumferential Inlet Distortion through a Representative Eleven Stage High-speed axial compressor. The common theme that threaded throughout these four tasks is the conceptual framework that consists of quantifying flow processes at the fadcompressor blade passage level to define the compressor performance characteristics needed for addressing physical phenomena such compressor aerodynamic instability and compressor response to flow distoriton with length scales larger than compressor blade-to-blade spacing at the system level. The results from these two levels can be synthesized to: (1) simulate compressor aerodynamic instability inception local to a blade rotor tip and its development from a local flow event into the nonlinear limit cycle instability that involves the entire compressor as was demonstrated in the first task; (2) determine the conditions under which compressor stability assessment based on two-dimensional model may not be adequate and the effects of self-induced flow distortion on compressor stability limit as in the second task; (3) quantify multistage compressor response to inlet distortion in stagnation pressure as illustrated in the fourth task; and (4) elucidate its potential applicability for compressor map generation under uniform as well as non-uniform inlet flow given three-dimensional Navier-Stokes solution for each individual blade row as was demonstrated in the third task.

Self-organization of S adatoms on Au(111): √3R30° rows at low coverage

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

Walen, Holly, E-mail: hwalen@iastate.edu; Liu, Da-Jiang; Oh, Junepyo

Using scanning tunneling microscopy, we observe an adlayer structure that is dominated by short rows of S atoms, on unreconstructed regions of a Au(111) surface. This structure forms upon adsorption of low S coverage (less than 0.1 monolayer) on a fully reconstructed clean surface at 300 K, then cooling to 5 K for observation. The rows adopt one of three orientations that are rotated by 30° from the close-packed directions of the Au(111) substrate, and adjacent S atoms in the rows are separated by √3 times the surface lattice constant, a. Monte Carlo simulations are performed on lattice-gas models, derivedmore » using a limited cluster expansion based on density functional theory energetics. Models which include long-range pairwise interactions (extending to 5a), plus selected trio interactions, successfully reproduce the linear rows of S atoms at reasonable temperatures.« less

Self-organization of S adatoms on Au(111): √3R30° rows at low coverage

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

Walen, Holly; Liu, Da -Jiang; Oh, Junepyo

Using scanning tunneling microscopy, we observe an adlayer structure that is dominated by short rows of S atoms, on unreconstructed regions of a Au(111) surface. This structure forms upon adsorption of low S coverage (less than 0.1 monolayer) on a fully reconstructed cleansurface at 300 K, then cooling to 5 K for observation. The rows adopt one of three orientations that are rotated by 30° from the close-packed directions of the Au(111) substrate, and adjacent S atoms in the rows are separated by √3 times the surface lattice constant, a. Monte Carlo simulations are performed on lattice-gas models, we derivedmore » using a limited cluster expansion based on density functional theory energetics. Furthermore, models which include long-range pairwise interactions (extending to 5a), plus selected trio interactions, successfully reproduce the linear rows of S atoms at reasonable temperatures.« less

Self-organization of S adatoms on Au(111): √3R30° rows at low coverage

DOE PAGES

Walen, Holly; Liu, Da -Jiang; Oh, Junepyo; ...

2015-07-06

Using scanning tunneling microscopy, we observe an adlayer structure that is dominated by short rows of S atoms, on unreconstructed regions of a Au(111) surface. This structure forms upon adsorption of low S coverage (less than 0.1 monolayer) on a fully reconstructed cleansurface at 300 K, then cooling to 5 K for observation. The rows adopt one of three orientations that are rotated by 30° from the close-packed directions of the Au(111) substrate, and adjacent S atoms in the rows are separated by √3 times the surface lattice constant, a. Monte Carlo simulations are performed on lattice-gas models, we derivedmore » using a limited cluster expansion based on density functional theory energetics. Furthermore, models which include long-range pairwise interactions (extending to 5a), plus selected trio interactions, successfully reproduce the linear rows of S atoms at reasonable temperatures.« less

Self-organization of S adatoms on Au(111): √3R30° rows at low coverage.

PubMed

Walen, Holly; Liu, Da-Jiang; Oh, Junepyo; Lim, Hyunseob; Evans, J W; Kim, Yousoo; Thiel, P A

2015-07-07

Using scanning tunneling microscopy, we observe an adlayer structure that is dominated by short rows of S atoms, on unreconstructed regions of a Au(111) surface. This structure forms upon adsorption of low S coverage (less than 0.1 monolayer) on a fully reconstructed clean surface at 300 K, then cooling to 5 K for observation. The rows adopt one of three orientations that are rotated by 30° from the close-packed directions of the Au(111) substrate, and adjacent S atoms in the rows are separated by √3 times the surface lattice constant, a. Monte Carlo simulations are performed on lattice-gas models, derived using a limited cluster expansion based on density functional theory energetics. Models which include long-range pairwise interactions (extending to 5a), plus selected trio interactions, successfully reproduce the linear rows of S atoms at reasonable temperatures.

Multifidelity, multidisciplinary optimization of turbomachines with shock interaction

NASA Astrophysics Data System (ADS)

Joly, Michael Marie

Research on high-speed air-breathing propulsion aims at developing aircraft with antipodal range and space access. Before reaching high speed at high altitude, the flight vehicle needs to accelerate from takeoff to scramjet takeover. Air turbo rocket engines combine turbojet and rocket engine cycles to provide the necessary thrust in the so-called low-speed regime. Challenges related to turbomachinery components are multidisciplinary, since both the high compression ratio compressor and the powering high-pressure turbine operate in the transonic regime in compact environments with strong shock interactions. Besides, lightweight is vital to avoid hindering the scramjet operation. Recent progress in evolutionary computing provides aerospace engineers with robust and efficient optimization algorithms to address concurrent objectives. The present work investigates Multidisciplinary Design Optimization (MDO) of innovative transonic turbomachinery components. Inter-stage aerodynamic shock interaction in turbomachines are known to generate high-cycle fatigue on the rotor blades compromising their structural integrity. A soft-computing strategy is proposed to mitigate the vane downstream distortion, and shown to successfully attenuate the unsteady forcing on the rotor of a high-pressure turbine. Counter-rotation offers promising prospects to reduce the weight of the machine, with fewer stages and increased load per row. An integrated approach based on increasing level of fidelity and aero-structural coupling is then presented and allows achieving a highly loaded compact counter-rotating compressor.

A Study of Fan Stage/Casing Interaction Models

NASA Technical Reports Server (NTRS)

Lawrence, Charles; Carney, Kelly; Gallardo, Vicente

2003-01-01

The purpose of the present study is to investigate the performance of several existing and new, blade-case interactions modeling capabilities that are compatible with the large system simulations used to capture structural response during blade-out events. Three contact models are examined for simulating the interactions between a rotor bladed disk and a case: a radial and linear gap element and a new element based on a hydrodynamic formulation. The first two models are currently available in commercial finite element codes such as NASTRAN and have been showed to perform adequately for simulating rotor-case interactions. The hydrodynamic model, although not readily available in commercial codes, may prove to be better able to characterize rotor-case interactions.

Blade motion and nutrient flux to the kelp, Eisenia arborea.

PubMed

Denny, Mark; Roberson, Loretta

2002-08-01

Marine algae rely on currents and waves to replenish the nutrients required for photosynthesis. The interaction of algal blades with flow often involves dynamic reorientations of the blade surface (pitching and flapping) that may in turn affect nutrient flux. As a first step toward understanding the consequences of blade motion, we explore the effect of oscillatory pitching on the flux to a flat plate and to two morphologies of the kelp Eisenia arborea. In slow flow (equivalent to a water velocity of 2.7 cm s(-1)), pitching increases the time-averaged flux to both kelp morphologies, but not to the plate. In fast flow (equivalent to 20 cm s(-1) in water), pitching has negligible effect on flux regardless of shape. For many aspects of flux, the flat plate is a reliable model for the flow-protected algal blade, but predictions made from the plate would substantially underestimate the flux to the flow-exposed blade. These measurements highlight the complexities of flow-related nutrient transport and the need to understand better the dynamic interactions among nutrient flux, blade motion, blade morphology, and water flow.

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.

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

NASA Technical Reports Server (NTRS)

Hoad, D. R.

1979-01-01

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

An Experimental Investigation of Steady and Unsteady Flow Field in an Axial Flow Turbine

NASA Technical Reports Server (NTRS)

Zaccaria, M.; Lakshminarayana, B.

1997-01-01

Measurements were made in a large scale single stage turbine facility. Within the nozzle passage measurements were made using a five hole probe, a two-component Laser Doppler Velocimeter (LDV), and a single sensor hot wire probe. These measurements showed weak secondary flows at midchord, and two secondary flow loss cores at the nozzle exit. The casing vortex loss core was the larger of the two. At the exit radial inward flow was found over the entire passage, and was more pronounced in the wake. Nozzle wake decay was found to be more rapid than for an isolated vane row due to the rotor's presence. The midspan rotor flow field was measured using a two-component LDV. Measurements were made from upstream of the rotor to a chord behind the rotor. The distortion of the nozzle wake as it passed through the rotor blade row was determined. The unsteadiness in the rotor flow field was determined. The decay of the rotor wake was also characterized.

Some design philosophy for reducing the community noise of advanced counter-rotation propellers

NASA Technical Reports Server (NTRS)

Dittmar, J. H.

1985-01-01

Advanced counter-rotation propellers have been indicated as possibly generating an unacceptable amount of noise for the people living near an airport. This report has explored ways to reduce this noise level, which is treated as being caused by the interaction of the upstream propeller wakes and vortices with the downstream propeller. The noise reduction techniques fall into two categories: (1) reducing the strength of the wakes and vortices, and (2) reducing the response of the downstream blades to them. The noise from the wake interaction was indicated as being reduced by increased propeller spacing and decreased blade drag coefficient. The vortex-interaction noise could be eliminated by having the vortex pass over the tips of the downstream blade, and it could be reduced by increased spacing or decreased initial circulation. The downstream blade response could be lessened by increasing the reduced frequency parameter omega or by phasing of the response from different sections to have a mutual cancellation effect. Uneven blade to blade spacing for the downstream blading was indicated as having a possible effect on the annoyance of counter-rotation propeller noise. Although there are undoubtedly additional methods of noise reduction not covered in this report, the inclusion of the design methods discussed would potentially result in a counter-rotation propeller that is acceptably quiet.

Rotor Broadband Noise Prediction with Comparison to Model Data

NASA Technical Reports Server (NTRS)

Brooks, Thomas F.; Burley, Casey L.

2001-01-01

This paper reports an analysis and prediction development of rotor broadband noise. The two primary components of this noise are Blade-Wake Interaction (BWI) noise, due to the blades' interaction with the turbulent wakes of the preceding blades, and "Self" noise, due to the development and shedding of turbulence within the blades' boundary layers. Emphasized in this report is the new code development for Self noise. The analysis and validation employs data from the HART program, a model BO-105 rotor wind tunnel test conducted in the German-Dutch Wind Tunnel (DNW). The BWI noise predictions are based on measured pressure response coherence functions using cross-spectral methods. The Self noise predictions are based on previously reported semiempirical modeling of Self noise obtained from isolated airfoil sections and the use of CAMRAD.Modl to define rotor performance and local blade segment flow conditions. Both BWI and Self noise from individual blade segments are Doppler shifted and summed at the observer positions. Prediction comparisons with measurements show good agreement for a range of rotor operating conditions from climb to steep descent. The broadband noise predictions, along with those of harmonic and impulsive Blade-Vortex Interaction (BVI) noise predictions, demonstrate a significant advance in predictive capability for main rotor noise.

Interactive multi-mode blade impact analysis

NASA Technical Reports Server (NTRS)

Alexander, A.; Cornell, R. W.

1978-01-01

The theoretical methodology used in developing an analysis for the response of turbine engine fan blades subjected to soft-body (bird) impacts is reported, and the computer program developed using this methodology as its basis is described. This computer program is an outgrowth of two programs that were previously developed for the purpose of studying problems of a similar nature (a 3-mode beam impact analysis and a multi-mode beam impact analysis). The present program utilizes an improved missile model that is interactively coupled with blade motion which is more consistent with actual observations. It takes into account local deformation at the impact area, blade camber effects, and the spreading of the impacted missile mass on the blade surface. In addition, it accommodates plate-type mode shapes. The analysis capability in this computer program represents a significant improvement in the development of the methodology for evaluating potential fan blade materials and designs with regard to foreign object impact resistance.

Fabrication of composite propfan blades for a cruise missile wind tunnel model

NASA Technical Reports Server (NTRS)

Fite, E. Brian

1993-01-01

This report outlines the procedures that were employed in fabricating prototype graphite-epoxy composite prop fan blades. These blades were used in wind tunnel tests that investigated prop fan propulsion system interactions with a missile airframe in order to study the feasibility of an advanced-technology-propfan-propelled missile. Major phases of the blade fabrication presented include machining of the master blade, mold fabrication, ply cutting and assembly, blade curing, and quality assurance. Specifically, four separate designs were fabricated, 18 blades of each geometry, using the same fabrication technique for each design.

Bird impact analysis package for turbine engine fan blades

NASA Technical Reports Server (NTRS)

Hirschbein, M. S.

1982-01-01

A computer program has been developed to analyze the gross structural response of turbine engine fan blades subjected to bird strikes. The program couples a NASTRAN finite element model and modal analysis of a fan blade with a multi-mode bird impact analysis computer program. The impact analysis uses the NASTRAN blade model and a fluid jet model of the bird to interactively calculate blade loading during a bird strike event. The analysis package is computationaly efficient, easy to use and provides a comprehensive history of the gross structual blade response. Example cases are presented for a representative fan blade.

The high Reynolds number flow through an axial-flow pump

NASA Astrophysics Data System (ADS)

Zierke, W. C.; Straka, W. A.; Taylor, P. D.

1993-11-01

The high Reynolds number pump (HIREP) facility at ARL Penn State has been used to perform a low-speed, large-scale experiment of the incompressible flow of water through a two-blade-row turbomachine. HIREP can involve blade chord Reynolds numbers as high as 6,000,000 and can accommodate a variety of instrumentation in both a stationary and a rotating frame of reference. The objectives of this experiment were as follows: to provide a database for comparison with three-dimensional, viscous (turbulent) flow computations; to evaluate the engineering models; and to improve our physical understanding of many of the phenomena involved in this complex flow field. The experimental results include a large quantity of data acquired throughout HIREP. A five-hole probe survey of the inlet flow 37.0 percent chord upstream of the inlet guide vane (IGV) leading edge is sufficient to give information for the inflow boundary conditions, while some static-pressure information is available to help establish an outflow boundary condition.

Acoustic and aerodynamic performance of a 1.83 meter (6 foot) diameter 1.2 pressure ratio fan (QF-6). [for short takeoff aircraft

NASA Technical Reports Server (NTRS)

Woodward, R. P.; Lucas, J. G.; Stakolich, E. G.

1974-01-01

A 1.2-pressure-ratio, 1.83-meter-(6-ft-) diameter experimental fan stage with characteristics suitable for use in STOL aircraft engines was tested for acoustic and aerodynamic performance. The design incorporated features for low noise, including absence of inlet guide vanes, low rotor-blade-tip speed, low aerodynamic blade loading, and long axial spacing between the rotor and stator rows. The stage was run with four nozzles of different area. The perceived noise along a 152.4 meter (500-ft) sideline was rear-quadrant dominated with a maximum design-point level of 103.9 PNdb. The acoustic 1/3-octave results were analytically separated into broadband and pure-tone components. It was found that the stage noise levels generally increase with a decrease in nozzle area, with this increase observed primarily in the broadband noise component. A stall condition was documented acoustically with a 90-percent-of-design-area nozzle.

Nonlinear (time domain) and linearized (time and frequency domain) solutions to the compressible Euler equations in conservation law form

NASA Technical Reports Server (NTRS)

Sreenivas, Kidambi; Whitfield, David L.

1995-01-01

Two linearized solvers (time and frequency domain) based on a high resolution numerical scheme are presented. The basic approach is to linearize the flux vector by expressing it as a sum of a mean and a perturbation. This allows the governing equations to be maintained in conservation law form. A key difference between the time and frequency domain computations is that the frequency domain computations require only one grid block irrespective of the interblade phase angle for which the flow is being computed. As a result of this and due to the fact that the governing equations for this case are steady, frequency domain computations are substantially faster than the corresponding time domain computations. The linearized equations are used to compute flows in turbomachinery blade rows (cascades) arising due to blade vibrations. Numerical solutions are compared to linear theory (where available) and to numerical solutions of the nonlinear Euler equations.

An Experimental Study of the Effect of Wake Passing on Turbine Blade Film Cooling

NASA Technical Reports Server (NTRS)

Heidmann, James D.; Lucci, Barbara L.; Reshotko, Eli

1997-01-01

The effect of wake passing on the showerhead film cooling performance of a turbine blade has been investigated experimentally. The experiments were performed in an annular turbine cascade with an upstream rotating row of cylindrical rods. Nickel thin-film gauges were used to determine local film effectiveness and Nusselt number values for various injectants, blowing ratios, and Strouhal numbers. Results indicated a 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 by a streamwise-constant decrement of 0.094.St. 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.

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.

Non-Linear Harmonic flow simulations of a High-Head Francis Turbine test case

NASA Astrophysics Data System (ADS)

Lestriez, R.; Amet, E.; Tartinville, B.; Hirsch, C.

2016-11-01

This work investigates the use of the non-linear harmonic (NLH) method for a high- head Francis turbine, the Francis99 workshop test case. The NLH method relies on a Fourier decomposition of the unsteady flow components in harmonics of Blade Passing Frequencies (BPF), which are the fundamentals of the periodic disturbances generated by the adjacent blade rows. The unsteady flow solution is obtained by marching in pseudo-time to a steady-state solution of the transport equations associated with the time-mean, the BPFs and their harmonics. Thanks to this transposition into frequency domain, meshing only one blade channel is sufficient, like for a steady flow simulation. Notable benefits in terms of computing costs and engineering time can therefore be obtained compared to classical time marching approach using sliding grid techniques. The method has been applied for three operating points of the Francis99 workshop high-head Francis turbine. Steady and NLH flow simulations have been carried out for these configurations. Impact of the grid size and near-wall refinement is analysed on all operating points for steady simulations and for Best Efficiency Point (BEP) for NLH simulations. Then, NLH results for a selected grid size are compared for the three different operating points, reproducing the tendencies observed in the experiment.

Effect of wake structure on blade-vortex interaction phenomena: Acoustic prediction and validation

NASA Technical Reports Server (NTRS)

Gallman, Judith M.; Tung, Chee; Schultz, Klaus J.; Splettstoesser, Wolf; Buchholz, Heino

1995-01-01

During the Higher Harmonic Control Aeroacoustic Rotor Test, extensive measurements of the rotor aerodynamics, the far-field acoustics, the wake geometry, and the blade motion for powered, descent, flight conditions were made. These measurements have been used to validate and improve the prediction of blade-vortex interaction (BVI) noise. The improvements made to the BVI modeling after the evaluation of the test data are discussed. The effects of these improvements on the acoustic-pressure predictions are shown. These improvements include restructuring the wake, modifying the core size, incorporating the measured blade motion into the calculations, and attempting to improve the dynamic blade response. A comparison of four different implementations of the Ffowcs Williams and Hawkings equation is presented. A common set of aerodynamic input has been used for this comparison.

Viscous Three-Dimensional Simulation of Flow in an Axial Low Pressure Compressor at Engine Icing Operating Points

NASA Technical Reports Server (NTRS)

Rigby, David L.; Ameri, Ali A.; Veres, Joe; Jorgenson, Philip C. E.

2017-01-01

Viscous three-dimensional simulations of the Honeywell ALF502R-5 low pressure compressor (sometimes called a booster) using the NASA Glenn code GlennHT have been carried out. A total of ten simulations were produced. Five operating points are investigated, with each point run with two different wall thermal conditions. These operating points are at, or near, points where engine icing has been determined to be likely. In the future, the results of this study will be used for further analysis such as predicting collection efficiency of ice particles and ice growth rates at various locations in the compressor. A mixing plane boundary condition is used between each blade row, resulting in convergence to steady state within each blade row. The k-omega turbulence model of Wilcox, combined with viscous grid spacing near the wall on the order of one, is used to resolve the turbulent boundary layers. For each of the operating points, heat transfer coefficients are generated on the blades and walls. The heat transfer coefficients are produced by running the operating point with two different wall thermal conditions and then solving simultaneously for the heat transfer coefficient and adiabatic wall temperature at each point. Average Nusselt numbers are calculated for the most relevant surfaces. The values are seen to scale with Reynolds number to approximately a power of 0.7. Additionally, images of surface distribution of Nusselt number are presented. Qualitative comparison between the five operating points show that there is relatively little change in the character of the distribution. The dominant observed effect is that of an overall scaling, which is expected due to Reynolds number differences. One interesting aspect about the Nusselt number distribution is observed on the casing (outer diameter) downstream of the exit guide vanes (EGVs). The Nusselt number is relatively high between the pairs of EGVs, with two lower troughs downstream of each EGV trailing edge. This is of particular interest since rather complex ice shapes have been observed in that region.

Study on rotational frequency noise in a centrifugal compressor for automobile turbochargers

NASA Astrophysics Data System (ADS)

Wakaki, Daichi; Sakuka, Yuta; Yamasaki, Nobuhiko; Yamagata, Akihiro

2014-02-01

The rotational frequency noise (also known as the pulsation noise) due to the mistuning of impeller blade rows introduced at the manufacturing stage of the impellers is observed in the small-sized centrifugal compressor for automobile turbochargers. The present paper addresses the elucidation of the generating mechanism and parameter dependency such as the kind and degree of mistuning. In order to analyze numerically the rotational frequency noise due to mistuning, the unsteady computational fluid dynamics (CFD) of the whole compressor including volute is executed, and the resultant time history of the pressure is fed into the spectral analysis.

Icing Branch Current Research Activities in Icing Physics

NASA Technical Reports Server (NTRS)

Vargas, Mario

2009-01-01

Current development: A grid block transformation scheme which allows the input of grids in arbitrary reference frames, the use of mirror planes, and grids with relative velocities has been developed. A simple ice crystal and sand particle bouncing scheme has been included. Added an SLD splashing model based on that developed by William Wright for the LEWICE 3.2.2 software. A new area based collection efficiency algorithm will be incorporated which calculates trajectories from inflow block boundaries to outflow block boundaries. This method will be used for calculating and passing collection efficiency data between blade rows for turbo-machinery calculations.

Full 3D Analysis of the GE90 Turbofan Primary Flowpath

NASA Technical Reports Server (NTRS)

Turner, Mark G.

2000-01-01

The multistage simulations of the GE90 turbofan primary flowpath components have been performed. The multistage CFD code, APNASA, has been used to analyze the fan, fan OGV and booster, the 10-stage high-pressure compressor and the entire turbine system of the GE90 turbofan engine. The code has two levels of parallel, and for the 18 blade row full turbine simulation has 87.3 percent parallel efficiency with 121 processors on an SGI ORIGIN. Grid generation is accomplished with the multistage Average Passage Grid Generator, APG. Results for each component are shown which compare favorably with test data.

Fluid-Structure interaction analysis and performance evaluation of a membrane blade

NASA Astrophysics Data System (ADS)

Saeedi, M.; Wüchner, R.; Bletzinger, K.-U.

2016-09-01

Examining the potential of a membrane blade concept is the goal of the current work. In the sailwing concept the surface of the wing, or the blade in this case, is made from pre-tensioned membranes which meet at the pre-tensioned edge cable at the trailing edge. Because of the dependency between membrane deformation and applied aerodynamic load, two-way coupled fluid-structure interaction analysis is necessary for evaluation of the aerodynamic performance of such a configuration. The in-house finite element based structural solver, CARAT++, is coupled with OpenFOAM in order to tackle the multi-physics problem. The main aerodynamic characteristics of the membrane blade including lift coefficient, drag coefficient and lift to drag ratio are compared with its rigid counterpart. A single non-rotating NREL phase VI blade is studied here as a first step towards analyzing the concept for the rotating case. Compared with the rigid blade, the membrane blade has a higher slope of the lift curve. For higher angles of attack, lift and drag coefficients as well as the lift to drag ratio is higher for the membrane blade. A single non-rotating blade is studied here as a first step towards analyzing the concept for the rotating case.

Evaluation of feasibility of prestressed concrete for use in wind turbine blades

NASA Technical Reports Server (NTRS)

Leiblein, S.; Londahl, D. S.; Furlong, D. B.; Dreier, M. E.

1979-01-01

A preliminary evaluation of the feasibility of the use of prestressed concrete as a material for low cost blades for wind turbines was conducted. A baseline blade design was achieved for an experimental wind turbine that met aerodynamic and structural requirements. Significant cost reductions were indicated for volume production. Casting of a model blade section showed no fabrication problems. Coupled dynamic analysis revealed that adverse rotor tower interactions can be significant with heavy rotor blades.

Non-Synchronous Vibration of Turbomachinery Airfoils

DTIC Science & Technology

2006-03-01

study and prevention of non-synchronous vibrations. Non-synchronous vibrations in turbine engine blades are the result of the interaction of an...was a modern fan vane blade known as the H2 case. This blade encountered NSV in experimental rig testing. An analysis was performed with TURBO ...design stage for flow over turbine engine blades . REFERENCES Anagnostopoulos, P., ed. Flow-Induced Vibrations in Engineering

Tracking Blade Tip Vortices for Numerical Flow Simulations of Hovering Rotorcraft

NASA Technical Reports Server (NTRS)

Kao, David L.

2016-01-01

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

The effects of free stream turbulence on the flow field through a compressor cascade

NASA Astrophysics Data System (ADS)

Muthanna Kolera, Chittiappa

The flow through a compressor cascade with tip leakage has been studied experimentally. The cascade of GE rotor B section blades had an inlet angle of 65.1°, a stagger angle of 56.9°, and a solidity of 1.08. The final turning angle of the cascade was 11.8°. This compressor configuration was representative of the core compressor of an aircraft engine. The cascade was operated with a tip gap of 1.65%, and operated at a Reynolds number based on the chord length (0.254 m) of 388,000. Measurements were made at 8 axial locations to reveal the structure of the flow as it evolved through the cascade. Measurements were also made to reveal the effects of grid generated turbulence on this flow. The data set is unique in that not only does it give a comparison of elevated free stream turbulence effects, but also documents the developing flow through the blade row of a compressor cascade with tip leakage. Measurements were made at a total of 8 locations 0.8, 0.23 axial chords upstream and 0, 0.27, 0.48, 0.77, 0.98, and 1.26 axial chords downstream of the leading edge of the blade row for both inflow turbulence cases. The measurements revealed the formation and development of the tip leakage vortex within the passage. The tip leakage vortex becomes apparent at approximately X/ca = 0.27 and dominated much of the endwall flow. The tip leakage vortex is characterized by high streamwise velocity deficits, high vorticity and high turbulence kinetic energy levels. The result showed that between 0.77 and 0.98 axial chords downstream of the leading edge, the vortex structure and behavior changes. The effects of grid generated turbulence were also documented. The results revealed significant effects on the flow field. The results showed a 4% decrease in the blade loading and a 20% reduction in the vorticity levels within tip leakage vortex. There was also a shift in the vortex path, showing a shift close to the suction side with grid generated turbulence, indicating the strength of the vortex was decreased. Circulation calculations showed this reduction, and also indicated that the tip leakage vortex increased in size by about 30%. The results revealed that overall, the turbulence kinetic energy levels in the tip leakage vortex were increased, with the most drastic change occurring at X/ca = 0.77.

The Study the Vibration Condition of the Blade of the Gas Turbine Engine with an All-metal Wire Rope Damper in the Area Mount of the Blade to the Disk

NASA Astrophysics Data System (ADS)

Melentjev, Vladimir S.; Gvozdev, Alexander S.

2018-01-01

Improving the reliability of modern turbine engines is actual task. This is achieved due to prevent a vibration damage of the operating blades. On the department of structure and design of aircraft engines have accumulated a lot of experimental data on the protection of the blades of the gas turbine engine from a vibration. In this paper we proposed a method for calculating the characteristics of wire rope dampers in the root attachment of blade of a gas turbine engine. The method is based on the use of the finite element method and transient analysis. Contact interaction (Lagrange-Euler method) between the compressor blade and the disc of the rotor has been taken into account. Contribution of contact interaction between details in damping of the system was measured. The proposed method provides a convenient way for the iterative selection of the required parameters the wire rope elastic-damping element. This element is able to provide the necessary protection from the vibration for the blade of a gas turbine engine.

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.

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.

Acoustic measurements from a rotor blade-vortex interaction noise experiment in the German-Dutch Wind Tunnel (DNW)

NASA Technical Reports Server (NTRS)

Martin, Ruth M.; Splettstoesser, W. R.; Elliott, J. W.; Schultz, K.-J.

1988-01-01

Acoustic data are presented from a 40 percent scale model of the 4-bladed BO-105 helicopter main rotor, measured in the large European aeroacoustic wind tunnel, the DNW. Rotor blade-vortex interaction (BVI) noise data in the low speed flight range were acquired using a traversing in-flow microphone array. The experimental apparatus, testing procedures, calibration results, and experimental objectives are fully described. A large representative set of averaged acoustic signals is presented.

Developments in blade shape design for a Darrieus vertical axis wind turbine

NASA Astrophysics Data System (ADS)

Ashwill, T. D.; Leonard, T. M.

1986-09-01

A new computer program package has been developed that determines the troposkein shape for a Darrieus Vertical Axis Wind Turbine Blade with any geometrical configuration or rotation rate. This package allows users to interact and develop a buildable blade whose shape closely approximates the troposkein. Use of this package can significantly reduce flatwise mean bending stresses in the blade and increase fatigue life.

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

NASA Technical Reports Server (NTRS)

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

1998-01-01

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

PREDICTING TURBINE STAGE PERFORMANCE

NASA Technical Reports Server (NTRS)

Boyle, R. J.

1994-01-01

This program was developed to predict turbine stage performance taking into account the effects of complex passage geometries. The method uses a quasi-3D inviscid-flow analysis iteratively coupled to calculated losses so that changes in losses result in changes in the flow distribution. In this manner the effects of both the geometry on the flow distribution and the flow distribution on losses are accounted for. The flow may be subsonic or shock-free transonic. The blade row may be fixed or rotating, and the blades may be twisted and leaned. This program has been applied to axial and radial turbines, and is helpful in the analysis of mixed flow machines. This program is a combination of the flow analysis programs MERIDL and TSONIC coupled to the boundary layer program BLAYER. The subsonic flow solution is obtained by a finite difference, stream function analysis. Transonic blade-to-blade solutions are obtained using information from the finite difference, stream function solution with a reduced flow factor. Upstream and downstream flow variables may vary from hub to shroud and provision is made to correct for loss of stagnation pressure. Boundary layer analyses are made to determine profile and end-wall friction losses. Empirical loss models are used to account for incidence, secondary flow, disc windage, and clearance losses. The total losses are then used to calculate stator, rotor, and stage efficiency. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 370/3033 under TSS with a central memory requirement of approximately 4.5 Megs of 8 bit bytes. This program was developed in 1985.

Design and fabrication of forward-swept counterrotation blade configuration for wind tunnel testing

NASA Technical Reports Server (NTRS)

Nichols, G. H.

1994-01-01

Work performed by GE Aircraft on advanced counterrotation blade configuration concepts for high speed turboprop system is described. Primary emphasis was placed on theoretically and experimentally evaluating the aerodynamic, aeromechanical, and acoustic performance of GE-defined counterrotating blade concepts. Several blade design concepts were considered. Feasibility studies were conducted to evaluate a forward-swept versus an aft-swept blade application and how the given blade design would affect interaction between rotors. Two blade designs were initially selected. Both designs involved in-depth aerodynamic, aeromechanical, mechanical, and acoustic analyses followed by the fabrication of forward-swept, forward rotor blade sets to be wind tunnel tested with an aft-swept, aft rotor blade set. A third blade set was later produced from a NASA design that was based on wind tunnel test results from the first two blade sets. This blade set had a stiffer outer ply material added to the original blade design, in order to reach the design point operating line. Detailed analyses, feasibility studies, and fabrication procedures for all blade sets are presented.

Turbofan noise generation. Volume 1: Analysis

NASA Astrophysics Data System (ADS)

Ventres, C. S.; Theobald, M. A.; Mark, W. D.

1982-07-01

Computer programs were developed which calculate the in-duct acoustic modes excited by a fan/stator stae operating at subsonic tip speed. Three noise source mechanisms are included: (1) sound generated by the rotor blades interacting with turbulence ingested into, or generated within, the inlet duct; (2) sound generated by the stator vanes interacting with the turbulent wakes of the rotors blades; and (3) sound generated by the stator vanes interacting with the mean velocity deficit wakes of the rotor blades. The fan/stator stage is modeled as an ensemble of blades and vanes of zero camber and thickness enclosed within an infinite hard-walled annular duct. Turbulence drawn into or generated within the inlet duct is modeled as nonhomogeneous and anisotropic random fluid motion, superimposed upon a uniform axial mean flow, and convected with that flow. Equations for the duct mode amplitudes, or expected values of the amplitudes, are derived.

Turbofan noise generation. Volume 1: Analysis

NASA Technical Reports Server (NTRS)

Ventres, C. S.; Theobald, M. A.; Mark, W. D.

1982-01-01

Computer programs were developed which calculate the in-duct acoustic modes excited by a fan/stator stae operating at subsonic tip speed. Three noise source mechanisms are included: (1) sound generated by the rotor blades interacting with turbulence ingested into, or generated within, the inlet duct; (2) sound generated by the stator vanes interacting with the turbulent wakes of the rotors blades; and (3) sound generated by the stator vanes interacting with the mean velocity deficit wakes of the rotor blades. The fan/stator stage is modeled as an ensemble of blades and vanes of zero camber and thickness enclosed within an infinite hard-walled annular duct. Turbulence drawn into or generated within the inlet duct is modeled as nonhomogeneous and anisotropic random fluid motion, superimposed upon a uniform axial mean flow, and convected with that flow. Equations for the duct mode amplitudes, or expected values of the amplitudes, are derived.

Endwall shape modification using vortex generators and fences to improve gas turbine cooling and effectiveness

NASA Astrophysics Data System (ADS)

Gokce, Zeki Ozgur

The gas turbine is one of the most important parts of the air-breathing jet engine. Hence, improving its efficiency and rendering it operable under high temperatures are constant goals for the aerospace industry. Two types of flow within the gas turbine are of critical relevance: The flow around the first row of stator blades (also known as the nozzle guide vane blade - NGV) and the cooling flow inside the turbine blade cooling channel. The subject of this thesis work was to search for methods that could improve the characteristics of these two types of flows, thus enabling superior engine performance. The innovative aspect of our work was to apply an endwall shape modification previously employed by non-aerospace industries for cooling applications, to the gas turbine cooling flow which is vital to aerospace propulsion. Since the costs of investigating the possible benefits of any idea via extensive experiments could be quite high, we decided to use computational fluid dynamics (CFD) followed by experimentation as our methodology. We decided to analyze the potential benefits of using vortex generators (VGs) as well as the rectangular endwall fence. Since the pin-fins used in cooling flow are circular cylinders, and since the boundary layer flow is mainly characterized by the leading edge diameter of the NGV blade, we modeled both the pin-fins and the NGV blade as vertical circular cylinders. The baseline case consisted of the cylinder(s) being subjected to cross flow and a certain amount of freestream turbulence. The modifications we made on the endwall consisted of rectangular fences. In the case of the cooling flow, we used triangular shaped, common flow up oriented, delta winglet type vortex generators as well as rectangular endwall fences. The channel contained singular cylinders as well as staggered rows of multiple cylinders. For the NGV flow, a rectangular endwall fence and a singular cylinder were utilized. Using extensive CFD modeling and analysis, we confirmed that placing a rectangular endwall fence upstream of the cylinder created additional turbulent mixing in the domain. This led to increased mixing of the cooler flow in the freestream and the hotter flow near the endwall. As a result, we showed that adding a rectangular fence created a 10% mean heat transfer increase downstream of the cylinder. When vortex generators are used, as the flow passes over the sharp edges of the vortex generators, it separates and continues downstream in a rolling, helical pattern. Combined with the effect generated by the orientation of the vortex generators, this flow structure mixes the higher momentum fluid in the freestream with lower momentum fluid in the boundary layer. Similar turbulent mixing behavior is observed over the entire domain, near the cylinders and the side walls. As a result, the heat transfer levels over the wall surfaces are increased and improved cooling is achieved. The improvements in heat transfer are obtained at the expense of acceptable pressure losses across the cooling channel. When the vortex generators are used, the CFD modeling studies showed that overall heat transfer improvements as high as 27% compared to the baseline case are observed inside a domain containing multiple rows of cylinders. A price in the form of 13% pressure loss increase across the channel is paid for the heat transfer benefits. Experiments conducted in the open loop wind tunnel of the Turbomachinery Aero-Heat Transfer Laboratory of the Department of Aerospace Engineering of Penn State University supported the general positive trend of these findings, with a 14% overall increase in heat transfer over the constant heat flux surface when vortex generators are installed, accompanied by an 8% increase in pressure loss. (Abstract shortened by UMI.)

Rotorcraft Brownout Advanced Understanding, Control, and Mitigation

DTIC Science & Technology

2014-10-31

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

A Novel Method for Reducing Rotor Blade-Vortex Interaction

NASA Technical Reports Server (NTRS)

Glinka, A. T.

2000-01-01

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

Intracortical pathways mediate nonlinear fast oscillation (>200 Hz) interactions within rat barrel cortex.

PubMed

Staba, Richard J; Ard, Tyler D; Benison, Alexander M; Barth, Daniel S

2005-05-01

Whisker evoked fast oscillations (FOs; >200 Hz) within the rodent posteromedial barrel subfield are thought to reflect very rapid integration of multiwhisker stimuli, yet the pathways mediating FO interactions remain unclear and may involve interactions within thalamus and/or cortex. In the present study using anesthetized rats, a cortical incision was made between sites representing the stimulated whiskers to determine how intracortical networks contributed to patterns of FOs. With cortex intact, simultaneous stimulation of a pair of whiskers aligned in a row evoked supralinear responses between sites separated by several millimeters. In contrast, stimulation of a nonadjacent pair of whiskers within an arc evoked FOs with no evidence for nonlinear interactions. However, stimulation of an adjacent pair of whiskers in an arc did evoke supralinear responses. After a cortical cut, supralinear interactions associated with FOs within a row were lost. These data indicate a distinct bias for stronger long-range connectivity that extends along barrel rows and that horizontal intracortical pathways exclusively mediate FO-related integration of tactile information.

Fluid-structure interaction modeling of wind turbines: simulating the full machine

NASA Astrophysics Data System (ADS)

Hsu, Ming-Chen; Bazilevs, Yuri

2012-12-01

In this paper we present our aerodynamics and fluid-structure interaction (FSI) computational techniques that enable dynamic, fully coupled, 3D FSI simulation of wind turbines at full scale, and in the presence of the nacelle and tower (i.e., simulation of the "full machine"). For the interaction of wind and flexible blades we employ a nonmatching interface discretization approach, where the aerodynamics is computed using a low-order finite-element-based ALE-VMS technique, while the rotor blades are modeled as thin composite shells discretized using NURBS-based isogeometric analysis (IGA). We find that coupling FEM and IGA in this manner gives a good combination of efficiency, accuracy, and flexibility of the computational procedures for wind turbine FSI. The interaction between the rotor and tower is handled using a non-overlapping sliding-interface approach, where both moving- and stationary-domain formulations of aerodynamics are employed. At the fluid-structure and sliding interfaces, the kinematic and traction continuity is enforced weakly, which is a key ingredient of the proposed numerical methodology. We present several simulations of a three-blade 5~MW wind turbine, with and without the tower. We find that, in the case of no tower, the presence of the sliding interface has no effect on the prediction of aerodynamic loads on the rotor. From this we conclude that weak enforcement of the kinematics gives just as accurate results as the strong enforcement, and thus enables the simulation of rotor-tower interaction (as well as other applications involving mechanical components in relative motion). We also find that the blade passing the tower produces a 10-12 % drop (per blade) in the aerodynamic torque. We feel this finding may be important when it comes to the fatigue-life analysis and prediction for wind turbine blades.

Wavy-Planform Helicopter Blades Make Less Noise

NASA Technical Reports Server (NTRS)

Brooks, Thomas F.

2004-01-01

Wavy-planform rotor blades for helicopters have been investigated for the first time in an effort to reduce noise. Two of the main sources of helicopter noise are blade/vortex interaction (BVI) and volume displacement. (The noise contributed by volume displacement is termed thickness noise.) The reduction in noise generated by a wavyplanform blade, relative to that generated by an otherwise equivalent straight-planform blade, affects both main sources: (1) the BVI noise is reduced through smoothing and defocusing of the aerodynamic loading on the blade and (2) the thickness noise is reduced by reducing gradients of thickness with respect to listeners on the ground.

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.

Theory for broadband Noise of Rotor and Stator Cascades with Inhomogeneous Inflow Turbulence Including Effects of Lean and Sweep

NASA Technical Reports Server (NTRS)

Hanson, Donald B.

2001-01-01

The problem of broadband noise generated by turbulence impinging on a downstream blade row is examined from a theoretical viewpoint. Equations are derived for sound power spectra in terms of 3 dimensional wavenumber spectra of the turbulence. Particular attention is given to issues of turbulence inhomogeneity associated with the near field of the rotor and variations through boundary layers. Lean and sweep of the rotor or stator cascade are also handled rigorously with a full derivation of the relevant geometry and definitions of lean and sweep angles. Use of the general theory is illustrated by 2 simple theoretical spectra for homogeneous turbulence. Limited comparisons are made with data from model fans designed by Pratt & Whitney, Allison, and Boeing. Parametric studies for stator noise are presented showing trends with Mach number, vane count, turbulence scale and intensity, lean, and sweep. Two conventions are presented to define lean and sweep. In the "cascade system" lean is a rotation out of its plane and sweep is a rotation of the airfoil in its plane. In the "duct system" lean is the leading edge angle viewing the fan from the front (along the fan axis) and sweep is the angle viewing the fan from the side (,perpendicular to the axis). It is shown that the governing parameter is sweep in the plane of the airfoil (which reduces the chordwise component of Mach number). Lean (out of the plane of the airfoil) has little effect. Rotor noise predictions are compared with duct turbulence/rotor interaction noise data from Boeing and variations, including blade tip sweep and turbulence axial and transverse scales are explored.

Forward sweep, low noise rotor blade

NASA Technical Reports Server (NTRS)

Brooks, Thomas F. (Inventor)

1994-01-01

A forward-swept, low-noise rotor blade includes an inboard section, an aft-swept section, and a forward-swept outboard section. The rotor blade reduces the noise of rotorcraft, including both standard helicopters and advanced systems such as tiltrotors. The primary noise reduction feature is the forward sweep of the planform over a large portion of the outer blade radius. The rotor blade also includes an aft-swept section. The purpose of the aft-swept region is to provide a partial balance to pitching moments produced by the outboard forward-swept portion of the blade. The noise source showing maximum noise reduction is blade-vortex interaction (BVI) noise. Also reduced are thickness, noise, high speed impulsive noise, cabin vibration, and loading noise.

F100(3) parallel compressor computer code and user's manual

NASA Technical Reports Server (NTRS)

Mazzawy, R. S.; Fulkerson, D. A.; Haddad, D. E.; Clark, T. A.

1978-01-01

The Pratt & Whitney Aircraft multiple segment parallel compressor model has been modified to include the influence of variable compressor vane geometry on the sensitivity to circumferential flow distortion. Further, performance characteristics of the F100 (3) compression system have been incorporated into the model on a blade row basis. In this modified form, the distortion's circumferential location is referenced relative to the variable vane controlling sensors of the F100 (3) engine so that the proper solution can be obtained regardless of distortion orientation. This feature is particularly important for the analysis of inlet temperature distortion. Compatibility with fixed geometry compressor applications has been maintained in the model.

Effects of geometric variables on rub characteristics of Ti-6Al-4V

NASA Technical Reports Server (NTRS)

Bill, R. C.; Wolak, J.; Wisander, D. W.

1981-01-01

Experiments simulating rub interactions between Ti-6Al-4V blade tips and various seal materials were conducted. The number of blade tips and the blade tip geometry were varied to determine their effects on rub forces and on wear phenomena. Contact was found to be quite unsteady for all blade tip geometries except for those incorporating deliberately rounded blade tips. The unsteady contact was characterized by long periods of rubbing contact and increasing blade tip that terminated in sudden rapid metal removal, sometimes accompanied by tearing and disruption of porous seal material under the rub surface. A model describing the blade tip loading is proposed and is based on the propagation of an elastic stress wave through the seal material as the seal material is dynamically compressed by the blade tip leading edge.

Acoustic and aerodynamic study of a pusher-propeller aircraft model

NASA Astrophysics Data System (ADS)

Soderman, Paul T.; Horne, W. Clifton

1990-09-01

An aerodynamic and acoustic study was made of a pusher-propeller aircraft model in the NASA-Ames 7 x 10 ft Wind Tunnel. The test section was changed to operate as an open jet. The 591 mm diameter unswept propeller was operated alone and in the wake of three empennages: an I tail, Y tail, and a V tail. The radiated noise and detailed wake properties were measured. Results indicate that the unsteady blade loading caused by the blade interactions with the wake mean velocity distribution had a strong effect on the harmonics of blade passage noise. The blade passage harmonics above the first were substantially increased in all horizontal directions by the empennage/propeller interaction. Directivity in the plane of the propeller was maximum perpendicular to the blade surface. Increasing the tail loading caused the propeller harmonics to increase 3 to 5 dB for an empennage/propeller spacing of 0.38 mean empennage chords. The interaction noise became weak as empennage propeller spacing was increased beyond 1.0 mean empennage chord lengths. Unlike the mean wake deficit, the wake turbulence had only a small effect on the propeller noise, that effect being a small increase in the broadband noise.

Acoustic and aerodynamic study of a pusher-propeller aircraft model

NASA Technical Reports Server (NTRS)

Soderman, Paul T.; Horne, W. Clifton

1990-01-01

An aerodynamic and acoustic study was made of a pusher-propeller aircraft model in the NASA-Ames 7 x 10 ft Wind Tunnel. The test section was changed to operate as an open jet. The 591 mm diameter unswept propeller was operated alone and in the wake of three empennages: an I tail, Y tail, and a V tail. The radiated noise and detailed wake properties were measured. Results indicate that the unsteady blade loading caused by the blade interactions with the wake mean velocity distribution had a strong effect on the harmonics of blade passage noise. The blade passage harmonics above the first were substantially increased in all horizontal directions by the empennage/propeller interaction. Directivity in the plane of the propeller was maximum perpendicular to the blade surface. Increasing the tail loading caused the propeller harmonics to increase 3 to 5 dB for an empennage/propeller spacing of 0.38 mean empennage chords. The interaction noise became weak as empennage propeller spacing was increased beyond 1.0 mean empennage chord lengths. Unlike the mean wake deficit, the wake turbulence had only a small effect on the propeller noise, that effect being a small increase in the broadband noise.

Active Control of Blade Tonals in Underwater Vehicles

DTIC Science & Technology

2006-12-01

Because the stator is a streamlined shape the wake deficit responsible for blade tonal noise is due mainly to surface drag, which can be thought of as a... wake deficit , the vortex rollup at this stage is not very repeatable. Therefore, this type of wake may not be the best suited for controlling blade ...sinusoidal and non-sinusoidal move profiles. This model was also able to capture the baseline wake deficit measured. 2-dimensional blade interaction was

Structural analysis of high-rpm composite propfan blades for a cruise missile wind tunnel model

NASA Technical Reports Server (NTRS)

Carek, David A.

1993-01-01

Analyses were performed on a high-speed composite blade set for the Department of Defense Propfan Missile Interactions Project. The final design iteration, which resulted in the CM2D-2 blade design, is described in this report. Mode shapes, integral order excitation, and stress margins were examined. In addition, geometric corrections were performed to compensate for blade deflection under operating conditions with respect to the aerodynamic design shape.

Numerical simulation of compressor endwall and casing treatment flow phenomena

NASA Technical Reports Server (NTRS)

Crook, A. J.; Greitzer, E. M.; Tan, C. S.; Adamczyk, J. J.

1992-01-01

A numerical study is presented of the flow in the endwall region of a compressor blade row, in conditions of operation with both smooth and grooved endwalls. The computations are first compared to velocity field measurements in a cantilevered stator/rotating hub configuration to confirm that the salient features are captured. Computations are then interrogated to examine the tip leakage flow structure since this is a dominant feature of the endwall region. In particular, the high blockage that can exist near the endwalls at the rear of a compressor blade passage appears to be directly linked to low total pressure fluid associated with the leakage flow. The fluid dynamic action of the grooved endwall, representative of the casing treatments that have been most successful in suppressing stall, is then simulated computationally and two principal effects are identified. One is suction of the low total pressure, high blockage fluid at the rear of the passage. The second is energizing of the tip leakage flow, most notably in the core of the leakage vortex, thereby suppressing the blockage at its source.

Assessment of Aerodynamic Challenges of a Variable-Speed Power Turbine for Large Civil Tilt-Rotor Application

NASA Technical Reports Server (NTRS)

Welch, Gerand E.

2010-01-01

The main rotors of the NASA Large Civil Tilt-Rotor notional vehicle operate over a wide speed-range (100% at take-off to 54% at cruise). The variable-speed power turbine, when coupled to a fixed-gear-ratio transmission, offers one approach to accomplish this speed variation. The key aero-challenges of the variable-speed power turbine are related to high work factors at cruise, where the power turbine operates at 54% of take-off speed, wide incidence variations into the vane, blade, and exit-guide-vane rows associated with the power-turbine speed change, and the impact of low aft-stage Reynolds number (transitional flow) at 28 kft cruise. Meanline and 2-D Reynolds-Averaged Navier- Stokes analyses are used to characterize the variable-speed power-turbine aerodynamic challenges and to outline a conceptual design approach that accounts for multi-point operation. Identified technical challenges associated with the aerodynamics of high work factor, incidence-tolerant blading, and low Reynolds numbers pose research needs outlined in the paper

Functional constraints on tooth morphology in carnivorous mammals

PubMed Central

2012-01-01

Background The range of potential morphologies resulting from evolution is limited by complex interacting processes, ranging from development to function. Quantifying these interactions is important for understanding adaptation and convergent evolution. Using three-dimensional reconstructions of carnivoran and dasyuromorph tooth rows, we compared statistical models of the relationship between tooth row shape and the opposing tooth row, a static feature, as well as measures of mandibular motion during chewing (occlusion), which are kinetic features. This is a new approach to quantifying functional integration because we use measures of movement and displacement, such as the amount the mandible translates laterally during occlusion, as opposed to conventional morphological measures, such as mandible length and geometric landmarks. By sampling two distantly related groups of ecologically similar mammals, we study carnivorous mammals in general rather than a specific group of mammals. Results Statistical model comparisons demonstrate that the best performing models always include some measure of mandibular motion, indicating that functional and statistical models of tooth shape as purely a function of the opposing tooth row are too simple and that increased model complexity provides a better understanding of tooth form. The predictors of the best performing models always included the opposing tooth row shape and a relative linear measure of mandibular motion. Conclusions Our results provide quantitative support of long-standing hypotheses of tooth row shape as being influenced by mandibular motion in addition to the opposing tooth row. Additionally, this study illustrates the utility and necessity of including kinetic features in analyses of morphological integration. PMID:22899809

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

NASA Technical Reports Server (NTRS)

Tweedt, Daniel L.

2014-01-01

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

Phenomenological modeling of abradable wear in turbomachines

NASA Astrophysics Data System (ADS)

Berthoul, Bérenger; Batailly, Alain; Stainier, Laurent; Legrand, Mathias; Cartraud, Patrice

2018-01-01

Abradable materials are widely used as coatings within compressor and turbine stages of modern aircraft engines in order to reduce operating blade-tip/casing clearances and thus maximize energy efficiency. However, rubbing occurrences between blade tips and coating liners may lead to high blade vibratory levels and endanger their structural integrity through fatigue mechanisms. Accordingly, there is a need for a better comprehension of the physical phenomena at play and for an accurate modeling of the interaction, in order to predict potentially unsafe events. To this end, this work introduces a phenomenological model of the abradable coating removal based on phenomena reported in the literature and accounting for key frictional and wear mechanisms including plasticity at junctions, ploughing, micro-rupture and machining. It is implemented within an in-house software solution dedicated to the prediction of full three-dimensional blade/abradable coating interactions within an aircraft engine low pressure compressor. Two case studies are considered. The first one compares the results of an experimental abradable test rig and its simulation. The second one deals with the simulation of interactions in a complete low-pressure compressor. The consistency of the model with experimental observations is underlined, and the impact of material parameter variations on the interaction and wear behavior of the blade is discussed. It is found that even though wear patterns are remarkably robust, results are significantly influenced by abradable coating material properties.

New techniques for experimental generation of two-dimensional blade-vortex interaction at low Reynolds numbers

NASA Technical Reports Server (NTRS)

Booth, E., Jr.; Yu, J. C.

1986-01-01

An experimental investigation of two dimensional blade vortex interaction was held at NASA Langley Research Center. The first phase was a flow visualization study to document the approach process of a two dimensional vortex as it encountered a loaded blade model. To accomplish the flow visualization study, a method for generating two dimensional vortex filaments was required. The numerical study used to define a new vortex generation process and the use of this process in the flow visualization study were documented. Additionally, photographic techniques and data analysis methods used in the flow visualization study are examined.

Experimental study of secondary flow in the presence of two jet arrays on the wall of a turbine blade

NASA Astrophysics Data System (ADS)

Seddini, Abdelali

A experimental study of the interaction of the secondary turbine flow with the cooling jets injected at the base of the blades is presented. Subsonic wind tunnel tests were carried on using a gas turbine wheel. The hot wire sensor and the five orifices pressure transducer used in the tests are described. The results allow a satisfactory description of the evolution of the jets in the space between blades and give some indications of the interactions of jets with secondary flows, passing vortices and between themselves.

Blades Forced Vibration Under Aero-Elastic Excitation Modeled by Van der Pol

NASA Astrophysics Data System (ADS)

Pust, Ladislav; Pesek, Ludek

This paper employs a new analytical approach to model the influence of aerodynamic excitation on the dynamics of a bladed cascade at the flutter state. The flutter is an aero-elastic phenomenon that is linked to the interaction of the flow and the traveling deformation wave in the cascade when only the damping of the cascade changes. As a case study the dynamic properties of the five-blade-bunch excited by the running harmonic external forces and aerodynamic self-excited forces are investigated. This blade-bunch is linked in the shroud by means of the viscous-elastic damping elements. The external running excitation depends on the ratio of stator and rotor blade numbers and corresponds to the real type of excitation in the steam turbine. The aerodynamic self-excited forces are modeled by two types of Van der Pol nonlinear models. The influence of the interaction of both types of self-excitation with the external running excitation is investigated on the response curves.

Programmable Aperture with MEMS Microshutter Arrays

NASA Technical Reports Server (NTRS)

Moseley, Samuel; Li, Mary; Kutyrev, Alexander; Kletetschka, Gunther; Fettig, Rainer

2011-01-01

A microshutter array (MSA) has been developed for use as an aperture array for multi-object selections in James Webb Space Telescope (JWST) technology. Light shields, molybdenum nitride (MoN) coating on shutters, and aluminum/aluminum oxide coatings on interior walls are put on each shutter for light leak prevention, and to enhance optical contrast. Individual shutters are patterned with a torsion flexure that permits shutters to open 90 deg. with a minimized mechanical stress concentration. The shutters are actuated magnetically, latched, and addressed electrostatically. Also, micromechanical features are tailored onto individual shutters to prevent stiction. An individual shutter consists of a torsion hinge, a shutter blade, a front electrode that is coated on the shutter blade, a backside electrode that is coated on the interior walls, and a magnetic cobalt-iron coating. The magnetic coating is patterned into stripes on microshutters so that shutters can respond to an external magnetic field for the magnetic actuation. A set of column electrodes is placed on top of shutters, and a set of row electrodes on sidewalls is underneath the shutters so that they can be electrostatically latched open. A linear permanent magnet is aligned with the shutter rows and is positioned above a flipped upside-down array, and sweeps across the array in a direction parallel to shutter columns. As the magnet sweeps across the array, sequential rows of shutters are rotated from their natural horizontal orientation to a vertical open position, where they approach vertical electrodes on the sidewalls. When the electrodes are biased with a sufficient electrostatic force to overcome the mechanical restoring force of torsion bars, shutters remain latched to vertical electrodes in their open state. When the bias is removed, or is insufficient, the shutters return to their horizontal, closed positions. To release a shutter, both the electrode on the shutter and the one on the back wall where the shutter sits are grounded. The shutters with one or both ungrounded electrodes are held open. Sub-micron bumps underneath light shields and silicon ribs on back walls are the two features to prevent stiction. These features ensure that the microshutter array functions properly in mechanical motions. The MSA technology can be used primarily in multi-object imaging and spectroscopy, photomask generation, light switches, and in the stepper equipment used to make integrated circuits and MEMS (microelectromechanical systems) devices.

Three-Dimensional Unsteady Simulation of Aerodynamics and Heat Transfer in a Modern High Pressure Turbine Stage

NASA Technical Reports Server (NTRS)

Shyam, Vikram; Ameri, Ali

2009-01-01

Unsteady 3-D RANS simulations have been performed on a highly loaded transonic turbine stage and results are compared to steady calculations as well as to experiment. A low Reynolds number k-epsilon turbulence model is employed to provide closure for the RANS system. A phase-lag boundary condition is used in the tangential direction. This allows the unsteady simulation to be performed by using only one blade from each of the two rows. The objective of this work is to study the effect of unsteadiness on rotor heat transfer and to glean any insight into unsteady flow physics. The role of the stator wake passing on the pressure distribution at the leading edge is also studied. The simulated heat transfer and pressure results agreed favorably with experiment. The time-averaged heat transfer predicted by the unsteady simulation is higher than the heat transfer predicted by the steady simulation everywhere except at the leading edge. The shock structure formed due to stator-rotor interaction was analyzed. Heat transfer and pressure at the hub and casing were also studied. Thermal segregation was observed that leads to the heat transfer patterns predicted by steady and unsteady simulations to be different.

Pre-Stall Behavior of a Transonic Axial Compressor Stage via Time-Accurate Numerical Simulation

NASA Technical Reports Server (NTRS)

Chen, Jen-Ping; Hathaway, Michael D.; Herrick, Gregory P.

2008-01-01

CFD calculations using high-performance parallel computing were conducted to simulate the pre-stall flow of a transonic compressor stage, NASA compressor Stage 35. The simulations were run with a full-annulus grid that models the 3D, viscous, unsteady blade row interaction without the need for an artificial inlet distortion to induce stall. The simulation demonstrates the development of the rotating stall from the growth of instabilities. Pressure-rise performance and pressure traces are compared with published experimental data before the study of flow evolution prior to the rotating stall. Spatial FFT analysis of the flow indicates a rotating long-length disturbance of one rotor circumference, which is followed by a spike-type breakdown. The analysis also links the long-length wave disturbance with the initiation of the spike inception. The spike instabilities occur when the trajectory of the tip clearance flow becomes perpendicular to the axial direction. When approaching stall, the passage shock changes from a single oblique shock to a dual-shock, which distorts the perpendicular trajectory of the tip clearance vortex but shows no evidence of flow separation that may contribute to stall.

Relevance of aerodynamic modelling for load reduction control strategies of two-bladed wind turbines

NASA Astrophysics Data System (ADS)

Luhmann, B.; Cheng, P. W.

2014-06-01

A new load reduction concept is being developed for the two-bladed prototype of the Skywind 3.5MW wind turbine. Due to transport and installation advantages both offshore and in complex terrain two-bladed turbine designs are potentially more cost-effective than comparable three-bladed configurations. A disadvantage of two-bladed wind turbines is the increased fatigue loading, which is a result of asymmetrically distributed rotor forces. The innovative load reduction concept of the Skywind prototype consists of a combination of cyclic pitch control and tumbling rotor kinematics to mitigate periodic structural loading. Aerodynamic design tools must be able to model correctly the advanced dynamics of the rotor. In this paper the impact of the aerodynamic modelling approach is investigated for critical operational modes of a two-bladed wind turbine. Using a lifting line free wake vortex code (FVM) the physical limitations of the classical blade element momentum theory (BEM) can be evaluated. During regular operation vertical shear and yawed inflow are the main contributors to periodic blade load asymmetry. It is shown that the near wake interaction of the blades under such conditions is not fully captured by the correction models of BEM approach. The differing prediction of local induction causes a high fatigue load uncertainty especially for two-bladed turbines. The implementation of both cyclic pitch control and a tumbling rotor can mitigate the fatigue loading by increasing the aerodynamic and structural damping. The influence of the time and space variant vorticity distribution in the near wake is evaluated in detail for different cyclic pitch control functions and tumble dynamics respectively. It is demonstrated that dynamic inflow as well as wake blade interaction have a significant impact on the calculated blade forces and need to be accounted for by the aerodynamic modelling approach. Aeroelastic simulations are carried out using the high fidelity multi body simulation software SIMPACK. The aerodynamic loads are calculated using ECN's AeroModule and NREL's BEM code Aerodynl3.

Positive ions of the first- and second-row transition metal hydrides

NASA Technical Reports Server (NTRS)

Pettersson, Lars G. M.; Bauschlicher, Charles W., Jr.; Langhoff, Stephen R.; Partridge, Harry

1987-01-01

Theoretical dissociation energies for the first- and second-row transition metal hydride positive ions are critically compared against recent experimental values obtained from ion beam reactive scattering methods. Theoretical spectroscopic parameters and dipole moments are presented for the ground and several low-lying excited states. The calculations employ large Gaussian basis sets and account for electron correlation using the single-reference single- and double-excitation configuration interaction and coupled-pair-functional methods. The Darwin and mass-velocity contributions to the relativistic energy are included in the all-electron calculations on the first-row systems using first-order perturbation theory, and in the second-row systems using the Hay and Wadt relativistic effective core potentials. The theoretical D(0) values for the second-row transition metal hydride positive ions should provide a critical measure of the experimental values, which are not as refined as many of those in the first transition row.

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.

Blade-mounted trailing edge flap control for BVI noise reduction

NASA Technical Reports Server (NTRS)

Hassan, A. A.; Charles, B. D.; Tadghighi, H.; Sankar, L. N.

1992-01-01

Numerical procedures based on the 2-D and 3-D full potential equations and the 2-D Navier-Stokes equations were developed to study the effects of leading and trailing edge flap motions on the aerodynamics of parallel airfoil-vortex interactions and on the aerodynamics and acoustics of the more general self-generated rotor blade vortex interactions (BVI). For subcritical interactions, the 2-D results indicate that the trailing edge flap can be used to alleviate the impulsive loads experienced by the airfoil. For supercritical interactions, the results show the necessity of using a leading edge flap, rather than a trailing edge flap, to alleviate the interaction. Results for various time dependent flap motions and their effect on the predicted temporal sectional loads, differential pressures, and the free vortex trajectories are presented. For the OLS model rotor, contours of a BVI noise metric were used to quantify the effects of the trailing edge flap on the size and directivity of the high/low intensity noise region(s). Average reductions in the BVI noise levels on the order of 5 dB with moderate power penalties on the order of 18 pct. for a four bladed rotor and 58 pct. for a two bladed rotor were obtained.

A Combined High and Low Cycle Fatigue Model for Life Prediction of Turbine Blades

PubMed Central

Yue, Peng; Yu, Zheng-Yong; Wang, Qingyuan

2017-01-01

Combined high and low cycle fatigue (CCF) generally induces the failure of aircraft gas turbine attachments. Based on the aero-engine load spectrum, accurate assessment of fatigue damage due to the interaction of high cycle fatigue (HCF) resulting from high frequency vibrations and low cycle fatigue (LCF) from ground-air-ground engine cycles is of critical importance for ensuring structural integrity of engine components, like turbine blades. In this paper, the influence of combined damage accumulation on the expected CCF life are investigated for turbine blades. The CCF behavior of a turbine blade is usually studied by testing with four load-controlled parameters, including high cycle stress amplitude and frequency, and low cycle stress amplitude and frequency. According to this, a new damage accumulation model is proposed based on Miner’s rule to consider the coupled damage due to HCF-LCF interaction by introducing the four load parameters. Five experimental datasets of turbine blade alloys and turbine blades were introduced for model validation and comparison between the proposed Miner, Manson-Halford, and Trufyakov-Kovalchuk models. Results show that the proposed model provides more accurate predictions than others with lower mean and standard deviation values of model prediction errors. PMID:28773064

Impeller tandem blade study with grid embedding for local grid refinement

NASA Technical Reports Server (NTRS)

Bache, George

1992-01-01

Flow non-uniformity at the discharge of high power density impellers can result in significant unsteady interactions between impeller blades and downstream diffuser vanes. These interactions result in degradation of both performance and pump reliability. The MSFC Pump Technology Team has recognized the importance of resolving this problem and has thus initiated the development and testing of a high head coefficient impeller. One of the primary goals of this program is to improve impeller performance and discharge flow uniformity. The objective of the present work is complimentary. Flow uniformity and performance gains were sought through the application of a tandem blade arrangement. The approach adopted was to numerically establish flow characteristics at the impeller discharge for the baseline MSFC impeller and then parametrically evaluate tandem blade configurations. A tandem design was sought that improves both impeller performance and discharge uniformity. The Navier-Stokes solver AEROVISC was used to conduct the study. Grid embedding is used to resolve local gradients while attempting to minimize model size. Initial results indicate that significant gains in flow uniformity can be achieved through the tandem blade concept and that blade clocking rather than slot location is the primary driver for flow uniformity.

A Combined High and Low Cycle Fatigue Model for Life Prediction of Turbine Blades.

PubMed

Zhu, Shun-Peng; Yue, Peng; Yu, Zheng-Yong; Wang, Qingyuan

2017-06-26

Combined high and low cycle fatigue (CCF) generally induces the failure of aircraft gas turbine attachments. Based on the aero-engine load spectrum, accurate assessment of fatigue damage due to the interaction of high cycle fatigue (HCF) resulting from high frequency vibrations and low cycle fatigue (LCF) from ground-air-ground engine cycles is of critical importance for ensuring structural integrity of engine components, like turbine blades. In this paper, the influence of combined damage accumulation on the expected CCF life are investigated for turbine blades. The CCF behavior of a turbine blade is usually studied by testing with four load-controlled parameters, including high cycle stress amplitude and frequency, and low cycle stress amplitude and frequency. According to this, a new damage accumulation model is proposed based on Miner's rule to consider the coupled damage due to HCF-LCF interaction by introducing the four load parameters. Five experimental datasets of turbine blade alloys and turbine blades were introduced for model validation and comparison between the proposed Miner, Manson-Halford, and Trufyakov-Kovalchuk models. Results show that the proposed model provides more accurate predictions than others with lower mean and standard deviation values of model prediction errors.

Time-Varying Loads of Co-Axial Rotor Blade Crossings

NASA Technical Reports Server (NTRS)

Schatzman, Natasha L.; Komerath, Narayanan; Romander, Ethan A.

2017-01-01

The blade crossing event of a coaxial counter-rotating rotor is a potential source of noise and impulsive blade loads. Blade crossings occur many times during each rotor revolution. In previous research by the authors, this phenomenon was analyzed by simulating two airfoils passing each other at specified speeds and vertical separation distances, using the compressible Navier-Stokes solver OVERFLOW. The simulations explored mutual aerodynamic interactions associated with thickness, circulation, and compressibility effects. Results revealed the complex nature of the aerodynamic impulses generated by upperlower airfoil interactions. In this paper, the coaxial rotor system is simulated using two trains of airfoils, vertically offset, and traveling in opposite directions. The simulation represents multiple blade crossings in a rotor revolution by specifying horizontal distances between each airfoil in the train based on the circumferential distance between blade tips. The shed vorticity from prior crossing events will affect each pair of upperlower airfoils. The aerodynamic loads on the airfoil and flow field characteristics are computed before, at, and after each airfoil crossing. Results from the multiple-airfoil simulation show noticeable changes in the airfoil aerodynamics by introducing additional fluctuation in the aerodynamic time history.

Simulation of unsteady flows through stator and rotor blades of a gas turbine using the Chimera method

NASA Technical Reports Server (NTRS)

Nakamura, S.; Scott, J. N.

1993-01-01

A two-dimensional model to solve compressible Navier-Stokes equations for the flow through stator and rotor blades of a turbine is developed. The flow domains for the stator and rotor blades are coupled by the Chimera method that makes grid generation easy and enhances accuracy because the area of the grid that have high turning of grid lines or high skewness can be eliminated from the computational domain after the grids are generated. The results of flow computations show various important features of unsteady flows including the acoustic waves interacting with boundary layers, Karman vortex shedding from the trailing edge of the stator blades, pulsating incoming flow to a rotor blade from passing stator blades, and flow separation from both suction and pressure sides of the rotor blades.

Computational Aerodynamic Analysis of Offshore Upwind and Downwind Turbines

DOE PAGES

Zhao, Qiuying; Sheng, Chunhua; Afjeh, Abdollah

2014-01-01

Aerodynamic interactions of the model NREL 5 MW offshore horizontal axis wind turbines (HAWT) are investigated using a high-fidelity computational fluid dynamics (CFD) analysis. Four wind turbine configurations are considered; three-bladed upwind and downwind and two-bladed upwind and downwind configurations, which operate at two different rotor speeds of 12.1 and 16 RPM. In the present study, both steady and unsteady aerodynamic loads, such as the rotor torque, blade hub bending moment, and base the tower bending moment of the tower, are evaluated in detail to provide overall assessment of different wind turbine configurations. Aerodynamic interactions between the rotor and tower are analyzed,more » including the rotor wake development downstream. The computational analysis provides insight into aerodynamic performance of the upwind and downwind, two- and three-bladed horizontal axis wind turbines.« less

Unsteady aerodynamic interaction effects on turbomachinery blade life and performance

NASA Technical Reports Server (NTRS)

Adamczyk, John J.

1992-01-01

This paper is an attempt to address the impact of a class of unsteady flows on the life and performance of turbomachinery blading. These class of flows to be investigated are those whose characteristic frequency is an integral multiple of rotor shaft speed. Analysis of data recorded downstream of a compressor and turbine rotor will reveal that this class of flows can be highly three-dimensional and may lead to the generation of secondary flows within downstream blading. By explicitly accounting for these unsteady flows in the design of turbomachinery blading for multistage applications, it may be possible to bring about gains in performance and blade life.

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.

Forward rotor vortex effects on counter rotating propeller noise

NASA Technical Reports Server (NTRS)

Laur, Michele; Squires, Becky; Nagel, Robert T.

1992-01-01

Three configurations of a model counter rotating propeller manipulate the blade tip flow by: placing the CRP at angle of attack, installing shrouds, and turning the upstream blades to provide forward sweep. Flow visualization and flow measurements with thermal anemometry show no evidence of a tip vortex; however, a leading edge vortex was detected on aft swept blades. The modifications served to alter the strength and/or path of the leading edge vortex. The vortical flow is eliminated by forward sweep on the upstream propeller blades. Far field acoustic data from each test indicate only small influences on the level and directivity of the BPFs. The interaction tone at the sum of the two BPF's was significantly altered in a consistent manner. As the vortex system varied, the interaction tone was affected: far field noise levels in the forward quandrant increased and the characteristic noise minimum near the plane of rotation became less pronounced and in some cases were eliminated. If the forward propeller leading edge vortex system does not impact the rear propeller in the standard manner, a net increase in the primary interaction tone occurs for the model tested. If the leading edge vortex is removed, the interaction tone increases.

A comprehensive PIV measurement campaign on a fully equipped helicopter model

NASA Astrophysics Data System (ADS)

De Gregorio, Fabrizio; Pengel, Kurt; Kindler, Kolja

2012-07-01

The flow field around a helicopter is characterised by its inherent complexity including effects of fluid-structure interference, shock-boundary layer interaction, and dynamic stall. Since the advancement of computational fluid dynamics and computing capabilities has led to an increasing demand for experimental validation data, a comprehensive wind tunnel test campaign of a fully equipped and motorised generic medium transport helicopter was conducted in the framework of the GOAHEAD project. Different model configurations (with or without main/tail rotor blades) and several flight conditions were investigated. In this paper, the results of the three-component velocity field measurements around the model are surveyed. The effect of the interaction between the main rotor wake and the fuselage for cruise/tail shake flight conditions was analysed based on the flow characteristics downstream from the rotor hub and the rear fuselage hatch. The results indicated a sensible increment of the intensity of the vortex shedding from the lower part of the fuselage and a strong interaction between the blade vortex filaments and the wakes shed by the rotor hub and by the engine exhaust areas. The pitch-up phenomenon was addressed, detecting the blade tip vortices impacting on the horizontal tail plane. For high-speed forward flight, the shock wave formation on the advancing blade was detected, measuring the location on the blade chord and the intensity. Furthermore, dynamic stall on the retreating main rotor blade in high-speed forward flight was observed at r/ R = 0.5 and 0.6. The analysis of the substructures forming the dynamic stall vortex revealed an unexpected spatial concentration suggesting a rotational stabilisation of large-scale structures on the blade.

Navier-Stokes Simulation of UH-60A Rotor/Wake Interaction Using Adaptive Mesh Refinement

NASA Technical Reports Server (NTRS)

Chaderjian, Neal M.

2017-01-01

Time-dependent Navier-Stokes simulations have been carried out for a flexible UH-60A rotor in forward flight, where the rotor wake interacts with the rotor blades. These flow conditions involved blade vortex interaction and dynamic stall, two common conditions that occur as modern helicopter designs strive to achieve greater flight speeds and payload capacity. These numerical simulations utilized high-order spatial accuracy and delayed detached eddy simulation. Emphasis was placed on understanding how improved rotor wake resolution affects the prediction of the normal force, pitching moment, and chord force of the rotor. Adaptive mesh refinement was used to highly resolve the turbulent rotor wake in a computationally efficient manner. Moreover, blade vortex interaction was found to trigger dynamic stall. Time-dependent flow visualization was utilized to provide an improved understanding of the numerical and physical mechanisms involved with three-dimensional dynamic stall.

Wake Geometry Effects on Rotor Blade-Vortex Interaction Noise Directivity

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

Turbofan noise generation. Volume 2: Computer programs

NASA Technical Reports Server (NTRS)

Ventres, C. S.; Theobald, M. A.; Mark, W. D.

1982-01-01

The use of a package of computer programs developed to calculate the in duct acoustic mods excited by a fan/stator stage operating at subsonic tip speed is described. The following three noise source mechanisms are included: (1) sound generated by the rotor blades interacting with turbulence ingested into, or generated within, the inlet duct; (2) sound generated by the stator vanes interacting with the turbulent wakes of the rotor blades; and (3) sound generated by the stator vanes interacting with the velocity deficits in the mean wakes of the rotor blades. The computations for three different noise mechanisms are coded as three separate computer program packages. The computer codes are described by means of block diagrams, tables of data and variables, and example program executions; FORTRAN listings are included.

Turbofan noise generation. Volume 2: Computer programs

NASA Astrophysics Data System (ADS)

Ventres, C. S.; Theobald, M. A.; Mark, W. D.

1982-07-01

The use of a package of computer programs developed to calculate the in duct acoustic mods excited by a fan/stator stage operating at subsonic tip speed is described. The following three noise source mechanisms are included: (1) sound generated by the rotor blades interacting with turbulence ingested into, or generated within, the inlet duct; (2) sound generated by the stator vanes interacting with the turbulent wakes of the rotor blades; and (3) sound generated by the stator vanes interacting with the velocity deficits in the mean wakes of the rotor blades. The computations for three different noise mechanisms are coded as three separate computer program packages. The computer codes are described by means of block diagrams, tables of data and variables, and example program executions; FORTRAN listings are included.

New Tools Being Developed for Engine- Airframe Blade-Out Structural Simulations

NASA Technical Reports Server (NTRS)

Lawrence, Charles

2003-01-01

One of the primary concerns of aircraft structure designers is the accurate simulation of the blade-out event. This is required for the aircraft to pass Federal Aviation Administration (FAA) certification and to ensure that the aircraft is safe for operation. Typically, the most severe blade-out occurs when a first-stage fan blade in a high-bypass gas turbine engine is released. Structural loading results from both the impact of the blade onto the containment ring and the subsequent instantaneous unbalance of the rotating components. Reliable simulations of blade-out are required to ensure structural integrity during flight as well as to guarantee successful blade-out certification testing. The loads generated by these analyses are critical to the design teams for several components of the airplane structures including the engine, nacelle, strut, and wing, as well as the aircraft fuselage. Currently, a collection of simulation tools is used for aircraft structural design. Detailed high-fidelity simulation tools are used to capture the structural loads resulting from blade loss, and then these loads are used as input into an overall system model that includes complete structural models of both the engines and the airframe. The detailed simulation (shown in the figure) includes the time-dependent trajectory of the lost blade and its interactions with the containment structure, and the system simulation includes the lost blade loadings and the interactions between the rotating turbomachinery and the remaining aircraft structural components. General-purpose finite element structural analysis codes are typically used, and special provisions are made to include transient effects from the blade loss and rotational effects resulting from the engine s turbomachinery. To develop and validate these new tools with test data, the NASA Glenn Research Center has teamed with GE Aircraft Engines, Pratt & Whitney, Boeing Commercial Aircraft, Rolls-Royce, and MSC.Software.

An Investigation of Unsteady Impeller-Diffuser Interactions in a Centrifugal Compressor

DTIC Science & Technology

1992-08-01

120 6.20 IDV Measument Positios ............................................................. 121 6.21 LDV...The motivation for radially oriented blades includes ease of manufacture and reduced stress in high speed machines. Backswept blades are used to

Containment of composite fan blades

NASA Technical Reports Server (NTRS)

Stotler, C. L.; Coppa, A. P.

1979-01-01

A lightweight containment was developed for turbofan engine fan blades. Subscale ballistic-type tests were first run on a number of concepts. The most promising configuration was selected and further evaluated by larger scale tests in a rotating test rig. Weight savings made possible by the use of this new containment system were determined and extrapolated to a CF6-size engine. An analytical technique was also developed to predict the released blades motion when involved in the blade/casing interaction process. Initial checkout of this procedure was accomplished using several of the tests run during the program.

Apparatus for Control of Stator Wakes

DTIC Science & Technology

2009-09-18

wake deficit . This has the effect of reducing the blade rate tonal noise of the propulsion rotor. 11 o CN 6 ...upstream of propeller propulsors the sharp wake deficits behind the stators result in unsteady loading and distinguishable peaks in the noise spectra at...trailing edge of a stator blade in order to fill its mean wake deficit to reduce unsteady loading on the rotor blades . Interaction between

Overview of aerothermodynamic loads definition study

NASA Technical Reports Server (NTRS)

Gaugler, Raymond E.

1991-01-01

The objective of the Aerothermodynamic Loads Definition Study is to develop methods of accurately predicting the operating environment in advanced Earth-to-Orbit (ETO) propulsion systems, such as the Space Shuttle Main Engine (SSME) powerhead. Development of time averaged and time dependent three dimensional viscous computer codes as well as experimental verification and engine diagnostic testing are considered to be essential in achieving that objective. Time-averaged, nonsteady, and transient operating loads must all be well defined in order to accurately predict powerhead life. Described here is work in unsteady heat flow analysis, improved modeling of preburner flow, turbulence modeling for turbomachinery, computation of three dimensional flow with heat transfer, and unsteady viscous multi-blade row turbine analysis.

Turbomachinery CFD on parallel computers

NASA Technical Reports Server (NTRS)

Blech, Richard A.; Milner, Edward J.; Quealy, Angela; Townsend, Scott E.

1992-01-01

The role of multistage turbomachinery simulation in the development of propulsion system models is discussed. Particularly, the need for simulations with higher fidelity and faster turnaround time is highlighted. It is shown how such fast simulations can be used in engineering-oriented environments. The use of parallel processing to achieve the required turnaround times is discussed. Current work by several researchers in this area is summarized. Parallel turbomachinery CFD research at the NASA Lewis Research Center is then highlighted. These efforts are focused on implementing the average-passage turbomachinery model on MIMD, distributed memory parallel computers. Performance results are given for inviscid, single blade row and viscous, multistage applications on several parallel computers, including networked workstations.

Heat transfer technology for internal passages of air-cooled blades for heavy-duty gas turbines.

PubMed

Weigand, B; Semmler, K; von Wolfersdorf, J

2001-05-01

The present review paper, although far from being complete, aims to give an overview about the present state of the art in the field of heat transfer technology for internal cooling of gas turbine blades. After showing some typical modern cooled blades, the different methods to enhance heat transfer in the internal passages of air-cooled blades are discussed. The complicated flows occurring in bends are described in detail, because of their increasing importance for modern cooling designs. A short review about testing of cooling design elements is given, showing the interaction of the different cooling features as well. The special focus of the present review has been put on the cooling of blades for heavy-duty gas turbines, which show several differences compared to aero-engine blades.

Examination of the collision force method for analyzing the responses of simple containment/deflection structures to impact by one engine rotor blade fragment

NASA Technical Reports Server (NTRS)

Zirin, R. M.; Witmer, E. A.

1972-01-01

An approximate collision analysis, termed the collision-force method, was developed for studying impact-interaction of an engine rotor blade fragment with an initially circular containment ring. This collision analysis utilizes basic mass, material property, geometry, and pre-impact velocity information for the fragment, together with any one of three postulated patterns of blade deformation behavior: (1) the elastic straight blade model, (2) the elastic-plastic straight shortening blade model, and (3) the elastic-plastic curling blade model. The collision-induced forces are used to predict the resulting motions of both the blade fragment and the containment ring. Containment ring transient responses are predicted by a finite element computer code which accommodates the large deformation, elastic-plastic planar deformation behavior of simple structures such as beams and/or rings. The effects of varying the values of certain parameters in each blade-behavior model were studied. Comparisons of predictions with experimental data indicate that of the three postulated blade-behavior models, the elastic-plastic curling blade model appears to be the most plausible and satisfactory for predicting the impact-induced motions of a ductile engine rotor blade and a containment ring against which the blade impacts.

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

NASA Technical Reports Server (NTRS)

Suder, Kenneth L.

1996-01-01

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

Morphological and mechanical properties of blades of Saccharina latissima

NASA Astrophysics Data System (ADS)

Vettori, Davide; Nikora, Vladimir

2017-09-01

Interactions between water flow and aquatic vegetation strongly depend on morphological and biomechanical characteristics of vegetation. Although any physical or numerical model that aims to replicate flow-vegetation interactions requires these characteristics, information on morphology and mechanics of vegetation living in coastal waters remains insufficient. The present study investigates the mechanical properties of blades of Saccharina latissima, a seaweed species spread along the shores of the UK and North East Atlantic. More than 50 seaweed samples with lengths spanning from 150 mm to 650 mm were collected from Loch Fyne (Scotland) and tested. Seaweed blades had a natural 'stretched droplet' shape with bullations in the central fascia and ruffled edges in the area close to the stipe. Their morphological features showed high variability for samples longer than 400 mm. The blades were almost neutrally buoyant, their material was found to be very flexible and ductile, being stiffer in longer blades. The laboratory tests showed that estimates of tensile Young's modulus appeared to be similar to bending Young's modulus suggesting a reasonable degree of isotropy in studied seaweed tissues.

Smart structure for small wind turbine blade

NASA Astrophysics Data System (ADS)

Supeni, E. E.; Epaarachchi, J. A.; Islam, M. M.; Lau, K. T.

2013-08-01

Wind energy is seen as a viable alternative energy option for future energy demand. The blades of wind turbines are generally regarded as the most critical component of the wind turbine system. Ultimately, the blades act as the prime mover of the whole system which interacts with the wind flow during the production of energy. During wind turbine operation the wind loading cause the deflection of the wind turbine blade which can be significant and affect the turbine efficiency. Such a deflection in wind blade not only will result in lower performance in electrical power generation but also increase of material degradation due high fatigue life and can significantly shorten the longevity for the wind turbine material. In harnessing stiffness of the blade will contribute massive weight factor and consequently excessive bending moment. To overcome this excessive deflection due to wind loading on the blade, it is feasible to use shape memory alloy (SMA) wires which has ability take the blade back to its optimal operational shape. This paper details analytical and experimental work being carried out to minimize blade flapping deflection using SMA.

Determining effects of turbine blades on fluid motion

DOEpatents

Linn, Rodman Ray [Los Alamos, NM; Koo, Eunmo [Los Alamos, NM

2012-05-01

Disclosed is a technique for simulating wind interaction with wind turbines. A turbine blade is divided into radial sections. The effect that each of these radial sections has on the velocities in Eulerian computational cells they overlap is determined. The effect is determined using Lagrangian techniques such that the calculations need not include wind components in the radial direction. A force on each radial section of turbine blade is determined. This force depends on the axial and azimuthal components of the fluid flow in the computational cell and the geometric properties of the turbine blade. The force on the turbine blade is fed back to effect the fluid flow in the computational cell for the next time step.

Determining effects of turbine blades on fluid motion

DOEpatents

Linn, Rodman Ray [Los Alamos, NM; Koo, Eunmo [Los Alamos, NM

2011-05-31

Disclosed is a technique for simulating wind interaction with wind turbines. A turbine blade is divided into radial sections. The effect that each of these radial sections has on the velocities in Eulerian computational cells they overlap is determined. The effect is determined using Lagrangian techniques such that the calculations need not include wind components in the radial direction. A force on each radial section of turbine blade is determined. This force depends on the axial and azimuthal components of the fluid flow in the computational cell and the geometric properties of the turbine blade. The force on the turbine blade is fed back to effect the fluid flow in the computational cell for the next time step.

An analysis for high Reynolds number inviscid/viscid interactions in cascades

NASA Technical Reports Server (NTRS)

Barnett, Mark; Verdon, Joseph M.; Ayer, Timothy C.

1993-01-01

An efficient steady analysis for predicting strong inviscid/viscid interaction phenomena such as viscous-layer separation, shock/boundary-layer interaction, and trailing-edge/near-wake interaction in turbomachinery blade passages is needed as part of a comprehensive analytical blade design prediction system. Such an analysis is described. It uses an inviscid/viscid interaction approach, in which the flow in the outer inviscid region is assumed to be potential, and that in the inner or viscous-layer region is governed by Prandtl's equations. The inviscid solution is determined using an implicit, least-squares, finite-difference approximation, the viscous-layer solution using an inverse, finite-difference, space-marching method which is applied along the blade surfaces and wake streamlines. The inviscid and viscid solutions are coupled using a semi-inverse global iteration procedure, which permits the prediction of boundary-layer separation and other strong-interaction phenomena. Results are presented for three cascades, with a range of inlet flow conditions considered for one of them, including conditions leading to large-scale flow separations. Comparisons with Navier-Stokes solutions and experimental data are also given.

Double row loop-coil configuration for high-speed electrodynamic maglev suspension, guidance, propulsion and guideway directional switching

DOEpatents

He, Jianliang; Rote, Donald M.

1996-01-01

A stabilization and propulsion system comprising a series of loop-coils arranged in parallel rows wherein two rows combine to form one of two magnetic rails. Levitation and lateral stability are provided when the induced field in the magnetic rails interacts with the superconducting magnets mounted on the magnetic levitation vehicle. The loop-coils forming the magnetic rails have specified dimensions and a specified number of turns and by constructing differently these specifications, for one rail with respect to the other, the angle of tilt of the vehicle can be controlled during directional switching. Propulsion is provided by the interaction of a traveling magnetic wave associated with the coils forming the rails and the super conducting magnets on the vehicle.

Double row loop-coil configuration for high-speed electrodynamic maglev suspension, guidance, propulsion and guideway directional switching

DOEpatents

He, J.; Rote, D.M.

1996-05-21

A stabilization and propulsion system are disclosed comprising a series of loop-coils arranged in parallel rows wherein two rows combine to form one of two magnetic rails. Levitation and lateral stability are provided when the induced field in the magnetic rails interacts with the superconducting magnets mounted on the magnetic levitation vehicle. The loop-coils forming the magnetic rails have specified dimensions and a specified number of turns and by constructing differently these specifications, for one rail with respect to the other, the angle of tilt of the vehicle can be controlled during directional switching. Propulsion is provided by the interaction of a traveling magnetic wave associated with the coils forming the rails and the superconducting magnets on the vehicle. 12 figs.

Nonlinear dynamic simulation of single- and multi-spool core engines

NASA Technical Reports Server (NTRS)

Schobeiri, T.; Lippke, C.; Abouelkheir, M.

1993-01-01

In this paper a new computational method for accurate simulation of the nonlinear dynamic behavior of single- and multi-spool core engines, turbofan engines, and power generation gas turbine engines is presented. In order to perform the simulation, a modularly structured computer code has been developed which includes individual mathematical modules representing various engine components. The generic structure of the code enables the dynamic simulation of arbitrary engine configurations ranging from single-spool thrust generation to multi-spool thrust/power generation engines under adverse dynamic operating conditions. For precise simulation of turbine and compressor components, row-by-row calculation procedures were implemented that account for the specific turbine and compressor cascade and blade geometry and characteristics. The dynamic behavior of the subject engine is calculated by solving a number of systems of partial differential equations, which describe the unsteady behavior of the individual components. In order to ensure the capability, accuracy, robustness, and reliability of the code, comprehensive critical performance assessment and validation tests were performed. As representatives, three different transient cases with single- and multi-spool thrust and power generation engines were simulated. The transient cases range from operating with a prescribed fuel schedule, to extreme load changes, to generator and turbine shut down.

Development and Hover Testing of the Active Elevon Rotor

DTIC Science & Technology

2012-05-01

typically aimed at reducing vibration, improving rotor performance, and/or reducing blade -vortex interaction (BVI) or in-plane noise . These efforts...will become unstable, either through a 1-DOF (degree of freedom) flutter or some kind of aeroservoelastic coupling with the rotor blade and/or wake ... blade CAEAs did exhibit electrical arcing (audible noise ), even at oscillatory voltages below ±200 V. This arcing/ noise suggests a latent deficiency

Feasibility study of an aerial manipulator interacting with a vertical wall

DTIC Science & Technology

2017-06-01

each blade . Some tests are run with different levels of PWM input and the resultant angular acceleration in each case is measured with the motion...Helicopter Near a Vertical Surface ...................29 Figure 15. Near-Wall Moment for a Single Blade Helicopter. Source: [30]. .............30...with canted propellers is proposed, so that each blade applies thrust with components in the vertical and in the horizontal plane. In Figure 10

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.

Surface interactions of Fusarium graminearum on barley.

PubMed

Imboden, Lori; Afton, Drew; Trail, Frances

2018-06-01

The filamentous fungus Fusarium graminearum, a devastating pathogen of barley (Hordeum vulgare L.), produces mycotoxins that pose a health hazard. To investigate the surface interactions of F. graminearum on barley, we focused on barley florets, as the most important infection site leading to grain contamination. The fungus interacted with silica-accumulating cells (trichomes and silica/cork cell pairs) on the host surface. We identified variation in trichome-type cells between two-row and six-row barley, and in the role of specific epidermal cells in the ingress of F. graminearum into barley florets. Prickle-type trichomes functioned to trap conidia and were sites of fungal penetration. Infections of more mature florets supported the spread of hyphae into the vascular bundles, whereas younger florets did not show this spread. These differences related directly to the timing and location of increases in silica content during maturation. Focal accumulation of cellulose in infected paleae of two-row and six-row barley indicated that the response is in part linked to trichome type. Overall, silica-accumulating epidermal cells had an expanded role in barley, serving to trap conidia, provide sites for fungal ingress and initiate resistance responses, suggesting a role for silica in pathogen establishment. © 2017 BSPP AND JOHN WILEY & SONS LTD.

Analysis of helicopter blade vortex structure by laser velocimetry

NASA Astrophysics Data System (ADS)

Boutier, A.; Lefèvre, J.; Micheli, F.

1996-05-01

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

Helicopter Non-Unique Trim Strategies for Blade-Vortex Interaction (BVI) Noise Reduction

DTIC Science & Technology

2016-01-22

levels of harmonic rotor noise are one of the key technical barriers preventing the widespread public acceptance of helicopters for commercial...transportation. Blade-Vortex Interaction (BVI) is one such mechanism of rotor noise. BVI noise is a problem for civilian helicopter terminal area...non-rotating frame) on the vehicle trim which in turn affects noise generation. For example, conventional single main rotor helicopters commonly

Transonic blade-vortex interactions - The far field

NASA Astrophysics Data System (ADS)

Lyrintzis, A. S.; George, A. R.

Numerical techniques are developed to predict midfield and far-field helicopter noise due to main-rotor blade-vortex interaction (BVI). The extension of the two-dimensional small-disturbance transonic flow code VTRAN2 (George and Chang, 1983) to the three-dimensional far field (via the Green-function approach of Kirchhoff) is described, and the treatment of oblique BVIs is discussed. Numerical results for a NACA 64A006 airfoil at Mach 0.82 are presented in extensive graphs and characterized in detail. The far-field BVI signature is shown to begin with a strongly forward-directed primary wave (from the original BVI), with an additional downward-directed wave in the case of type C shock motion on the blade.

Reduction of Unsteady Forcing in a Vaned, Contra-Rotating Transonic Turbine Configuration

NASA Technical Reports Server (NTRS)

Clark, John

2010-01-01

HPT blade unsteadiness in the presence of a downstream vane consistent with contra-rotation is characterized by strong interaction at the first harmonic of downstream vane passing. E An existing stage-and-one-half transonic turbine rig design was used as a baseline to investigate means of reducing such a blade-vane interaction. E Methods assessed included: Aerodynamic shaping of HPT blades 3D stacking of the downstream vane Steady pressure-side blowing E Of the methods assessed, a combination of vane bowing and steady pressure-side blowing produced the most favorable result. E Transonic turbine experiments are planned to assess predictive accuracy for the baseline turbine and any design improvements.

Takeoff/approach noise for a model counterrotation propeller with a forward-swept upstream rotor

NASA Technical Reports Server (NTRS)

Woodward, Richard P.; Hall, David G.; Podboy, Gary G.; Jeracki, Robert J.

1993-01-01

A scale model of a counterrotating propeller with forward-swept blades in the forward rotor and aft-swept blades in the aft rotor (designated F39/A31) has been tested in the NASA Lewis 9- by 15-Foot Anechoic Wind Tunnel. This paper presents aeroacoustic results at a takeoff/approach condition of Mach 0.20. Laser Doppler velocimeter results taken in a plane between the two rotors are also included to quantify the interaction flow field. The intention of the forward-swept design is to reduce the magnitude of the forward rotor tip vortex and/or wakes which impinge on the aft rotor, thus lowering the interaction tone levels. A reference model propeller (designated F31/A31), having aft-swept blades in both rotors, was also tested. Aeroelastic performance of the F39/A31 propeller was disappointing. The forward rotor tip region tended to untwist toward higher effective blade angles under load. The forward rotor also exhibited steady state blade flutter at speeds and loadings well below the design condition. The noise results, based on sideline acoustic data, show that the interaction tone levels were up to 8 dB higher with the forward-swept design compared to those for the reference propeller at similar operating conditions, with these tone level differences extending down to lower propeller speeds where flutter did not occur. These acoustic results are for a poorly-performing forward-swept propeller. It is quite possible that a properly-designed forward-swept propeller would exhibit substantial interaction tone level reductions.

Three-Dimensional Aerodynamic Instabilities In Multi-Stage Axial Compressors

NASA Technical Reports Server (NTRS)

Tan, Choon S.; Gong, Yifang; Suder, Kenneth L. (Technical Monitor)

2001-01-01

This thesis presents the conceptualization and development of a computational model for describing three-dimensional non-linear disturbances associated with instability and inlet distortion in multistage compressors. Specifically, the model is aimed at simulating the non-linear aspects of short wavelength stall inception, part span stall cells, and compressor response to three-dimensional inlet distortions. The computed results demonstrated the first-of-a-kind capability for simulating short wavelength stall inception in multistage compressors. The adequacy of the model is demonstrated by its application to reproduce the following phenomena: (1) response of a compressor to a square-wave total pressure inlet distortion; (2) behavior of long wavelength small amplitude disturbances in compressors; (3) short wavelength stall inception in a multistage compressor and the occurrence of rotating stall inception on the negatively sloped portion of the compressor characteristic; (4) progressive stalling behavior in the first stage in a mismatched multistage compressor; (5) change of stall inception type (from modal to spike and vice versa) due to IGV stagger angle variation, and "unique rotor tip incidence" at these points where the compressor stalls through short wavelength disturbances. The model has been applied to determine the parametric dependence of instability inception behavior in terms of amplitude and spatial distribution of initial disturbance, and intra-blade-row gaps. It is found that reducing the inter-blade row gaps suppresses the growth of short wavelength disturbances. It is also concluded from these parametric investigations that each local component group (rotor and its two adjacent stators) has its own instability point (i.e. conditions at which disturbances are sustained) for short wavelength disturbances, with the instability point for the compressor set by the most unstable component group. For completeness, the methodology has been extended to describe finite amplitude disturbances in high-speed compressors. Results are presented for the response of a transonic compressor subjected to inlet distortions.

Simulation of the Vibratory Condition of the Compressor Blade with a Pressed wire Material “MR” Damper Which Located Around the Root Attachment

NASA Astrophysics Data System (ADS)

Gvozdev, Alexander S.; Melentjev, Vladimir S.

2018-01-01

When you create a modern gas turbine engines urgent task is to improve the reliability by preventing fatigue damages of rotor blades. Such damage is largely determined by the level of vibration stresses. In this paper, using the finite element method and transient analysis of propose a method calculating the damping characteristics of the plates of the pressed wire material “MR” around the root attachment of the compressor blades of a gas turbine engine. Where taken into account contact interaction between the blades and the impeller disk.

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.

Helicopter rotor wake geometry and its influence in forward flight. Volume 1: Generalized wake geometry and wake effect on rotor airloads and performance

NASA Technical Reports Server (NTRS)

Egolf, T. A.; Landgrebe, A. J.

1983-01-01

An analytic investigation to generalize wake geometry of a helicopter rotor in steady level forward flight and to demonstrate the influence of wake deformation in the prediction of rotor airloads and performance is described. Volume 1 presents a first level generalized wake model based on theoretically predicted tip vortex geometries for a selected representative blade design. The tip vortex distortions are generalized in equation form as displacements from the classical undistorted tip vortex geometry in terms of vortex age, blade azimuth, rotor advance ratio, thrust coefficient, and number of blades. These equations were programmed to provide distorted wake coordinates at very low cost for use in rotor airflow and airloads prediction analyses. The sensitivity of predicted rotor airloads, performance, and blade bending moments to the modeling of the tip vortex distortion are demonstrated for low to moderately high advance ratios for a representative rotor and the H-34 rotor. Comparisons with H-34 rotor test data demonstrate the effects of the classical, predicted distorted, and the newly developed generalized wake models on airloads and blade bending moments. Use of distorted wake models results in the occurrence of numerous blade-vortex interactions on the forward and lateral sides of the rotor disk. The significance of these interactions is related to the number and degree of proximity to the blades of the tip vortices. The correlation obtained with the distorted wake models (generalized and predicted) is encouraging.

Variable-Speed Power-Turbine Research at Glenn Research Center

NASA Technical Reports Server (NTRS)

Welch, Gerard E.; McVetta, Ashlie B.; Stevens, Mark A.; Howard, Samuel A.; Giel, Paul W.; Ameri, Ali, A.; To, Waiming; Skoch, Gary J.; Thurman, Douglas R.

2012-01-01

The main rotors of the NASA Large Civil Tilt-Rotor (LCTR) notional vehicle operate over a wide speed-range, from 100 percent at takeoff to 54 percent at cruise. The variable-speed power turbine (VSPT) offers one approach by which to effect this speed variation. VSPT aerodynamics challenges include high work factors at cruise, wide (40 to 60 ) incidence-angle variations in blade and vane rows over the speed range, and operation at low Reynolds numbers. Rotordynamics challenges include potential responsiveness to shaft modes within the 50 percent VSPT speed-range. A research effort underway at NASA Glenn Research Center, intended to address these key aerodynamic and rotordynamic challenges, is described. Conceptual design and 3-D multistage RANS and URANS analyses, conducted internally and under contract, provide expected VSPT sizing, stage-count, performance and operability information, and maps for system studies. Initial steps toward experimental testing of incidence-tolerant blading in a transonic linear cascade are described, and progress toward development/improvement of a simulation capability for multistage turbines with low Reynolds number transitional flow is summarized. Preliminary rotordynamics analyses indicate that viable concept engines with 50 percent VSPT shaft-speed range. Assessments of potential paths toward VSPT component-level testing are summarized.

Manipulation of upstream rotor leading edge vortex and its effects on counter rotating propeller noise

NASA Technical Reports Server (NTRS)

Squires, Becky

1993-01-01

The leading edge vortex of a counter rotating propeller (CRP) model was altered by using shrouds and by turning the upstream rotors to a forward sweep configuration. Performance, flow, and acoustic data were used to determine the effect of vortex impingement on the noise signature of the CRP system. Forward sweep was found to eliminate the leading edge vortex of the upstream blades. Removal of the vortex had little effect on the tone noise at the forward and rear blade passing frequencies (BPF's) but significantly altered both the sound pressure level and directivity of the interaction tone which occurs at the sum of the two BPF's. A separate manipulation of the leading edge vortex performed by installing shrouds of various inlet length on the CRP verified that diverting the vortex path increases the noise level of the interaction tone. An unexpected link has been established between the interaction tone and the leading edge vortex-blade interaction phenomenon.

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

NASA Technical Reports Server (NTRS)

Chan, Tony; Dejong, Frederik J.

1993-01-01

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

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.

Active Robust Control of Elastic Blade Element Containing Magnetorheological Fluid

NASA Astrophysics Data System (ADS)

Sivrioglu, Selim; Cakmak Bolat, Fevzi

2018-03-01

This research study proposes a new active control structure to suppress vibrations of a small-scale wind turbine blade filled with magnetorheological (MR) fluid and actuated by an electromagnet. The aluminum blade structure is manufactured using the airfoil with SH3055 code number which is designed for use on small wind turbines. An interaction model between MR fluid and the electromagnetic actuator is derived. A norm based multi-objective H2/H∞ controller is designed using the model of the elastic blade element. The H2/H∞ controller is experimentally realized under the impact and steady state aerodynamic load conditions. The results of experiments show that the MR fluid is effective for suppressing vibrations of the blade structure.

DoD High Performance Computing Modernization Program FY16 Annual Report

DTIC Science & Technology

2018-05-02

vortex shedding from rotor blade tips using adaptive mesh refinement gives Helios the unique capability to assess the interaction of these vortices...with the fuselage and nearby rotor blades . Helios provides all the benefits for rotary-winged aircraft that Kestrel does for fixed-wing aircraft...rotor blade upgrade of the CH-47F Chinook helicopter to achieve up to an estimated 2,000 pounds increase in hover thrust (~10%) with limited

Experimental and numerical simulation of a rotor/stator interaction event localized on a single blade within an industrial high-pressure compressor

NASA Astrophysics Data System (ADS)

Batailly, Alain; Agrapart, Quentin; Millecamps, Antoine; Brunel, Jean-François

2016-08-01

This contribution addresses a confrontation between the experimental simulation of a rotor/stator interaction case initiated by structural contacts with numerical predictions made with an in-house numerical strategy. Contrary to previous studies carried out within the low-pressure compressor of an aircraft engine, this interaction is found to be non-divergent: high amplitudes of vibration are experimentally observed and numerically predicted over a short period of time. An in-depth analysis of experimental data first allows for a precise characterization of the interaction as a rubbing event involving the first torsional mode of a single blade. Numerical results are in good agreement with experimental observations: the critical angular speed, the wear patterns on the casing as well as the blade dynamics are accurately predicted. Through out the article, the in-house numerical strategy is also confronted to another numerical strategy that may be found in the literature for the simulation of rubbing events: key differences are underlined with respect to the prediction of non-linear interaction phenomena.

Rotor noise due to blade-turbulence interaction

NASA Astrophysics Data System (ADS)

Ishimaru, K.

1983-01-01

The time-averaged intensity density function of the acoustic radiation from rotating blades is derived by replacing blades with rotating dipoles. This derivation is done under the following turbulent inflow conditions: turbulent ingestion with no inlet strut wakes, inflow turbulence elongation and contraction with no inlet strut wakes, and inlet strut wakes. Dimensional analysis reveals two non-dimensional parameters which play important roles in generating the blade-passing frequency tone and its multiples. The elongation and contraction of inflow turbulence has a strong effect on the generation of the blade-passing frequency tone and its multiples. Increasing the number of rotor blades widens the peak at the blade-passing frequency and its multiples. Increasing the rotational speed widens the peak under the condition that the non-dimensional parameter involving the rotational speed is fixed. The number of struts and blades should be chosen so that (the least common multiple of them)-(rotational speed) is in the cutoff range of Sears' function, in order to minimize the effect of the mean flow deficit on the time averaged intensity density function.

Compressor seal rub energetics study

NASA Technical Reports Server (NTRS)

Laverty, W. F.

1978-01-01

The rub mechanics of compressor abradable blade tip seals at simulated engine conditions were investigated. Twelve statistically planned, instrumented rub tests were conducted with titanium blades and Feltmetal fibermetal rubstrips. The tests were conducted with single stationary blades rubbing against seal material bonded to rotating test disks. The instantaneous rub torque, speed, incursion rate and blade temperatures were continuously measured and recorded. Basic rub parameters (incursion rate, rub depth, abradable density, blade thickness and rub velocity) were varied to determine the effects on rub energy and heat split between the blade, rubstrip surface and rub debris. The test data was reduced, energies were determined and statistical analyses were completed to determine the primary and interactive effects. Wear surface morphology, profile measurements and metallographic analysis were used to determine wear, glazing, melting and material transfer. The rub energies for these tests were most significantly affected by the incursion rate while rub velocity and blade thickness were of secondary importance. The ratios of blade wear to seal wear were representative of those experienced in engine operation of these seal system materials.

DNS investigation of the dynamical behaviour of trailing vortices in unbaffled stirred vessels at transitional Reynolds numbers

NASA Astrophysics Data System (ADS)

Başbuǧ, S.; Papadakis, G.; Vassilicos, J. C.

2017-06-01

Flow in an unbaffled stirred vessel agitated by a four-bladed radial impeller is investigated by using direct numerical simulations at Re = 320 and 1600. We observe fluctuations in the power consumption with a peak frequency at ca. three times the impeller rotational speed for both Reynolds numbers. It is discovered that these fluctuations are associated with a periodic event in the wake of the blades, which involves alternating growth and decay of the upper and lower cores of the trailing vortex pair as well as up-and-down swinging motion of the radial jet. Moreover, the phase relation between the wakes of the different blades is examined in detail. Further studies using fractal-shaped blades show that the exact blade shape does not have a strong influence on this phenomenon. However, the wake interaction between the blades, hence the number of blades, has a direct influence on the unsteadiness of trailing vortices.

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.

On the effect of moving blade grid on the flow field characteristics

NASA Astrophysics Data System (ADS)

Procházka, Pavel; Uruba, Václav; Pešek, Luděk; Bula, VÍtězslav

2018-06-01

The motivation of this paper is the continual development of the blades for the last stage of the steam turbine. The biggest problem is the slenderness of such blades and the extreme sensitivity to aeroelastic vibrations (flutter) caused by the instabilities present in the flow. This experimental research is dealing with the aeroelastic binding of the moving blades located in the blade grid with the flow field and vice versa. A parallelogram is used to ensure one order of freedom of the blade. The grid has five blades in total, three of them are driven by force control using three shakers. The deviation as well as force response is measured by strain gauges and dynamometers. The flow field statistical as well as dynamical characteristics are measured by optical method Particle Image Velocimetry. The grid is connected to the blow-down wind tunnel with velocity range up to 40 m/s. The aeroelastic binding is investigated in dependency on used actuation frequency and maximal amplitude (the intensity of force actuation) and on different Reynolds numbers. The flow field and the wake behind each individual blade are studied and the maximal interaction is examined for individual inter-blade phase angle of the grid.

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

NASA Technical Reports Server (NTRS)

Dittmar, James H.

2000-01-01

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

Vortex dynamics during blade-vortex interactions

NASA Astrophysics Data System (ADS)

Peng, Di; Gregory, James W.

2015-05-01

Vortex dynamics during parallel blade-vortex interactions (BVIs) were investigated in a subsonic wind tunnel using particle image velocimetry (PIV). Vortices were generated by applying a rapid pitch-up motion to an airfoil through a pneumatic system, and the subsequent interactions with a downstream, unloaded target airfoil were studied. The blade-vortex interactions may be classified into three categories in terms of vortex behavior: close interaction, very close interaction, and collision. For each type of interaction, the vortex trajectory and strength variation were obtained from phase-averaged PIV data. The PIV results revealed the mechanisms of vortex decay and the effects of several key parameters on vortex dynamics, including separation distance (h/c), Reynolds number, and vortex sense. Generally, BVI has two main stages: interaction between vortex and leading edge (vortex-LE interaction) and interaction between vortex and boundary layer (vortex-BL interaction). Vortex-LE interaction, with its small separation distance, is dominated by inviscid decay of vortex strength due to pressure gradients near the leading edge. Therefore, the decay rate is determined by separation distance and vortex strength, but it is relatively insensitive to Reynolds number. Vortex-LE interaction will become a viscous-type interaction if there is enough separation distance. Vortex-BL interaction is inherently dominated by viscous effects, so the decay rate is dependent on Reynolds number. Vortex sense also has great impact on vortex-BL interaction because it changes the velocity field and shear stress near the surface.

Whirl Flutter Stability of Two-Bladed Proprotor/Pylon Systems In High Speed Flight

NASA Technical Reports Server (NTRS)

Singh, Beerinder; Chopra, Inderjit; Pototzky, A. (Technical Monitor)

2002-01-01

The lack of polar symmetry in two-bladed rotors leads to equations of motion with periodic coefficients in axial flight, which is contrary to three or more bladed rotors that result in constant coefficient equations. With periodic coefficients, the analysis becomes involved, as a result very few studies have been directed towards the analysis of two-bladed rotors. In this paper, the aeroelastic stability of two-bladed proprotor/pylon/wing combinations is examined in high speed axial flight. Several parametric studies are carried out to illustrate the special nature of two-bladed proprotors and to better understand the mechanism of whirl-flutter in such rotors. The wing beam bending mode for two-bladed rotors is found to be stable over the range of parameters examined, a behaviour very different from three-bladed rotors. Also, the wing torsion mode exhibits a new type of instability similar to a wing torsional divergence scouring at I/rev frequency. This type of behaviour is not seen in three and more bladed rotors. The interaction between wing chordwise bending and torsion modes is found to be much greater in the case of two-bladed rotors and, over the range of parameters considered, these two modes govern the stability of the system.

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

NASA Astrophysics Data System (ADS)

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

2013-12-01

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

Darrieus rotor aerodynamics

NASA Astrophysics Data System (ADS)

Klimas, P. C.

1982-05-01

A summary of the progress of modeling the aerodynamic effects on the blades of a Darrieus wind turbine is presented. Interference is discussed in terms of blade/blade wake interaction and improvements in single and multiple stream tube models, of vortex simulations of blades and their wakes, and a hybrid momentum/vortex code to combine fast computation time with interference-describing capabilities. An empirical model has been developed for treating the properties of dynamic stall such as airfoil geometry, Reynolds number, reduced frequency, angle-of-attack, and Mach number. Pitching circulation has been subjected to simulation as potential flow about a two-dimensional flat plate, along with applications of the concepts of virtual camber and virtual incidence, with a cambered airfoil operating in a rectilinear flowfield. Finally, a need to develop a loading model suitable for nonsymmetrical blade sections is indicated, as well as blade behavior in a dynamic, curvilinear regime.

Aerodynamic Analysis of Morphing Blades

NASA Astrophysics Data System (ADS)

Harris, Caleb; Macphee, David; Carlisle, Madeline

2016-11-01

Interest in morphing blades has grown with applications for wind turbines and other aerodynamic blades. This passive control method has advantages over active control methods such as lower manufacturing and upkeep costs. This study has investigated the lift and drag forces on individual blades with experimental and computational analysis. The goal has been to show that these blades delay stall and provide larger lift-to-drag ratios at various angles of attack. Rigid and flexible airfoils were cast from polyurethane and silicone respectively, then lift and drag forces were collected from a load cell during 2-D testing in a wind tunnel. Experimental data was used to validate computational models in OpenFOAM. A finite volume fluid-structure-interaction solver was used to model the flexible blade in fluid flow. Preliminary results indicate delay in stall and larger lift-to-drag ratios by maintaining more optimal angles of attack when flexing. Funding from NSF REU site Grant EEC 1358991 is greatly appreciated.

General approach and scope. [rotor blade design optimization

NASA Technical Reports Server (NTRS)

Adelman, Howard M.; Mantay, Wayne R.

1989-01-01

This paper describes a joint activity involving NASA and Army researchers at the NASA Langley Research Center to develop optimization procedures aimed at improving the rotor blade design process by integrating appropriate disciplines and accounting for all of the important interactions among the disciplines. The disciplines involved include rotor aerodynamics, rotor dynamics, rotor structures, airframe dynamics, and acoustics. The work is focused on combining these five key disciplines in an optimization procedure capable of designing a rotor system to satisfy multidisciplinary design requirements. Fundamental to the plan is a three-phased approach. In phase 1, the disciplines of blade dynamics, blade aerodynamics, and blade structure will be closely coupled, while acoustics and airframe dynamics will be decoupled and be accounted for as effective constraints on the design for the first three disciplines. In phase 2, acoustics is to be integrated with the first three disciplines. Finally, in phase 3, airframe dynamics will be fully integrated with the other four disciplines. This paper deals with details of the phase 1 approach and includes details of the optimization formulation, design variables, constraints, and objective function, as well as details of discipline interactions, analysis methods, and methods for validating the procedure.

Full-Potential Modeling of Blade-Vortex Interactions. Degree awarded by George Washington Univ., Feb. 1987

NASA Technical Reports Server (NTRS)

Jones, Henry E.

1997-01-01

A study of the full-potential modeling of a blade-vortex interaction was made. A primary goal of this study was to investigate the effectiveness of the various methods of modeling the vortex. The model problem restricts the interaction to that of an infinite wing with an infinite line vortex moving parallel to its leading edge. This problem provides a convenient testing ground for the various methods of modeling the vortex while retaining the essential physics of the full three-dimensional interaction. A full-potential algorithm specifically tailored to solve the blade-vortex interaction (BVI) was developed to solve this problem. The basic algorithm was modified to include the effect of a vortex passing near the airfoil. Four different methods of modeling the vortex were used: (1) the angle-of-attack method, (2) the lifting-surface method, (3) the branch-cut method, and (4) the split-potential method. A side-by-side comparison of the four models was conducted. These comparisons included comparing generated velocity fields, a subcritical interaction, and a critical interaction. The subcritical and critical interactions are compared with experimentally generated results. The split-potential model was used to make a survey of some of the more critical parameters which affect the BVI.

Helicopter Blade-Vortex Interaction Noise with Comparisons to CFD Calculations

NASA Technical Reports Server (NTRS)

McCluer, Megan S.

1996-01-01

A comparison of experimental acoustics data and computational predictions was performed for a helicopter rotor blade interacting with a parallel vortex. The experiment was designed to examine the aerodynamics and acoustics of parallel Blade-Vortex Interaction (BVI) and was performed in the Ames Research Center (ARC) 80- by 120-Foot Subsonic Wind Tunnel. An independently generated vortex interacted with a small-scale, nonlifting helicopter rotor at the 180 deg azimuth angle to create the interaction in a controlled environment. Computational Fluid Dynamics (CFD) was used to calculate near-field pressure time histories. The CFD code, called Transonic Unsteady Rotor Navier-Stokes (TURNS), was used to make comparisons with the acoustic pressure measurement at two microphone locations and several test conditions. The test conditions examined included hover tip Mach numbers of 0.6 and 0.7, advance ratio of 0.2, positive and negative vortex rotation, and the vortex passing above and below the rotor blade by 0.25 rotor chords. The results show that the CFD qualitatively predicts the acoustic characteristics very well, but quantitatively overpredicts the peak-to-peak sound pressure level by 15 percent in most cases. There also exists a discrepancy in the phasing (about 4 deg) of the BVI event in some cases. Additional calculations were performed to examine the effects of vortex strength, thickness, time accuracy, and directionality. This study validates the TURNS code for prediction of near-field acoustic pressures of controlled parallel BVI.

The Vortex Lattice Method for the Rotor-Vortex Interaction Problem

NASA Technical Reports Server (NTRS)

Padakannaya, R.

1974-01-01

The rotor blade-vortex interaction problem and the resulting impulsive airloads which generate undesirable noise levels are discussed. A numerical lifting surface method to predict unsteady aerodynamic forces induced on a finite aspect ratio rectangular wing by a straight, free vortex placed at an arbitrary angle in a subsonic incompressible free stream is developed first. Using a rigid wake assumption, the wake vortices are assumed to move downsteam with the free steam velocity. Unsteady load distributions are obtained which compare favorably with the results of planar lifting surface theory. The vortex lattice method has been extended to a single bladed rotor operating at high advance ratios and encountering a free vortex from a fixed wing upstream of the rotor. The predicted unsteady load distributions on the model rotor blade are generally in agreement with the experimental results. This method has also been extended to full scale rotor flight cases in which vortex induced loads near the tip of a rotor blade were indicated. In both the model and the full scale rotor blade airload calculations a flat planar wake was assumed which is a good approximation at large advance ratios because the downwash is small in comparison to the free stream at large advance ratios. The large fluctuations in the measured airloads near the tip of the rotor blade on the advance side is predicted closely by the vortex lattice method.

Heat transfer and pressure measurements for the SSME fuel-side turbopump

NASA Technical Reports Server (NTRS)

Dunn, Michael G.

1990-01-01

A measurement program is currently underway at the Calspan-UB Research Center (CUBRC) which utilizes the Rocketdyne two-state fuel-side turbine with the engine geometric configuration reproduced. This is a full two-state turbine for which the vane rows and the blades are the engine hardware currently used on the Space Shuttle turbopump. A status report is provided for the experimental program and a description of the instrumentation and the measurements to be performed. The specific items that will be illustrated and described are as follows: (1) the gas flow path, (2) the heat-flux instrumentation, (3) the surface-pressure instrumentation, (4) the experimental conditions for which data will be obtained, and (5) the specific measurements that will be performed.

A microprocessor-based automation test system for the experiment of the multi-stage compressor

NASA Astrophysics Data System (ADS)

Zhang, Huisheng; Lin, Chongping

1991-08-01

An automation test system that is controlled by the microprocessor and used in the multistage compressor experiment is described. Based on the analysis of the compressor experiment performances, a complete hardware system structure is set up. It is composed of a IBM PC/XT computer, a large scale sampled data system, the moving machine with three directions, the scanners, the digital instrumentation and some output devices. A program structure of real-time software system is described. The testing results show that this test system can take the measure of many parameter magnitudes in the blade row places and on a boundary layer in different states. The automatic extent and the accuracy of experiment is increased and the experimental cost is reduced.

Performance of two-stage fan with larger dampers on first-stage rotor

NASA Technical Reports Server (NTRS)

Urasek, D. C.; Cunnan, W. S.; Stevans, W.

1979-01-01

The performance of a two stage, high pressure-ratio fan, having large, part-span vibration dampers on the first stage rotor is presented and compared with an identical aerodynamically designed fan having smaller dampers. Comparisons of the data for the two damper configurations show that with increased damper size: (1) very high losses in the damper region reduced overall efficiency of first stage rotor by approximately 3 points, (2) the overall performance of each blade row, downstream of the damper was not significantly altered, although appreciable differences in the radial distributions of various performance parameters were noted, and (3) the lower performance of the first stage rotor decreased the overall fan efficiency more than 1 percentage point.

An experimental study of windturbine noise from blade-tower wake interaction

NASA Astrophysics Data System (ADS)

Marcus, E. N.; Harris, W. L.

1983-04-01

A program of experiments has been conducted to study the impulsive noise of a horizontal axis windturbine. These tests were performed on a 1/53 scale model of the DOE-NASA MOD-1 windturbine. Experiments were performed in the M.I.T. 5 x 7-1/2 ft Anechoic Windtunnel Facility. The impulsive noise of a horizontal axis windturbine is observed to result from repeated blade passage through the mean velocity deficit induced in the lee of the windturbine support tower. The two factors which most influence this noise are rotation speed and tower drag coefficient. The intensity of noise from blade tower wake interaction is predicted to increase with the fourth power of the RPM and the second power of the tower drag coefficient. These predictions are confirmed in experiments. Further experiments are also presented in order to observe directionality of the acoustic field as well as the acoustic influence of tower shape and blade number.

Three-Dimensional Unsteady Simulation of a Modern High Pressure Turbine Stage Using Phase Lag Periodicity: Analysis of Flow and Heat Transfer

NASA Technical Reports Server (NTRS)

Shyam, Vikram; Ameri, Ali; Luk, Daniel F.; Chen, Jen-Ping

2010-01-01

Unsteady three-dimensional RANS simulations have been performed on a highly loaded transonic turbine stage and results are compared to steady calculations as well as experiment. A low Reynolds number k- turbulence model is employed to provide closure for the RANS system. A phase-lag boundary condition is used in the periodic direction. This allows the unsteady simulation to be performed by using only one blade from each of the two rows. The objective of this paper is to study the effect of unsteadiness on rotor heat transfer and to glean any insight into unsteady flow physics. The role of the stator wake passing on the pressure distribution at the leading edge is also studied. The simulated heat transfer and pressure results agreed favorably with experiment. The time-averaged heat transfer predicted by the unsteady simulation is higher than the heat transfer predicted by the steady simulation everywhere except at the leading edge. The shock structure formed due to stator-rotor interaction was analyzed. Heat transfer and pressure at the hub and casing were also studied. Thermal segregation was observed that leads to the heat transfer patterns predicted by steady and unsteady simulations to be different.

The Role of Flow Diagnostic Techniques in Fan and Open Rotor Noise Modeling

NASA Technical Reports Server (NTRS)

Envia, Edmane

2016-01-01

A principal source of turbomachinery noise is the interaction of the rotating and stationary blade rows with the perturbations in the airstream through the engine. As such, a lot of research has been devoted to the study of the turbomachinery noise generation mechanisms. This is particularly true of fan and open rotors, both of which are the major contributors to the overall noise output of modern aircraft engines. Much of the research in fan and open rotor noise has been focused on developing theoretical models for predicting their noise characteristics. These models, which run the gamut from the semi-empirical to fully computational ones, are, in one form or another, informed by the description of the unsteady flow-field in which the propulsors (i.e., the fan and open rotors) operate. Not surprisingly, the fidelity of the theoretical models is dependent, to a large extent, on capturing the nuances of the unsteady flowfield that have a direct role in the noise generation process. As such, flow diagnostic techniques have proven to be indispensible in identifying the shortcoming of theoretical models and in helping to improve them. This presentation will provide a few examples of the role of flow diagnostic techniques in assessing the fidelity and robustness of the fan and open rotor noise prediction models.

The structure and development of streamwise vortex arrays embedded in a turbulent boundary layer. Ph.D. Thesis - Case Western Reserve Univ.

NASA Technical Reports Server (NTRS)

Wendt, Bruce J.; Greber, Isaac; Hingst, Warren R.

1991-01-01

An investigation of the structure and development of streamwise vortices embedded in a turbulent boundary layer was conducted. The vortices were generated by a single spanwise row of rectangular vortex generator blades. A single embedded vortex was examined, as well as arrays of embedded counter rotating vortices produced by equally spaced vortex generators. Measurements of the secondary velocity field in the crossplane provided the basis for characterization of vortex structure. Vortex structure was characterized by four descriptors. The center of each vortex core was located at the spanwise and normal position of peak streamwise vorticity. Vortex concentration was characterized by the magnitude of the peak streamwise vorticity, and the vortex strength by its circulation. Measurements of the secondary velocity field were conducted at two crossplane locations to examine the streamwise development of the vortex arrays. Large initial spacings of the vortex generators produced pairs of strong vortices which tended to move away from the wall region while smaller spacings produced tight arrays of weak vortices close to the wall. A model of vortex interaction and development is constructed using the experimental results. The model is based on the structure of the Oseen Vortex. Vortex trajectories are modelled by including the convective effects of neighbors.

Acoustic and aerodynamic performance of a 1.83-meter (6-ft) diameter 1.25-pressure-ratio fan (QF-8)

NASA Technical Reports Server (NTRS)

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

1976-01-01

A 1.25-pressure-ratio 1.83-meter (6-ft) tip diameter experimental fan stage with characteristics suitable for engine application on STOL aircraft was tested for acoustic and aerodynamic performance. The design incorporated proven features for low noise, including absence of inlet guide vanes, low rotor blade tip speed, low aerodynamic blade loading, and long axial spacing between the rotor and stator blade rows. The fan was operated with five exhaust nozzle areas. The stage noise levels generally increased with a decrease in nozzle area. Separation of the acoustic one-third octave results into broadband and pure-tone components showed the broadband noise to be greater than the corresponding pure-tone components. The sideline perceived noise was highest in the rear quadrants. The acoustic results of QF-8 were compared with those of two similar STOL application fans in the test series. The QF-8 had somewhat higher relative noise levels than those of the other two fans. The aerodynamic results of QF-8 and the other two fans were compared with corresponding results from 50.8-cm (20-in.) diam scale models of these fans and design values. Although the results for the full-scale and scale models of the other two fans were in reasonable agreement for each design, the full-scale fan QF-8 results showed poor performance compared with corresponding model results and design expectations. Facility effects of the full-scale fan QF-8 installation were considered in analyzing this discrepancy.

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.

Flow structure generated by perpendicular blade vortex interaction and implications for helicopter noise predictions

NASA Technical Reports Server (NTRS)

Devenport, William J.; Glegg, Stewart A. L.

1995-01-01

This report summarizes accomplishments and progress for the period ending April 1995. Much of the work during this period has concentrated on preparation for an analysis of data produced by an extensive wind tunnel test. Time has also been spent further developing an empirical theory to account for the effects of blade-vortex interaction upon the circulation distribution of the vortex and on preliminary measurements aimed at controlling the vortex core size.

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.

Development and Validation of a Multidisciplinary Tool for Accurate and Efficient Rotorcraft Noise Prediction (MUTE)

NASA Technical Reports Server (NTRS)

Liu, Yi; Anusonti-Inthra, Phuriwat; Diskin, Boris

2011-01-01

A physics-based, systematically coupled, multidisciplinary prediction tool (MUTE) for rotorcraft noise was developed and validated with a wide range of flight configurations and conditions. MUTE is an aggregation of multidisciplinary computational tools that accurately and efficiently model the physics of the source of rotorcraft noise, and predict the noise at far-field observer locations. It uses systematic coupling approaches among multiple disciplines including Computational Fluid Dynamics (CFD), Computational Structural Dynamics (CSD), and high fidelity acoustics. Within MUTE, advanced high-order CFD tools are used around the rotor blade to predict the transonic flow (shock wave) effects, which generate the high-speed impulsive noise. Predictions of the blade-vortex interaction noise in low speed flight are also improved by using the Particle Vortex Transport Method (PVTM), which preserves the wake flow details required for blade/wake and fuselage/wake interactions. The accuracy of the source noise prediction is further improved by utilizing a coupling approach between CFD and CSD, so that the effects of key structural dynamics, elastic blade deformations, and trim solutions are correctly represented in the analysis. The blade loading information and/or the flow field parameters around the rotor blade predicted by the CFD/CSD coupling approach are used to predict the acoustic signatures at far-field observer locations with a high-fidelity noise propagation code (WOPWOP3). The predicted results from the MUTE tool for rotor blade aerodynamic loading and far-field acoustic signatures are compared and validated with a variation of experimental data sets, such as UH60-A data, DNW test data and HART II test data.

Preliminary Investigation of the Shock-Boundary Layer Interaction in a Simulated Fan Passage

DTIC Science & Technology

1991-03-01

unlimited 2b DECLASSIFICATION/DOWNGRADING SCHEDULE 4 PERFORMING ORGANIZATION REPORT NUMBER(S) 5 MONITORING ORGANIZATION REPORT NUMBER(S) 6a NAME OF...Figure 4. Vortex Generator Jets Configuration [Ref. 2] 27 Figure 5 . Cascade Geometry 28 Figure 6. Schematic of Transonic Cascade Wind Tunnel 29 Figure 7... 65 Figure A9. Test Section Top Blade 66 Figure A1O. Test Section Middle Blade 67 Figure A 11. Test Section Lower Blade 68 Figure A12. Pressure Tap

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

NASA Astrophysics Data System (ADS)

Roger, Michel; Schram, Christophe; Moreau, Stéphane

2014-01-01

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

Bolted joints in graphite-epoxy composites

NASA Technical Reports Server (NTRS)

Hart-Smith, L. J.

1976-01-01

All-graphite/epoxy laminates and hybrid graphite-glass/epoxy laminates were tested. The tests encompassed a range of geometries for each laminate pattern to cover the three basic failure modes - net section tension failure through the bolt hole, bearing and shearout. Static tensile and compressive loads were applied. A constant bolt diameter of 6.35 mm (0.25 in.) was used in the tests. The interaction of stress concentrations associated with multi-row bolted joints was investigated by testing single- and double-row bolted joints and open-hole specimens in tension. For tension loading, linear interaction was found to exist between the bearing stress reacted at a given bolt hole and the remaining tension stress running by that hole to be reacted elsewhere. The interaction under compressive loading was found to be non-linear. Comparative tests were run using single-lap bolted joints and double-lap joints with pin connection. Both of these joint types exhibited lower strengths than were demonstrated by the corresponding double-lap joints. The analysis methods developed here for single bolt joints are shown to be capable of predicting the behavior of multi-row joints.

Soybean Yield and Heterodera glycines Population Dynamics as Affected by Cultural Practices in Major Production Areas of the United States and Canada

USDA-ARS?s Scientific Manuscript database

Little information is available on the interactive effects of tillage and row spacing on yield of soybean and population dynamics of H. glycines. This study investigated the effects of rotation of soybean and corn, tillage, row spacing, and cultivar on yield of soybean and population dynamics of H. ...

Parametric Investigation of the Effect of Hub Pitching Moment on Blade Vortex Interaction (BVI) Noise of an Isolated Rotor

DTIC Science & Technology

2016-05-19

1 Parametric Investigation of the Effect of Hub Pitching Moment on Blade Vortex Interaction (BVI) Noise of an Isolated Rotor Carlos Malpica...changing the hub pitching moment for an isolated rotor, trimmed in nominal 80 knot, 6 and 12 deg descent, flight conditions, alters the miss distance...compensate for the uncomfortable change in fuselage pitch attitude introduced by a fuselage-mounted X-force controller. NOMENCLATURE xM C rolling

A parametric study of transonic blade-vortex interaction noise

NASA Technical Reports Server (NTRS)

Lyrintzis, A. S.

1991-01-01

Several parameters of transonic blade-vortex interactions (BVI) are being studied and some ideas for noise reduction are introduced and tested using numerical simulation. The model used is the two-dimensional high frequency transonic small disturbance equation with regions of distributed vorticity (VTRAN2 code). The far-field noise signals are obtained by using the Kirchhoff method with extends the numerical 2-D near-field aerodynamic results to the linear acoustic 3-D far-field. The BVI noise mechanisms are explained and the effects of vortex type and strength, and angle of attack are studied. Particularly, airfoil shape modifications which lead to noise reduction are investigated. The results presented are expected to be helpful for better understanding of the nature of the BVI noise and better blade design.

Visualization and analysis of vortex-turbine intersections in wind farms.

PubMed

Shafii, Sohail; Obermaier, Herald; Linn, Rodman; Koo, Eunmo; Hlawitschka, Mario; Garth, Christoph; Hamann, Bernd; Joy, Kenneth I

2013-09-01

Characterizing the interplay between the vortices and forces acting on a wind turbine's blades in a qualitative and quantitative way holds the potential for significantly improving large wind turbine design. This paper introduces an integrated pipeline for highly effective wind and force field analysis and visualization. We extract vortices induced by a turbine's rotation in a wind field, and characterize vortices in conjunction with numerically simulated forces on the blade surfaces as these vortices strike another turbine's blades downstream. The scientifically relevant issue to be studied is the relationship between the extracted, approximate locations on the blades where vortices strike the blades and the forces that exist in those locations. This integrated approach is used to detect and analyze turbulent flow that causes local impact on the wind turbine blade structure. The results that we present are based on analyzing the wind and force field data sets generated by numerical simulations, and allow domain scientists to relate vortex-blade interactions with power output loss in turbines and turbine life expectancy. Our methods have the potential to improve turbine design to save costs related to turbine operation and maintenance.

Top surface blade residues and the central channel water molecules are conserved in every repeat of the integrin-like β-propeller structures.

PubMed

Denesyuk, Alexander; Denessiouk, Konstantin; Johnson, Mark S

2018-02-01

An integrin-like β-propeller domain contains seven repeats of a four-stranded antiparallel β-sheet motif (blades). Previously we described a 3D structural motif within each blade of the integrin-type β-propeller. Here, we show unique structural links that join different blades of the β-propeller structure, which together with the structural motif for a single blade are repeated in a β-propeller to provide the functional top face of the barrel, found to be involved in protein-protein interactions and substrate recognition. We compare functional top face diagrams of the integrin-type β-propeller domain and two non-integrin type β-propeller domains of virginiamycin B lyase and WD Repeat-Containing Protein 5. Copyright © 2017 Elsevier Inc. All rights reserved.

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

NASA Astrophysics Data System (ADS)

Kaneko, Masanao; Tsujita, Hoshio

2015-04-01

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

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.

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

NASA Astrophysics Data System (ADS)

Rule, John Allen

1997-08-01

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

Discrete-frequency and broadband noise radiation from diesel engine cooling fans

NASA Astrophysics Data System (ADS)

Kim, Geon-Seok

This effort focuses on measuring and predicting the discrete-frequency and broadband noise radiated by diesel engine cooling fans. Unsteady forces developed by the interaction of the fan blade with inlet flow are the dominant source for both discrete-frequency and broadband noise of the subject propeller fan. In many cases, a primary source of discrepancy between fan noise prediction and measurement is due to incomplete description of the fan inflow. Particularly, in such engine cooling systems where space is very limited, it would be very difficult, if not, impossible to measure the fan inflow velocity field using the conventional, stationary hot-wire method. Instead, the fan inflow was measured with two-component x-type hot-film probes attached very close to the leading edge of a rotating blade. One of the advantages of the blade-mounted-probe measurement technique is that it measures velocities relative to the rotating probe, which enables the acquired data to be applied directly in many aerodynamic theories that have been developed for the airfoil fixed-coordinate system. However, the velocity time data measured by this technique contains the spatially non-uniform mean velocity field along with the temporal fluctuations. A phase-locked averaging technique was successfully employed to decompose the velocity data into time-invariant flow distortions and fluctuations due to turbulence. The angles of attack of the fan blades, obtained from inlet flow measurements, indicate that the blades are stalled. The fan's radiated noise was measured without contamination from the engine noise by driving the fan with an electric motor. The motor operated at a constant speed while a pair of speed controllable pulleys controlled the fan speed. Narrowband and 1/3-octave band sound power of the cooling fan was measured by using the comparison method with a reference sound source in a reverberant room. The spatially non-uniform mean velocity field was used in axial-flow fan noise theory to predict the discrete-frequency noise at the blade passing frequency (BPF) and harmonics. The unsteady lift was predicted by considering transverse and longitudinal velocity fluctuations. The influences due to an upstream finger guard were also investigated. The radiated sound power spectra that were measured for the fan are shown to be in excellent agreement with those predicted. The agreement between prediction and measurement is only fair at the fundamental BPF tone. Further experimental investigations revealed that the interaction noise between the fan blades and a shroud surrounding the fan might be the dominant source for the radiation at the first harmonic. The space-time correlation functions of the inflow velocity fluctuations were measured and utilized in stochastic lifting surface theory to calculate the unsteady blade lift and resulting broadband fan noise. The integral length scale of the inlet flow was found to be much smaller than the blade-to-blade separate distance of the fan. Therefore, contributions from blade-to-blade correlations of the various elements on different blades were found to be negligible and hence ignored; only the correlations between the strip elements on a given blade were considered. The cross-correlations measured between elements separated by more than the integral length scale were also found to be negligibly small. The predicted broadband sound power spectra agree well with those measured for frequencies greater than 400 Hz. There are deviations between the predictions and measurements at lower frequencies. These are likely due to fan blade stall, and possibly, anomalies in the noise measurement environment. In order to reduce the sound radiation at the blade rate tones, the baseline fan was replaced with a skewed fan. The backward skew angle of 30° was found to effectively reduce the 2nd and higher harmonics of the blade rate tone. The interaction of the shroud opening and the blade tips dominates the sound level at the fundamental tone. This tone was successfully reduced by incorporating a serrated shroud opening. Overall, a 2.8 dB sound power level reduction was achieved by applying the skewed fan and the serrated shroud opening simultaneously. Almost all blade rate tone levels were reduced without adversely affecting the flow performance of the system.

The Spectral and Statistical Properties of Turbulence Generated by a Vortex/Blade-Tip Interaction

NASA Technical Reports Server (NTRS)

Devenport, William J.; Wittmer, Kenneth S.; Wenger, Christian W.

1997-01-01

The perpendicular interaction of a streamwise vortex with the tip of a lifting blade was studied in incompressible flow to provide information useful to the accurate prediction of helicopter rotor noise and the understanding of vortex dominated turbulent flows. The vortex passed 0.3 chord lengths to the suction side of the blade tip, providing a weak interaction. Single and two-point turbulence measurements were made using sub-miniature four sensor hot-wire probes 15 chord lengths downstream of the blade trailing edge; revealing the mean velocity and Reynolds stress tensor distributions of the turbulence, as well as its spanwise length scales as a function of frequency. The single point measurements show the flow downstream of the blade to be dominated by the interaction of the original tip vortex and the vortex shed by the blade. These vortices rotate about each other under their mutual induction, winding up the turbulent wakes of the blades. This interaction between the vortices appears to be the source of new turbulence in their cores and in the region between them. This turbulence appears to be responsible for some decay in the core of the original vortex, not seen when the blade is removed. The region between the vortices is not only a region of comparatively large stresses, but also one of intense turbulence production. Velocity autospectra measured near its center suggests the presence quasi-periodic large eddies with axes roughly parallel to a line joining the vortex cores. Detailed two-point measurements were made on a series of spanwise cuts through the flow so as to reveal the turbulence scales as they would be seen along the span of an intersecting airfoil. The measurements were made over a range of probe separations that enabled them to be analyzed not only in terms of coherence and phase spectra but also in terms of wave-number frequency (kappa-omega) spectra, computed by transforming the measured cross-spectra with respect to the spanwise separation of the probes. These data clearly show the influence of the coherent eddies in the spiral wake and the turbulent region between the cores. These eddies produce distinct peaks in the upwash velocity kappa-omega spectra, and strong anisotropy manifested both in the decay of the kappa-omega spectrum at larger wave-numbers and in differences between the kappa-omega spectra of different components. None of these features are represented in the von Karman spectrum for isotropic turbulence that is often used in broadband noise computations. Wave-number frequency spectra measured in the cores appear to show some evidence that the turbulence outside sets tip core waves, as has previously been hypothesized. These spectra also provide for the first time a truly objective method for distinguishing velocity fluctuations produced by core wandering from other motions.

Unsteady loading of a vertical-axis turbine in the interaction with an upstream deflector

NASA Astrophysics Data System (ADS)

Kim, Daegyoum; Gharib, Morteza

2014-01-01

Torque generation and flow distribution of a lift-based vertical-axis turbine with an upstream deflecting plate are investigated in water tunnel experiments. The deployment of a deflector in front of a lift-based turbine is a promising approach to increase local flow velocity and enhance energy conversion efficiency without consideration for complicated control. For the turbine with the deflector, the phase during which the blade passes near the front end of the turbine has a major contribution to torque increase from the case without the deflector. Meanwhile, the deflector can have a negative effect in torque generation at the phase when the blade moves upstream against free stream if the turbine is placed close to the deflector in a crosswise direction. The change of nearby flow distribution by the deflector is also examined to find its correlation with torque generation. When the blade rotates through the near-wake region of the deflector, the blade can collides with the vortical structure shed from the deflector. This interaction causes significant torque fluctuation.

Simulation of Rotary-Wing Near-Wake Vortex Structures Using Navier-Stokes CFD Methods

NASA Technical Reports Server (NTRS)

Kenwright, David; Strawn, Roger; Ahmad, Jasim; Duque, Earl; Warmbrodt, William (Technical Monitor)

1997-01-01

This paper will use high-resolution Navier-Stokes computational fluid dynamics (CFD) simulations to model the near-wake vortex roll-up behind rotor blades. The locations and strengths of the trailing vortices will be determined from newly-developed visualization and analysis software tools applied to the CFD solutions. Computational results for rotor nearwake vortices will be used to study the near-wake vortex roll up for highly-twisted tiltrotor blades. These rotor blades typically have combinations of positive and negative spanwise loading and complex vortex wake interactions. Results of the computational studies will be compared to vortex-lattice wake models that are frequently used in rotorcraft comprehensive codes. Information from these comparisons will be used to improve the rotor wake models in the Tilt-Rotor Acoustic Code (TRAC) portion of NASA's Short Haul Civil Transport program (SHCT). Accurate modeling of the rotor wake is an important part of this program and crucial to the successful design of future civil tiltrotor aircraft. The rotor wake system plays an important role in blade-vortex interaction noise, a major problem for all rotorcraft including tiltrotors.

OFF-DESIGN PERFORMANCE OF RADIAL INFLOW TURBINES

NASA Technical Reports Server (NTRS)

Wasserbauer, C. A.

1994-01-01

This program calculates off design performance of radial inflow turbines. The program uses a one dimensional solution of flow conditions through the turbine along the main streamline. The loss model accounts for stator, rotor, incidence, and exit losses. Program features include consideration of stator and rotor trailing edge blockage and computation of performance to limiting load. Stator loss (loss in kinetic energy across the stator) is proportional to the average kinetic energy in the blade row and is represented in the program by an equation which includes a stator loss coefficient determined from design point performance and then assumed to be constant for the off design calculations. Minimum incidence loss does not occur at zero incidence angle with respect to the rotor blade, but at some optimum flow angle. At high pressure ratios the level of rotor inlet velocity seemed to have an excessive influence on the loss. Using the component of velocity in the direction of the optimum flow angle gave better correlations with experimental results. Overall turbine geometry and design point values of efficiency, pressure ratio, and mass flow are needed as input information. The output includes performance and velocity diagram parameters for any number of given speeds over a range of turbine pressure ratio. The program has been implemented on the IBM 7094 and operates in batch mode.

A Simple Interactive Program to Design Supercavitating Propeller Blades.

DTIC Science & Technology

1982-06-01

Torque coefficient CT - Thrust-load coefficient c - Blade chord length D - Propeller diameter, assumed Dopt - Propeller diameter, optimum F - Blade force...2 ci) (10) where: C = drag to lift ratio A. W ideal advance ratio. At this point, solutions can be made for the radial pitch, ( D x ir XX), and the...t/ D )0 .7 x D ൏ 1C (13) 0.7 C where: C Ia coefficient of lift fora finite foil. Reference 2 states that the optimum value for C . is 0.16 and that

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

Air-structure coupling features analysis of mining contra-rotating axial flow fan cascade

NASA Astrophysics Data System (ADS)

Chen, Q. G.; Sun, W.; Li, F.; Zhang, Y. J.

2013-12-01

The interaction between contra-rotating axial flow fan blade and working gas has been studied by means of establishing air-structure coupling control equation and combining Computational Fluid Dynamics (CFD) and Computational solid mechanics (CSM). Based on the single flow channel model, the Finite Volume Method was used to make the field discrete. Additionally, the SIMPLE algorithm, the Standard k-ε model and the Arbitrary Lagrangian-Eulerian dynamic grids technology were utilized to get the airflow motion by solving the discrete governing equations. At the same time, the Finite Element Method was used to make the field discrete to solve dynamic response characteristics of blade. Based on weak coupling method, data exchange from the fluid solver and the solid solver was processed on the coupling interface. Then interpolation was used to obtain the coupling characteristics. The results showed that the blade's maximum amplitude was on the tip of the last-stage blade and aerodynamic force signal could reflect the blade working conditions to some extent. By analyzing the flow regime in contra-rotating axial flow fan, it could be found that the vortex core region was mainly in the blade surface, the hub and the blade clearance. In those regions, the turbulence intensity was very high. The last-stage blade's operating life is shorter than that of the pre-stage blade due to the fatigue fracture occurs much more easily on the last-stage blade which bears more stress.

Ducting arrangement for cooling a gas turbine structure

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

Lee, Ching-Pang; Morrison, Jay A.

2015-07-21

A ducting arrangement (10) for a can annular gas turbine engine, including: a duct (12, 14) disposed between a combustor (16) and a first row of turbine blades and defining a hot gas path (30) therein, the duct (12, 14) having raised geometric features (54) incorporated into an outer surface (80); and a flow sleeve (72) defining a cooling flow path (84) between an inner surface (78) of the flow sleeve (72) and the duct outer surface (80). After a cooling fluid (86) traverses a relatively upstream raised geometric feature (90), the inner surface (78) of the flow sleeve (72)more » is effective to direct the cooling fluid (86) toward a landing (94) separating the relatively upstream raised geometric feature (90) from a relatively downstream raised geometric feature (94).« less

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

NASA Technical Reports Server (NTRS)

Stewart, Warner L; Schum, Harold J; Wong, Robert Y

1952-01-01

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

Virtual Reality Used to Serve the Glenn Engineering Community

NASA Technical Reports Server (NTRS)

Carney, Dorothy V.

2001-01-01

There are a variety of innovative new visualization tools available to scientists and engineers for the display and analysis of their models. At the NASA Glenn Research Center, we have an ImmersaDesk, a large, single-panel, semi-immersive display device. This versatile unit can interactively display three-dimensional images in visual stereo. Our challenge is to make this virtual reality platform accessible and useful to researchers. An example of a successful application of this computer technology is the display of blade out simulations. NASA Glenn structural dynamicists, Dr. Kelly Carney and Dr. Charles Lawrence, funded by the Ultra Safe Propulsion Project under Base R&T, are researching blade outs, when turbine engines lose a fan blade during operation. Key objectives of this research include minimizing danger to the aircraft via effective blade containment, predicting destructive loads due to the imbalance following a blade loss, and identifying safe, cost-effective designs and materials for future engines.

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.

Stability and adaptability of soybean cultivars in Minas Gerais.

PubMed

Soares, I O; Bruzi, A T; Zambiazzi, E V; Guilherme, S R; Bianchi, M C; Silva, K B; Fronza, V; Teixeira, C M

2017-08-17

Genotypes x environment (G x E) interaction consists of different behavior of genotypes cultivated in different environments. This interaction occurs due to the performance variation of each genotype in different environments. To reduce the effect of the interaction in soybean crops, some studies have been reported in the literature to study their adaptability and stability. However, these studies are still scarce in Minas Gerais State. Thus, the aim of this study was to verify the adaptability and stability of soybean cultivars and identify the cultivars that contribute least to the G x E interaction in Minas Gerais. Six soybean cultivars were evaluated in 9 different environments. The plots were composed of 4 rows of 5 m with a spacing of 0.5 m between rows, and only the two central rows were harvested. The inoculation with Bradyrhizobium japonicum was performed through sowing furrow. The fertilization followed the recommendations of the Soil Fertility Commission of Minas Gerais. Grain yield was evaluated in kg/ha after conversion to 13% moisture. After individual analysis, the joint analysis was performed by grouping the phenotypic means by the Scott and Knott (1974) test. Wricke's ecovalence methodologies and the Annicchiarico confidence index were applied for the adaptability and stability analysis. The interaction was decomposed into a simple and a complex part. The cultivars BRSMG 820RR and BRSMG 760SRR have wide adaptability and stability. The first one presents a better index of confidence and a small contribution to the interaction.

Fluid Structure Interaction in a Turbine Blade

NASA Technical Reports Server (NTRS)

Gorla, Rama S. R.

2004-01-01

An unsteady, three dimensional Navier-Stokes solution in rotating frame formulation for turbomachinery applications is presented. Casting the governing equations in a rotating frame enabled the freezing of grid motion and resulted in substantial savings in computer time. The turbine blade was computationally simulated and probabilistically evaluated in view of several uncertainties in the aerodynamic, structural, material and thermal variables that govern the turbine blade. The interconnection between the computational fluid dynamics code and finite element structural analysis code was necessary to couple the thermal profiles with the structural design. The stresses and their variations were evaluated at critical points on the Turbine blade. Cumulative distribution functions and sensitivity factors were computed for stress responses due to aerodynamic, geometric, mechanical and thermal random variables.

Mapping Wind Farm Loads and Power Production - A Case Study on Horns Rev 1

NASA Astrophysics Data System (ADS)

Galinos, Christos; Dimitrov, Nikolay; Larsen, Torben J.; Natarajan, Anand; Hansen, Kurt S.

2016-09-01

This paper describes the development of a wind turbine (WT) component lifetime fatigue load variation map within an offshore wind farm. A case study on the offshore wind farm Horns Rev I is conducted with this purpose, by quantifying wake effects using the Dynamic Wake Meandering (DWM) method, which has previously been validated based on CFD, Lidar and full scale load measurements. Fully coupled aeroelastic load simulations using turbulent wind conditions are conducted for all wind directions and mean wind speeds between cut-in and cut-out using site specific turbulence level measurements. Based on the mean wind speed and direction distribution, the representative 20-year lifetime fatigue loads are calculated. It is found that the heaviest loaded WT is not the same when looking at blade root, tower top or tower base components. The blade loads are mainly dominated by the wake situations above rated wind speed and the highest loaded blades are in the easternmost row as the dominating wind direction is from West. Regarding the tower components, the highest loaded WTs are also located towards the eastern central location. The turbines with highest power production are, not surprisingly, the ones facing a free sector towards west and south. The power production results of few turbines are compared with SCADA data. The results of this paper are expected to have significance for operation and maintenance planning, where the schedules for inspection and service activities can be adjusted to the requirements arising from the varying fatigue levels. Furthermore, the results can be used in the context of remaining fatigue lifetime assessment and planning of decommissioning.

Visualization and Analysis of Vortex-Turbine Intersections in Wind Farms.

PubMed

Shafii, Sohail; Obermaier, Harald; Linn, Rodman; Koo, Eunmo; Hlawitschka, Mario; Garth, Christoph; Hamann, Bernd; Joy, Kenneth

2013-02-13

Characterizing the interplay between the vortices and forces acting on a wind turbine's blades in a qualitative and quantitative way holds the potential for significantly improving large wind turbine design. The paper introduces an integrated pipeline for highly effective wind and force field analysis and visualization. We extract vortices induced by a turbine's rotation in a wind field, and characterize vortices in conjunction with numerically simulated forces on the blade surfaces as these vortices strike another turbine's blades downstream. The scientifically relevant issue to be studied is the relationship between the extracted, approximate locations on the blades where vortices strike the blades and the forces that exist in those locations. This integrated approach is used to detect and analyze turbulent flow that causes local impact on the wind turbine blade structure. The results that we present are based on analyzing the wind and force field data sets generated by numerical simulations, and allow domain scientists to relate vortex-blade interactions with power output loss in turbines and turbine life-expectancy. Our methods have the potential to improve turbine design in order to save costs related to turbine operation and maintenance.

Investigation of three-dimensional flow field in a turbine including rotor/stator interaction. I - Design development and performance of the research facility

NASA Technical Reports Server (NTRS)

Lakshminarayana, B.; Camci, C.; Halliwell, I.; Zaccaria, M.

1992-01-01

A description of the Axial Flow Turbine Research Facility (AFTRF) installed at the Turbomachinery Laboratory of the Pennsylvania State University is presented in this paper. The facility diameter is 91.66 cm (3 feet) and the hub-to-tip ratio of the blading is 0.73. The flow path consists of turbulence generating grid, 23 nozzle vane and 29 rotor blades followed by outlet guide vanes. The blading design, carried out by General Electric Company personnel, embody modern HP turbine design philosophy, loading and flow coefficient, reaction, aspect ratio, and blade turning angles; all within the current aircraft engine design turbine practice. State-of-the-art quasi-3D blade design techniques were used to design the vane and the blade shapes. The vanes and blades are heavily instrumented with fast response pressure, shear stress, and velocity probes and have provision for flow visualization and laser Doppler anemometer measurement. Furthermore, provision has been made for detailed nozzle wake, rotor wake and boundary layer surveys. A 150 channel slip ring unit is used for transmitting the rotor data to a stationary instrumentation system. All the design objectives have been met.

Development of advanced blade pitching kinematics for cycloturbines and cyclorotors

NASA Astrophysics Data System (ADS)

Adams, Zachary Howard

Cycloturbines and cyclorotors are established concepts for extracting freesteam fluid energy and producing thrust which promise to exceed the performance of traditional horizontal axis turbines and rotors while maintaining unique operational advantages. However, their potential is not yet realized in widespread applications. A central barrier to their proliferation is the lack of fundamental understanding of the aerodynamic interaction between the turbine and the freestream flow. In particular, blade pitch must be precisely actuated throughout the revolution to achieve the proper blade angle of attack and maximize performance. So far, there is no adequate method for determining or implementing the optimal blade pitching kinematics for cyclorotors or cycloturbines. This dissertation bridges the pitching deficiency by introducing a novel low order model to predict improved pitch kinematics, experimentally demonstrating improved performance, and evaluating flow physics with a high order Navier-Stokes computational code. The foundation for developing advanced blade pitch motions is a low order model named Fluxline Theory. Fluid calculations are performed in a coordinate system fixed to streamlines whose spatial locations are not pre-described in order to capture the flow expansion/contraction and bending through the turbine. A transformation then determines the spatial location of streamlines through the rotor disk and finally blade element method integrations determine the power and forces produced. Validation against three sets of extant cycloturbine experimental data demonstrates improvement over other existing streamtube models. Fluxline Theory was extended by removing dependence on a blade element model to better understand how turbine-fluid interaction impacts thrust and power production. This pure momentum variation establishes a cycloturbine performance limit similar to the Betz Limit for horizontal axis wind turbines, as well as the fluid deceleration required to achieve optimum performance. A novel inverse method was developed implementing a new semi-empirical curvilinear flow blade aerodynamic coefficient model to predict optimum cycloturbine blade pitch waveforms from the ideal fluid deceleration. These improved blade pitch waveforms were evaluated on a 1.37m diameter by 1.37m span cycloturbine to definitively characterize their improvement over existing blade pitch motions and demonstrate the practicality of a variable blade pitch system. The Fluxline Optimal pitching kinematics outperformed sinusoidal and fixed pitching kinematics. The turbine achieved a mean gross aerodynamic power coefficient of 0.44 (95% confidence interval: [0.388,0.490]) and 0.52 (95% confidence interval: [0.426,0.614]) at tip speed ratios (TSRs) of 1.5 and 2.25 respectively which exceeds all other low TSR vertical axis wind turbines. Two-dimensional incompressible Reynolds-averaged Navier-Stokes computational fluid dynamic simulations were used to characterize higher order effects of the blade interaction with the fluid. These simulations suggest Fluxline Optimal pitch kinematics achieve high power coefficients by evenly extracting energy from the flow without blade stall or detached turbine wakes. Fluxline Theory was adapted to inform the design of high efficiency cyclorotors by incorporating the concept of rotor angle of attack as well as a power and drag loss model for blade support structure. A blade element version of this theory predicts rotor performance. For hovering, a simplified variation of the theory instructs that cyclorotors will achieve the greatest power loading at low disk loadings with high solidity blades pitched to maximum lift coefficient. Increasing lift coefficients in the upstream portion of the rotor disproportionately increases performance compared to magnifying lift in the downstream portion. This suggests airfoil sections that counter curvilinear flow effects could improve hovering efficiency. Additionally, the simplified hovering theory explains the cyclorotor side force which was observed experimentally, but never adequately explained. In contrast, a separate simplified version of the theory for high speed forward flight points to better rotor performance with a low solidity, high disk loading rotor operated at high advance ratios. High rotor aspect ratios will improve performance in both hover and forward flight. A new mechanical blade pitch mechanism was designed to actuate the high efficiency blade pitch motions predicted by Fluxline Theory for both cyclorotors and cycloturbines. The mechanism optimizes blade pitch at all operating conditions via different cross sections of a three dimensionally contoured cam. Varying the position of the cam accounts for changing wind direction and velocity on a cycloturbine, or for pilot-controlled thrust vectoring, forward speed, and aircraft angle of attack as a cyclorotor. A simplified variation of the mechanism, which implemented fully aerodynamically-shrouded blade pitch links, performed flawlessly on the cycloturbine experiment.

Effect of applied force and blade speed on histopathology of bone during resection by sagittal saw.

PubMed

James, Thomas P; Chang, Gerard; Micucci, Steven; Sagar, Amrit; Smith, Eric L; Cassidy, Charles

2014-03-01

A sagittal saw is commonly used for resection of bone during joint replacement surgery. During sawing, heat is generated that can lead to an increase in temperature at the resected surface. The aim of this study was to determine the effect of applied thrust force and blade speed on generating heat. The effect of these factors and their interactions on cutting temperature and bone health were investigated with a full factorial Design of Experiments approach for two levels of thrust force, 15 N and 30 N, and for two levels of blade oscillation rate, 12,000 and 18,000 cycles per minute (cpm). In addition, a preliminary study was conducted to eliminate blade wear as a confounding factor. A custom sawing fixture was used to crosscut samples of fresh bovine cortical bone while temperature in the bone was measured by thermocouple (n=40), followed by measurements of the depth of thermal necrosis by histopathological analysis (n=200). An analysis of variance was used to determine the significance of the factor effects on necrotic depth as evidenced by empty lacunae. Both thrust force and blade speed demonstrated a statistically significant effect on the depth of osteonecrosis (p<0.05), while the interaction of thrust force with blade speed was not significant (p=0.22). The minimum necrotic depth observed was 0.50mm, corresponding to a higher level of force and blade speed (30 N, 18,000 cpm). Under these conditions, a maximum temperature of 93°C was measured at 0.3mm from the kerf. With a decrease in both thrust force and blade speed (15N, 12,000 cpm), the temperature in the bone increased to 109°C, corresponding to a nearly 50% increase in depth of the necrotic zone to 0.74 mm. A predictive equation for necrotic depth in terms of thrust force and blade speed was determined through regression analysis and validated by experiment. The histology results imply that an increase in applied thrust force is more effective in reducing the depth of thermal damage to surrounding bone than an increase in blade speed. Copyright © 2013 IPEM. Published by Elsevier Ltd. All rights reserved.

A comparison of cylindrical and row trenched cooling holes with alignment angle of 0 degree near the combustor endwall

NASA Astrophysics Data System (ADS)

Kianpour, E.; Nor Azwadi, C. S.; Golshokouh, I.

2013-12-01

We studied the effects of cylindrical and row trenched cooling holes with alignment angle of 0° at BR=3.18 on the film cooling performance near the endwall surface of a combustor simulator. In this research, a three-dimensional presentation of gas turbine engine was simulated and analyzed with a commercial finite volume package FLUENT 6.2.26 to gain fundamental data. The current study has been performed with Reynolds-averaged Navier-Stokes turbulence model (RANS) on internal cooling passages. This combustor simulator combined the interaction of two rows of dilution jets, which were staggered in the stream wise direction and aligned in the span wise direction. The entire findings of the study declared that with using the row trenched holes near the enwall surface; film cooling effectiveness is doubled compared to the cooling performance of baseline case.

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.

Interaction of impeller and guide vane in a series-designed axial-flow pump

NASA Astrophysics Data System (ADS)

Kim, S.; Choi, Y. S.; Lee, K. Y.; Kim, J. H.

2012-11-01

In this paper, the interaction of the impeller and guide vane in a series-designed axial-flow pump was examined through the implementation of a commercial CFD code. The impeller series design refers to the general design procedure of the base impeller shape which must satisfy the various flow rate and head requirements by changing the impeller setting angle and number of blades of the base impeller. An arc type meridional shape was used to keep the meridional shape of the hub and shroud with various impeller setting angles. The blade angle and the thickness distribution of the impeller were designed as an NACA airfoil type. In the design of the guide vane, it was necessary to consider the outlet flow condition of the impeller with the given setting angle. The meridional shape of the guide vane were designed taking into consideration the setting angle of the impeller, and the blade angle distribution of the guide vane was determined with a traditional design method using vane plane development. In order to achieve the optimum impeller design and guide vane, three-dimensional computational fluid dynamics and the DOE method were applied. The interaction between the impeller and guide vane with different combination set of impeller setting angles and number of impeller blades was addressed by analyzing the flow field of the computational results.

Experimental Study of under-platform Damper Kinematics in Presence of Blade Dynamics

NASA Astrophysics Data System (ADS)

Botto, D.; Gastaldi, C.; Gola, M. M.; Umer, M.

2018-01-01

Among the different devices used in the aerospace industries under-platform dampers are widely used in turbo engines to mitigate the blade vibration. Nevertheless, the damper behaviour is not easy to simulate and engineers have been working in order to improve the accuracy with which theoretical contact models predict the damper behaviour. Majority of the experimental setups collect experimental data in terms of blade amplitude reduction which do not increase the knowledge about the damper dynamics and therefore the uncertainty on the damper behaviour remains a big issue. In this paper, a novel test rig has been purposely designed to accommodate a single blade and two under-platform dampers to deeply investigate the damper-blade interactions. In this test bench, a contact force measuring system was designed to extensively measure the damper contact forces. Damper kinematics is rebuilt by using the relative displacement measured between damper and blade. This paper describes the concept behind the new approach, shows the details of new test rig and discusses experimental results by comparing with previously measured results on an old experimental setup.

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

NASA Technical Reports Server (NTRS)

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

1983-01-01

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

Evaluation of a doubly-swept blade tip for rotorcraft noise reduction

NASA Technical Reports Server (NTRS)

Wake, Brian E.; Egolf, T. Alan

1992-01-01

A computational study was performed for a doubly-swept rotor blade tip to determine its benefit for high-speed impulsive (HSI) and blade-vortex interaction (BVI) noise. This design consists of aft and forward sweep. For the HSI-noise computations, unsteady Euler calculations were performed for several variations to a rotor blade geometry. A doubly-swept planform was predicted to increase the delocalizing Mach number to 0.94 (representative of a 200+ kt helicopter). For the BVI-noise problem, it had been hypothesized that the doubly-swept blade tip, by producing a leading-edge vortex, would reduce the tip-vortex effect on BVI noise. A procedure was used in which the tip vortex velocity profile computed by a Navier-Stokes solver was used to compute the inflow associated with BVI. This inflow was used by a Euler solver to compute the unsteady pressures for an acoustic analysis. The results of this study were inconclusive due to the difficulty in accurately predicting the viscous tip vortex downstream of the blade. Also, for the condition studied, no leading-edge vortex formed at the tip.

Simulation of Aircraft Engine Blade-Out Structural Dynamics

NASA Technical Reports Server (NTRS)

Lawrence, Charles; Carney, Kelly; Gallardo, Vicente

2001-01-01

A primary concern of aircraft structure designers is the accurate simulation of the blade-out event and the subsequent windmilling of the engine. Reliable simulations of the blade-out event are required to insure structural integrity during flight as well as to guarantee successful blade-out certification testing. The system simulation includes the lost blade loadings and the interactions between the rotating turbomachinery and the remaining aircraft structural components. General-purpose finite element structural analysis codes such as MSC NASTRAN are typically used and special provisions are made to include transient effects from the blade loss and rotational effects resulting from the engine's turbomachinery. The present study provides the equations of motion for rotordynamic response including the effect of spooldown speed and rotor unbalance and examines the effects of these terms on a cantilevered rotor. The effect of spooldown speed is found to be greater with increasing spooldown rate. The parametric term resulting from the mass unbalance has a more significant effect on the rotordynamic response than does the spooldown term. The parametric term affects both the peak amplitudes as well as the resonant frequencies of the rotor.

Simulation of Aircraft Engine Blade-Out Structural Dynamics. Revised

NASA Technical Reports Server (NTRS)

Lawrence, Charles; Carney, Kelly; Gallardo, Vicente

2001-01-01

A primary concern of aircraft structure designers is the accurate simulation of the blade-out event and the subsequent windmilling of the engine. Reliable simulations of the blade-out event are required to insure structural integrity during flight as well as to guarantee successful blade-out certification testing. The system simulation includes the lost blade loadings and the interactions between the rotating turbomachinery and the remaining aircraft structural components. General-purpose finite element structural analysis codes such as MSC NASTRAN are typically used and special provisions are made to include transient effects from the blade loss and rotational effects resulting from the engine's turbomachinery. The present study provides the equations of motion for rotordynamic response including the effect of spooldown speed and rotor unbalance and examines the effects of these terms on a cantilevered rotor. The effect of spooldown speed is found to be greater with increasing spooldown rate. The parametric term resulting from the mass unbalance has a more significant effect on the rotordynamic response than does the spooldown term. The parametric term affects both the peak amplitudes as well as the resonant frequencies of the rotor.

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).

A viscous-inviscid interactive compressor calculations

NASA Technical Reports Server (NTRS)

Johnston, W.; Sockol, P. M.

1978-01-01

A viscous-inviscid interactive procedure for subsonic flow is developed and applied to an axial compressor stage. Calculations are carried out on a two-dimensional blade-to-blade region of constant radius assumed to occupy a mid-span location. Hub and tip effects are neglected. The Euler equations are solved by MacCormack's method, a viscous marching procedure is used in the boundary layers and wake, and an iterative interaction scheme is constructed that matches them in a way that incorporates information related to momentum and enthalpy thicknesses as well as the displacement thickness. The calculations are quasi-three-dimensional in the sense that the boundary layer and wake solutions allow for the presence of spanwise (radial) velocities.

HART-II: Prediction of Blade-Vortex Interaction Loading

NASA Technical Reports Server (NTRS)

Lim, Joon W.; Tung, Chee; Yu, Yung H.; Burley, Casey L.; Brooks, Thomas; Boyd, Doug; vanderWall, Berend; Schneider, Oliver; Richard, Hugues; Raffel, Markus

2003-01-01

During the HART-I data analysis, the need for comprehensive wake data was found including vortex creation and aging, and its re-development after blade-vortex interaction. In October 2001, US Army AFDD, NASA Langley, German DLR, French ONERA and Dutch DNW performed the HART-II test as an international joint effort. The main objective was to focus on rotor wake measurement using a PIV technique along with the comprehensive data of blade deflections, airloads, and acoustics. Three prediction teams made preliminary correlation efforts with HART-II data: a joint US team of US Army AFDD and NASA Langley, German DLR, and French ONERA. The predicted results showed significant improvements over the HART-I predicted results, computed about several years ago, which indicated that there has been better understanding of complicated wake modeling in the comprehensive rotorcraft analysis. All three teams demonstrated satisfactory prediction capabilities, in general, though there were slight deviations of prediction accuracies for various disciplines.