14 CFR 27.903 - Engines.

Code of Federal Regulations, 2010 CFR

2010-01-01

... usage. (b) Engine or drive system cooling fan blade protection. (1) If an engine or rotor drive system... fan blade fails. This must be shown by showing that— (i) The fan blades are contained in case of failure; (ii) Each fan is located so that a failure will not jeopardize safety; or (iii) Each fan blade...

14 CFR 27.903 - Engines.

Code of Federal Regulations, 2011 CFR

2011-01-01

... usage. (b) Engine or drive system cooling fan blade protection. (1) If an engine or rotor drive system... fan blade fails. This must be shown by showing that— (i) The fan blades are contained in case of failure; (ii) Each fan is located so that a failure will not jeopardize safety; or (iii) Each fan blade...

14 CFR 27.903 - Engines.

Code of Federal Regulations, 2013 CFR

2013-01-01

... usage. (b) Engine or drive system cooling fan blade protection. (1) If an engine or rotor drive system... fan blade fails. This must be shown by showing that— (i) The fan blades are contained in case of failure; (ii) Each fan is located so that a failure will not jeopardize safety; or (iii) Each fan blade...

An efficient constraint to account for mistuning effects in the optimal design of engine rotors

NASA Technical Reports Server (NTRS)

Murthy, Durbha V.; Pierre, Christophe; Ottarsson, Gisli

1992-01-01

Blade-to-blade differences in structural properties, unavoidable in practice due to manufacturing tolerances, can have significant influence on the vibratory response of engine rotor blade. Accounting for these differences, also known as mistuning, in design and in optimization procedures is generally not possible. This note presents an easily calculated constraint that can be used in design and optimization procedures to control the sensitivity of final designs to mistuning.

75 FR 4774 - Takes of Marine Mammals Incidental to Specified Activities; St. George Reef Light Station...

Federal Register 2010, 2011, 2012, 2013, 2014

2010-01-29

... two-bladed main and tail rotors which are fitted with noise-attenuating blade tip caps that would... disturbances caused by the helicopters rotors and engine. It is likely that the initial helicopter approach to... relatively quiet rotors, and behavioral habituation on the part [[Page 4776

Spin Testing for Durability Began on a Self-Tuning Impact Damper for Turbomachinery Blades

NASA Technical Reports Server (NTRS)

Duffy, Kirsten; Mehmed, Oral

2003-01-01

NASA and Pratt & Whitney will collaborate under a Space Act Agreement to perform spin testing of the impact damper to verify damping effectiveness and durability. Pratt & Whitney will provide the turbine blade and damper hardware for the tests. NASA will provide the facility and perform the tests. Effectiveness and durability will be investigated during and after sustained sweeps of rotor speed through resonance. Tests of a platform wedge damper are also planned to compare its effectiveness with that of the impact damper. Results from baseline tests without dampers will be used to measure damping effectiveness. The self-tuning impact damper combines two damping methods-the tuned mass damper and the impact damper. It consists of a ball located within a cavity in the blade. This ball rolls back and forth on a spherical trough under centrifugal load (tuned mass damper) and can strike the walls of the cavity (impact damper). The ball s rolling natural frequency is proportional to the rotor speed and can be designed to follow an engine-order line (integer multiple of rotor speed). Aerodynamic forcing frequencies typically follow these engineorder lines, and a damper tuned to the engine order will most effectively reduce blade vibrations when the resonant frequency equals the engine-order forcing frequency. This damper has been tested in flat plates and turbine blades in the Dynamic Spin Facility. During testing, a pair of plates or blades rotates in vacuum. Excitation is provided by one of three methods--eddy-current engine-order excitation (ECE), electromechanical shakers, and magnetic bearing excitation. The eddy-current system consists of magnets located circumferentially around the rotor. As a blade passes a magnet, a force is imparted on the blade. The number of magnets used can be varied to change the desired engine order of the excitation. The magnets are remotely raised or lowered to change the magnitude of the force on the blades. The other two methods apply force to the rotating shaft itself at frequencies independent of the rotor speed. During testing, blade vibration is monitored with strain gauges and laser displacement probes.

Helicopter rotor and engine sizing for preliminary performance estimation

NASA Technical Reports Server (NTRS)

Talbot, P. D.; Bowles, J. V.; Lee, H. C.

1986-01-01

Methods are presented for estimating some of the more fundamental design variables of single-rotor helicopters (tip speed, blade area, disk loading, and installed power) based on design requirements (speed, weight, fuselage drag, and design hover ceiling). The well-known constraints of advancing-blade compressibility and retreating-blade stall are incorporated into the estimation process, based on an empirical interpretation of rotor performance data from large-scale wind-tunnel tests. Engine performance data are presented and correlated with a simple model usable for preliminary design. When approximate results are required quickly, these methods may be more convenient to use and provide more insight than large digital computer programs.

The General Electric F404 - Engine of the RAAF’s New Fighter.

DTIC Science & Technology

1985-07-01

turbine stages, high pressure and low pressure, stationary and rotating, are cooled, as well as rotors, cooling plates, blade and vane platforms and...such engine components as turbine rotor blading . disks and seals. This has led to the development of design methods that enable extended usage to...Scientific Adviser RAN Aircraft Maintenance and Flight Trials Unit Directorate of Naval Aircraft Engineering Directorate of Naval Aviation Policy

Updated Assessment of an Open Rotor Airplane Using an Advanced Blade Design

NASA Technical Reports Server (NTRS)

Hendricks, Eric S.; Berton, Jeffrey J.; Haller, William J.; Tong, Michael T.; Guynn, Mark D.

2013-01-01

Application of open rotor propulsion systems (historically referred to as "advanced turboprops" or "propfans") to subsonic transport aircraft received significant attention and research in the 1970s and 1980s when fuel efficiency was the driving focus of aeronautical research. Recent volatility in fuel prices and concern for aviation's environmental impact have renewed interest in open rotor propulsion, and revived research by NASA and a number of engine manufacturers. Over the last few years, NASA has revived and developed analysis capabilities to assess aircraft designs with open rotor propulsion systems. These efforts have been described in several previous papers along with initial results from applying these capabilities. The initial results indicated that open rotor engines have the potential to provide large reductions in fuel consumption and emissions. Initial noise analysis indicated that current noise regulations can be met with modern baseline blade designs. Improved blades incorporating low-noise features are expected to result in even lower noise levels. This paper describes improvements to the initial assessment, plus a follow-on study using a more advanced open rotor blade design to power the advanced singleaisle transport. The predicted performance and environmental results of these two advanced open rotor concepts are presented and compared.

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.

Numerical investigations on axial and radial blade rubs in turbo-machinery

NASA Astrophysics Data System (ADS)

Abdelrhman, Ahmed M.; Tang, Eric Sang Sung; Salman Leong, M.; Al-Qrimli, Haidar F.; Rajamohan, G.

2017-07-01

In the recent years, the clearance between the rotor blades and stator/casing had been getting smaller and smaller prior improving the aerodynamic efficiency of the turbomachines as demand in the engineering field. Due to the clearance reduction between the blade tip and the rotor casing and between rotor blades and stator blades, axial and radial blade rubbing could be occurred, especially at high speed resulting into complex nonlinear vibrations. The primary aim of this study is to address the blade axial rubbing phenomenon using numerical analysis of rotor system. A comparison between rubbing caused impacts of axial and radial blade rubbing and rubbing forces are also aims of this study. Tow rotor models (rotor-stator and rotor casing models) has been designed and sketched using SOILDSWORKS software. ANSYS software has been used for the simulation and the numerical analysis. The rubbing conditions were simulated at speed range of 1000rpm, 1500rpm and 2000rpm. Analysis results for axial blade rubbing showed the appearance of blade passing frequency and its multiple frequencies (lx, 2x 3x etc.) and these frequencies will more excited with increasing the rotational speed. Also, it has been observed that when the rotating speed increased, the rubbing force and the harmonics frequencies in x, y and z-direction become higher and severe. The comparison study showed that axial blade rub is more dangerous and would generate a higher vibration impacts and higher blade rubbing force than radial blade rub.

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

Federal Register 2010, 2011, 2012, 2013, 2014

2010-03-17

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

The response of turbine engine rotors to interference rubs

NASA Technical Reports Server (NTRS)

Kascak, A. F.

1980-01-01

A method was developed for the direct integration of a rotor dynamics system experiencing a blade loss induced rotor rub. Both blade loss and rotor rub were simulated on a rotor typical of a small gas turbine. A small change in the coefficient of friction (from 0.1 to 0.2) caused the rotor to change from forward to backward whirl and to theoretically destroy itself in a few rotations. This method provides an analytical capability to study the susceptibility of rotors to rub induced backward whirl problems.

Multiscale/Multifunctional Probabilistic Composite Fatigue

NASA Technical Reports Server (NTRS)

Chamis, Christos C.

2010-01-01

A multilevel (multiscale/multifunctional) evaluation is demonstrated by applying it to three different sample problems. These problems include the probabilistic evaluation of a space shuttle main engine blade, an engine rotor and an aircraft wing. The results demonstrate that the blade will fail at the highest probability path, the engine two-stage rotor will fail by fracture at the rim and the aircraft wing will fail at 109 fatigue cycles with a probability of 0.9967.

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

Reduction of gas flow nonuniformity in gas turbine engines by means of gas-dynamic methods

NASA Astrophysics Data System (ADS)

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

2017-08-01

Gas flow nonuniformity is one of the main sources of rotor blade vibrations in the gas turbine engines. Usually, the flow circumferential nonuniformity occurs near the annular frames, located in the flow channel of the engine. This leads to the increased dynamic stresses in blades and as a consequence to the blade damage. The goal of the research was to find an acceptable method of reducing the level of gas flow nonuniformity as the source of dynamic stresses in the rotor blades. Two different methods were investigated during this research. Thus, this study gives the ideas about methods of improving the flow structure in gas turbine engine. On the basis of existing conditions (under development or existing engine) it allows the selection of the most suitable method for reducing gas flow nonuniformity.

Effect of Rotor- and Stator-Blade Modifications on Surge Performance of an 11-Stage Axial-Flow Compressor. I - Original Production Compressor of XJ40-WE-6 Engine

NASA Technical Reports Server (NTRS)

Finger, Harold B.; Essig, Robert H.; Conrad, E. William

1952-01-01

An investigation to increase the compressor surge-limit pressure ratio of the XJ40-WE-6 turbojet engine at high equivalent speeds was conducted at the NACA Lewis altitude wind tunnel. This report evaluates the compressor modifications which were restricted to (1) twisting rotor blades (in place) to change blade section angles and (2) inserting new stator diaphragms with different blade angles. Such configuration changes could be incorporated quickly and easily in existing engines at overhaul depots. It was found that slight improvements in the compressor surge limit were possible by compressor blade adjustment. However, some of the modifications also reduced the engine air flow and hence penalized the thrust. The use of a mixer assembly at the compressor outlet improved the surge limit with no appreciable thrust penalty.

Multi-Scale/Multi-Functional Probabilistic Composite Fatigue

NASA Technical Reports Server (NTRS)

Chamis, Christos C.

2008-01-01

A multi-level (multi-scale/multi-functional) evaluation is demonstrated by applying it to three different sample problems. These problems include the probabilistic evaluation of a space shuttle main engine blade, an engine rotor and an aircraft wing. The results demonstrate that the blade will fail at the highest probability path, the engine two-stage rotor will fail by fracture at the rim and the aircraft wing will fail at 109 fatigue cycles with a probability of 0.9967.

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

NASA Technical Reports Server (NTRS)

Cochran, Reeves P.; Dengler, Robert P.

1961-01-01

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

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

NASA Technical Reports Server (NTRS)

Abdul-Aziz, Ali

1996-01-01

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

14 CFR 33.19 - Durability.

Code of Federal Regulations, 2010 CFR

2010-01-01

... design of the compressor and turbine rotor cases must provide for the containment of damage from rotor... outside the compressor and turbine rotor cases must be defined. (b) Each component of the propeller blade... STANDARDS: AIRCRAFT ENGINES Design and Construction; General § 33.19 Durability. (a) Engine design and...

14 CFR 33.19 - Durability.

Code of Federal Regulations, 2012 CFR

2012-01-01

... design of the compressor and turbine rotor cases must provide for the containment of damage from rotor... outside the compressor and turbine rotor cases must be defined. (b) Each component of the propeller blade... STANDARDS: AIRCRAFT ENGINES Design and Construction; General § 33.19 Durability. (a) Engine design and...

14 CFR 33.19 - Durability.

Code of Federal Regulations, 2014 CFR

2014-01-01

... design of the compressor and turbine rotor cases must provide for the containment of damage from rotor... outside the compressor and turbine rotor cases must be defined. (b) Each component of the propeller blade... STANDARDS: AIRCRAFT ENGINES Design and Construction; General § 33.19 Durability. (a) Engine design and...

14 CFR 33.19 - Durability.

Code of Federal Regulations, 2011 CFR

2011-01-01

... design of the compressor and turbine rotor cases must provide for the containment of damage from rotor... outside the compressor and turbine rotor cases must be defined. (b) Each component of the propeller blade... STANDARDS: AIRCRAFT ENGINES Design and Construction; General § 33.19 Durability. (a) Engine design and...

14 CFR 33.19 - Durability.

Code of Federal Regulations, 2013 CFR

2013-01-01

... design of the compressor and turbine rotor cases must provide for the containment of damage from rotor... outside the compressor and turbine rotor cases must be defined. (b) Each component of the propeller blade... STANDARDS: AIRCRAFT ENGINES Design and Construction; General § 33.19 Durability. (a) Engine design and...

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

NASA Technical Reports Server (NTRS)

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

1973-01-01

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

Power turbine ventilation system

NASA Technical Reports Server (NTRS)

Wakeman, Thomas G. (Inventor); Brown, Richard W. (Inventor)

1991-01-01

Air control mechanism within a power turbine section of a gas turbine engine. The power turbine section includes a rotor and at least one variable pitch propulsor blade. The propulsor blade is coupled to and extends radially outwardly of the rotor. A first annular fairing is rotatable with the propulsor blade and interposed between the propulsor blade and the rotor. A second fairing is located longitudinally adjacent to the first fairing. The first fairing and the second fairing are differentially rotatable. The air control mechanism includes a platform fixedly coupled to a radially inner end of the propulsor blade. The platform is generally positioned in a first opening and a first fairing. The platform and the first fairing define an outer space. In a first position corresponding with a first propulsor blade pitch, the platform is substantially conformal with the first fairing. In a second position corresponding with the second propulsor blade pitch, an edge portion of the platform is displaced radially outwardly from the first fairing. When the blades are in the second position and rotating about the engine axis, the displacement of the edge portion with respect to the first fairing allows air to flow from the outer space to the annular cavity.

DFVLR rotorcraft: Construction and engineering

NASA Technical Reports Server (NTRS)

Langer, H. J.

1984-01-01

A helicopter rotor test stand is described. Full scale helicopter components can be tested such as hingeless fiberglass rotors and two blade rotor with flapping hinge, or a hybrid system. The facility is used to test stability, rotor components and downwind components.

Passive damping of composite blades using embedded piezoelectric modules or shape memory alloy wires: a comparative study

NASA Astrophysics Data System (ADS)

Bachmann, F.; de Oliveira, R.; Sigg, A.; Schnyder, V.; Delpero, T.; Jaehne, R.; Bergamini, A.; Michaud, V.; Ermanni, P.

2012-07-01

Emission reduction from civil aviation has been intensively addressed in the scientific community in recent years. The combined use of novel aircraft engine architectures such as open rotor engines and lightweight materials offer the potential for fuel savings, which could contribute significantly in reaching gas emissions targets, but suffer from vibration and noise issues. We investigated the potential improvement of mechanical damping of open rotor composite fan blades by comparing two integrated passive damping systems: shape memory alloy wires and piezoelectric shunt circuits. Passive damping concepts were first validated on carbon fibre reinforced epoxy composite plates and then implemented in a 1:5 model of an open rotor blade manufactured by resin transfer moulding (RTM). A two-step process was proposed for the structural integration of the damping devices into a full composite fan blade. Forced vibration measurements of the plates and blade prototypes quantified the efficiency of both approaches, and their related weight penalty.

Aeroelastic response and stability of tiltrotors with elastically-coupled composite rotor blades. Ph.D. Thesis

NASA Technical Reports Server (NTRS)

Nixon, Mark W.

1993-01-01

There is a potential for improving the performance and aeroelastic stability of tiltrotors through the use of elastically-coupled composite rotor blades. To study the characteristics of tiltrotors with these types of rotor blades it is necessary to formulate a new analysis which has the capabilities of modeling both a tiltrotor configuration and an anisotropic rotor blade. Background for these formulations is established in two preliminary investigations. In the first, the influence of several system design parameters on tiltrotor aeroelastic stability is examined for the high-speed axial flight mode using a newly-developed rigid-blade analysis with an elastic wing finite element model. The second preliminary investigation addresses the accuracy of using a one-dimensional beam analysis to predict frequencies of elastically-coupled highly-twisted rotor blades. Important aspects of the new aeroelastic formulations are the inclusion of a large steady pylon angle which controls tilt of the rotor system with respect to the airflow, the inclusion of elastic pitch-lag coupling terms related to rotor precone, the inclusion of hub-related degrees of freedom which enable modeling of a gimballed rotor system and engine drive-train dynamics, and additional elastic coupling terms which enable modeling of the anisotropic features for both the rotor blades and the tiltrotor wing. Accuracy of the new tiltrotor analysis is demonstrated by a comparison of the results produced for a baseline case with analytical and experimental results reported in the open literature. Two investigations of elastically tailored blades on a baseline tiltrotor are then conducted. One investigation shows that elastic bending-twist coupling of the rotor blade is a very effective means for increasing the flutter velocity of a tiltrotor, and the magnitude of coupling required does not have an adverse effect on performance or blade loads. The second investigation shows that passive blade twist control via elastic extension-twist coupling of the rotor blade has the capability of significantly improving tiltrotor aerodynamic performance. This concept, however, is shown to have, in general, a negative impact on stability characteristics.

Effect of Variable Chord Length on Transonic Axial Rotor Performance Investigated

NASA Technical Reports Server (NTRS)

Suder, Kenneth L.

2002-01-01

During the life of any gas turbine, blade erosion is present, especially for those units that are exposed to unfiltered air, such as aviation turbofan engines. The effect of this erosion is to reduce the blade chord progressively from the midspan to the tip region and to roughen and distort the blade surface. The effects of roughness on rotor performance have been documented by Suder et al. and Roberts. These papers indicate that the penalty for leading-edge roughness and erosion can be significant. Turbofan operators, therefore, restore chord length at routine maintenance intervals to regain performance before deterioration is too severe to salvage blades. As the rotor blades erode, the leading edge becomes rough - blunt and distorted from the nominal shape - and the aerodynamic performance suffers. Nominal performance can be recovered by recontouring the leading edges. This process, which inherently shortens the blade chord, can be used until the blade chord erodes to the stall limit. Below this chord length, which varies among engine-compressor types, a decrease of stall margin is likely. After compressor blade rework that includes leading edge recontouring, the blades have different chord lengths, ranging from blades that are near nominal chord length down to those near the stall chord limit. Furthermore, as blades erode below the stall limit, they must be replaced with new blades that have the full nominal chord length. Consequently, a set of compressor blades with varying chord lengths will be installed into each turbofan engine that goes through a complete maintenance cycle. The question arises, "Does fan or compressor performance depend on the order in which mixed-chord blades are installed into a fan or compressor disk?"

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

NASA Astrophysics Data System (ADS)

Guo, Haotian; Duan, Fajie; Zhang, Jilong

2016-01-01

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

77 FR 4650 - Airworthiness Directives; General Electric Company Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2012-01-31

... blade borescope inspection (BSI) or a failed engine core vibration survey, establishes a new lower life... LPT rotor stage 3 disk removal after a failed HPT blade BSI or a failed engine core vibration survey... engine test cell as part of an engine manual performance run fulfill the vibration survey requirements of...

Development of a Noninterference Technique for Measurement of Turbine Engine Compressor Blade Stress

DTIC Science & Technology

1980-06-01

TECHNIQUE FOR MEASUREMENT OF TURBINE ENGINE COMPRESSOR BLADE STRESS 7 A U T H O R ( s ) P . E. M c C a r t y a n d J . W. Thompson , J r...e a e a ~ and tdentJ~ by b|ock numbe~ A noninterference technique for measuring stress in compressor blades of turbine engines is being developed...43 4 AEDC-TR-79-78 1.0 INTRODUCTION 1.1 BACKGROUND Compressor rotor blades in turbojet engines are subjected to

Transonic Fan/Compressor Rotor Design Study. Volume 4

DTIC Science & Technology

1982-02-01

amd Identify by block number) Fan Aircraft Engines Compressor Blade Thickness Rotor Camber Distribution Aerodesign Throat Margin Aerodynamics 20...COMPRESSOR ROTOR DESIGN STUDY Volume IV D.E. Parker and M.R. Simonson General Electric Company Aircraft Engine Business Group Advanced Technology...Compressor Research Group Chief, Technology Branch FOR THE COMMANDER H. IVAN BUS Director, Turbine Engine Division If your address has changed, if you

Transonic Fan/Compressor Rotor Design Study. Volume 2

DTIC Science & Technology

1982-02-01

Identity by block number) Fan Aircraft Engines Compressor Blade Thickness Rotor Camber Distribution Aerodesign Throat Margin Aerodynamics 20. 1ABSRACT...COMPRESSOR ROTOR DESIGN STUDY Volume II D.E. Parker and M.R. Simonson General Electric Company / Aircraft Engine Business Group Advanced Technology...Research Group Chief, Technology Branch FOR THE COMMANDER H. IVAN BUSH Director, Turbine Engine Division . If your address has changed, if you wish to be

Transient dynamics of a flexible rotor with squeeze film dampers

NASA Technical Reports Server (NTRS)

Buono, D. F.; Schlitzer, L. D.; Hall, R. G., III; Hibner, D. H.

1978-01-01

A series of simulated blade loss tests are reported on a test rotor designed to operate above its second bending critical speed. A series of analyses were performed which predicted the transient behavior of the test rig for each of the blade loss tests. The scope of the program included the investigation of transient rotor dynamics of a flexible rotor system, similar to modern flexible jet engine rotors, both with and without squeeze film dampers. The results substantiate the effectiveness of squeeze film dampers and document the ability of available analytical methods to predict their effectiveness and behavior.

Rotor burst protection program: Statistics on aircraft gas turbine engine rotor failures that occurred in US commercial aviation during 1975

NASA Technical Reports Server (NTRS)

Delucia, R. A.; Mangano, G. J.

1977-01-01

Statistics on gas turbine rotor failures that have occurred in U.S. commercial aviation during 1975 are presented. The compiled data were analyzed to establish: (1) The incidence of rotor failures and the number of contained and uncontained rotor bursts; (2) The distribution of rotor bursts with respect to engine rotor component; i.e., fan, compressor or turbine; (3) The type of rotor fragment (disk, rim or blade) typically generated at burst; (4) The cause of failure; (5) The type of engines involved; and (6) The flight condition at the time of failure.

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

NASA Astrophysics Data System (ADS)

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

2016-08-01

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

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

NASA Technical Reports Server (NTRS)

Mueller, A. W.

1984-01-01

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

Caution: Precision Error in Blade Alignment Results in Faulty Unsteady CFD Simulation

NASA Astrophysics Data System (ADS)

Lewis, Bryan; Cimbala, John; Wouden, Alex

2012-11-01

Turbomachinery components experience unsteady loads at several frequencies. The rotor frequency corresponds to the time for one rotor blade to rotate between two stator vanes, and is normally dominant for rotor torque oscillations. The guide vane frequency corresponds to the time for two rotor blades to pass by one guide vane. The machine frequency corresponds to the machine RPM. Oscillations at the machine frequency are always present due to minor blade misalignments and imperfections resulting from manufacturing defects. However, machine frequency oscillations should not be present in CFD simulations if the mesh is free of both blade misalignment and surface imperfections. The flow through a Francis hydroturbine was modeled with unsteady Reynolds-Averaged Navier-Stokes (URANS) CFD simulations and a dynamic rotating grid. Spectral analysis of the unsteady torque on the rotor blades revealed a large component at the machine frequency. Close examination showed that one blade was displaced by 0 .0001° due to round-off errors during mesh generation. A second mesh without blade misalignment was then created. Subsequently, large machine frequency oscillations were not observed for this mesh. These results highlight the effect of minor geometry imperfections on CFD solutions. This research was supported by a grant from the DoE and a National Defense Science and Engineering Graduate Fellowship.

Preliminary design study of advanced composite blade and hub and nonmechanical control system for the tilt-rotor aircraft. Volume 1: Engineering studies

NASA Technical Reports Server (NTRS)

Alexander, H. R.; Smith, K. E.; Mcveigh, M. A.; Dixon, P. G.; Mcmanus, B. L.

1979-01-01

Composite structures technology is applied in a preliminary design study of advanced technology blades and hubs for the XV-15 tilt rotor research demonstrator aircraft. Significant improvements in XV-15 hover and cruise performance are available using blades designed for compatibility with the existing aircraft, i.e., blade installation would not require modification of the airframe, hub or upper controls. Provision of a low risk nonmechanical control system was also studied, and a development specification is given.

Dynamics of Rotating Multi-component Turbomachinery Systems

NASA Technical Reports Server (NTRS)

Lawrence, Charles

1993-01-01

The ultimate objective of turbomachinery vibration analysis is to predict both the overall, as well as component dynamic response. To accomplish this objective requires complete engine structural models, including multistages of bladed disk assemblies, flexible rotor shafts and bearings, and engine support structures and casings. In the present approach each component is analyzed as a separate structure and boundary information is exchanged at the inter-component connections. The advantage of this tactic is that even though readily available detailed component models are utilized, accurate and comprehensive system response information may be obtained. Sample problems, which include a fixed base rotating blade and a blade on a flexible rotor, are presented.

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

NASA Technical Reports Server (NTRS)

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

1981-01-01

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

The design of an air-cooled metallic high temperature radial turbine

NASA Technical Reports Server (NTRS)

Snyder, Philip H.; Roelke, Richard J.

1988-01-01

Recent trends in small advanced gas turbine engines call for higher turbine inlet temperatures. Advances in radial turbine technology have opened the way for a cooled metallic radial turbine capable of withstanding turbine inlet temperatures of 2500 F while meeting the challenge of high efficiency in this small flow size range. In response to this need, a small air-cooled radial turbine has been designed utilizing internal blade coolant passages. The coolant flow passage design is uniquely tailored to simultaneously meet rotor cooling needs and rotor fabrication constraints. The rotor flow-path design seeks to realize improved aerodynamic blade loading characteristics and high efficiency while satisfying rotor life requirements. An up-scaled version of the final engine rotor is currently under fabrication and, after instrumentation, will be tested in the warm turbine test facility at the NASA Lewis Research Center.

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

NASA Technical Reports Server (NTRS)

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

2000-01-01

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

Influence of cross section variations on the structural behaviour of composite rotor blades

NASA Astrophysics Data System (ADS)

Rapp, Helmut; Woerndle, Rudolf

1991-09-01

A highly sophisticated structural analysis is required for helicopter rotor blades with nonhomogeneous cross sections made from nonisotropic material. Combinations of suitable analytical techniques with FEM-based techniques permit a cost effective and sufficiently accurate analysis of these complicated structures. It is determined that in general the 1D engineering theory of bending combined with 2D theories for determining the cross section properties is sufficient to describe the structural blade behavior.

Assessment of community noise for a medium-range airplane with open-rotor engines

NASA Astrophysics Data System (ADS)

Kopiev, V. F.; Shur, M. L.; Travin, A. K.; Belyaev, I. V.; Zamtfort, B. S.; Medvedev, Yu. V.

2017-11-01

Community noise of a hypothetical medium-range airplane equipped with open-rotor engines is assessed by numerical modeling of the aeroacoustic characteristics of an isolated open rotor with the simplest blade geometry. Various open-rotor configurations are considered at constant thrust, and the lowest-noise configuration is selected. A two-engine medium-range airplane at known thrust of bypass turbofan engines at different segments of the takeoff-landing trajectory is considered, after the replacement of those engines by the open-rotor engines. It is established that a medium-range airplane with two open-rotor engines meets the requirements of Chapter 4 of the ICAO standard with a significant margin. It is shown that airframe noise makes a significant contribution to the total noise of an airplane with open-rotor engines at landing.

Transonic Fan/Compressor Rotor Design Study. Volume 3

DTIC Science & Technology

1982-02-01

KEY WORDS (Continue on revere. old. $1 nocoeoary and identify by block nuvb.,) Fan Aircraft Engines Compressor Blade Thickne)s Rotor Camber...COMPRESSOR ’Q ROTOR DESIGN STUDY Volume III D.E. Parker and M.R. Simonson CZ) General Electric Company Aircraft Engine Business Group Advanced...Compressor Research Group Chief, Technology Branch FOR THE COMMANDER H. WAN BI Director, Turbine Engine Division ŕ *If your address has changed, if you wish

Rocket Engine Turbine Blade Surface Pressure Distributions Experiment and Computations

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

Benchmarking aerodynamic prediction of unsteady rotor aerodynamics of active flaps on wind turbine blades using ranging fidelity tools

NASA Astrophysics Data System (ADS)

Barlas, Thanasis; Jost, Eva; Pirrung, Georg; Tsiantas, Theofanis; Riziotis, Vasilis; Navalkar, Sachin T.; Lutz, Thorsten; van Wingerden, Jan-Willem

2016-09-01

Simulations of a stiff rotor configuration of the DTU 10MW Reference Wind Turbine are performed in order to assess the impact of prescribed flap motion on the aerodynamic loads on a blade sectional and rotor integral level. Results of the engineering models used by DTU (HAWC2), TUDelft (Bladed) and NTUA (hGAST) are compared to the CFD predictions of USTUTT-IAG (FLOWer). Results show fairly good comparison in terms of axial loading, while alignment of tangential and drag-related forces across the numerical codes needs to be improved, together with unsteady corrections associated with rotor wake dynamics. The use of a new wake model in HAWC2 shows considerable accuracy improvements.

Development of a wind energy converter with single blade rotor

NASA Astrophysics Data System (ADS)

Hipp, K.

1984-06-01

Wind energy converters with high tip speed ratio and a capacity of up to 50 kW in a 8.5 /msec wind speed were developed. Units with 12 m diameter rotors were tested. The concept of a cost favorable plant as a high speed engine with a supercritically running one blade rotor (soft bearing), gust balance out, automatic blade adjustment to ensure favorable starting qualities, proves to be a success. The single rectangular blade non-twisted with the profile NACA 23012/18 has no dynamic problems. The application of a centrifugal governor, i.e., vane like a Maxwell slat, operating only by rotation about a fixed hinge axis in order to attain adequate constant rotational speed of the plant, is not satisfactory.

Initial Assessment of Open Rotor Propulsion Applied to an Advanced Single-Aisle Aircraft

NASA Technical Reports Server (NTRS)

Guynn, Mark D.; Berton, Jeffrey J.; Hendricks, Eric S.; Tong, Michael T.; Haller, William J.; Thurman, Douglas R.

2011-01-01

Application of high speed, advanced turboprops, or propfans, to subsonic transport aircraft received significant attention and research in the 1970s and 1980s when fuel efficiency was the driving focus of aeronautical research. Recent volatility in fuel prices and concern for aviation s environmental impact have renewed interest in unducted, open rotor propulsion, and revived research by NASA and a number of engine manufacturers. Unfortunately, in the two decades that have passed since open rotor concepts were thoroughly investigated, NASA has lost experience and expertise in this technology area. This paper describes initial efforts to re-establish NASA s capability to assess aircraft designs with open rotor propulsion. Specifically, methodologies for aircraft-level sizing, performance analysis, and system-level noise analysis are described. Propulsion modeling techniques have been described in a previous paper. Initial results from application of these methods to an advanced single-aisle aircraft using open rotor engines based on historical blade designs are presented. These results indicate open rotor engines have the potential to provide large reductions in fuel consumption and emissions. Initial noise analysis indicates that current noise regulations can be met with old blade designs and modern, noiseoptimized blade designs are expected to result in even lower noise levels. Although an initial capability has been established and initial results obtained, additional development work is necessary to make NASA s open rotor system analysis capability on par with existing turbofan analysis capabilities.

Single rotor turbine engine

DOEpatents

Platts, David A.

2002-01-01

There has been invented a turbine engine with a single rotor which cools the engine, functions as a radial compressor, pushes air through the engine to the ignition point, and acts as an axial turbine for powering the compressor. The invention engine is designed to use a simple scheme of conventional passage shapes to provide both a radial and axial flow pattern through the single rotor, thereby allowing the radial intake air flow to cool the turbine blades and turbine exhaust gases in an axial flow to be used for energy transfer. In an alternative embodiment, an electric generator is incorporated in the engine to specifically adapt the invention for power generation. Magnets are embedded in the exhaust face of the single rotor proximate to a ring of stationary magnetic cores with windings to provide for the generation of electricity. In this alternative embodiment, the turbine is a radial inflow turbine rather than an axial turbine as used in the first embodiment. Radial inflow passages of conventional design are interleaved with radial compressor passages to allow the intake air to cool the turbine blades.

Blade platform seal for ceramic/metal rotor assembly

DOEpatents

Wertz, John L.

1982-01-01

A combination ceramic and metal turbine rotor for use in high temperature gas turbine engines includes a metal rotor disc having a rim with a plurality of circumferentially spaced blade root retention slots therein to receive a plurality of ceramic blades, each including side platform segments thereon and a dovetail configured root slidably received in one of the slots. Adjacent ones of the platform segments including edge portions thereon closely spaced when the blades are assembled to form expansion gaps in an annular flow surface for gas passage through the blades and wherein the assembly further includes a plurality of unitary seal members on the rotor connected to its rim and each including a plurality of spaced, axially extending, flexible fingers that underlie and conform to the edge portions of the platform segments and which are operative at turbine operating temperatures and speeds to distribute loading on the platform segments as the fingers are seated against the underside of the blade platforms to seal the gaps without undesirably stressing thin web ceramic sections of the platform.

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.

Finite element analysis of metal matrix composite blade

NASA Astrophysics Data System (ADS)

Isai Thamizh, R.; Velmurugan, R.; Jayagandhan, R.

2016-10-01

In this work, compressor rotor blade of a gas turbine engine has been analyzed for stress, maximum displacement and natural frequency using ANSYS software for determining its failure strength by simulating the actual service conditions. Static stress analysis and modal analysis have been carried out using Ti-6Al-4V alloy, which is currently used in compressor blade. The results are compared with those obtained using Ti matrix composites reinforced with SiC. The advantages of using metal matrix composites in the gas turbine compressor blades are investigated. From the analyses carried out, it seems that composite rotor blades have lesser mass, lesser tip displacement and lower maximum stress values.

Dynamic response characteristics of dual flow-path integrally bladed rotors

NASA Astrophysics Data System (ADS)

Beck, Joseph A.; Brown, Jeffrey M.; Scott-Emuakpor, Onome E.; Cross, Charles J.; Slater, Joseph C.

2015-02-01

New turbine engine designs requiring secondary flow compression often look to dual flow-path integrally bladed rotors (DFIBRs) since these stages have the ability to perform work on the secondary, or bypassed, flow-field. While analogous to traditional integrally bladed rotor stages, DFIBR designs have many differences that result in unique dynamic response characteristics that must be understood to avoid fatigue. This work investigates these characteristics using reduced-order models (ROMs) that incorporate mistuning through perturbations to blade frequencies. This work provides an alternative to computationally intensive geometric-mistuning approaches for DFIBRs by utilizing tuned blade mode reductions and substructure coupling in cyclic coordinates. Free and forced response results are compared to full finite element model (FEM) solutions to determine if any errors are related to the reduced-order model formulation reduction methods. It is shown that DFIBRs have many more frequency veering regions than their single flow-path integrally blade rotor (IBR) counterparts. Modal families are shown to transition between system, inner-blade, and outer-blade motion. Furthermore, findings illustrate that while mode localization of traditional IBRs is limited to a single or small subset of blades, DFIBRs can have modal energy localized to either an inner- or outer-blade set resulting in many blades responding above tuned levels. Lastly, ROM forced response predictions compare well to full FEM predictions for the two test cases shown.

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.

Strain Gage Signal Interpretation.

DTIC Science & Technology

1986-02-01

blades and vanes in many engines have been collected, played back and examined. The engine types encompass GE’s stable of turbine engines from the small...aeromechanical engineer . 1.3 SUMMARY OF RESULTS Strain gage signals from vibrating rotor blades and vanes were collected, examined, classified, and generalized...turboprops, to turbojets and to the large high bypass turbofan engines . Test conditions include all the phases that are investigated

77 FR 60288 - Airworthiness Directives; Pratt & Whitney (P&W) Division Turbofan Engines

Federal Register 2010, 2011, 2012, 2013, 2014

2012-10-03

... reassembly of the 2nd stage high-pressure turbine (HPT) rotor and 3rd stage LPT rotor at the next HPT and LPT... found. (2) At the next high-pressure turbine (HPT) overhaul, re- assemble the 2nd stage HPT rotor blades... fractures in the low-pressure turbine (LPT) of certain PW4000-94'' and PW4000-100'' turbofan engines. This...

Dynamic modeling and vibration characteristics analysis of the aero-engine dual-rotor system with Fan blade out

NASA Astrophysics Data System (ADS)

Yu, Pingchao; Zhang, Dayi; Ma, Yanhong; Hong, Jie

2018-06-01

Fan Blade Out (FBO) from a running rotor of the turbofan engine will not only introduce the sudden unbalance and inertia asymmetry into the rotor, but also apply large impact load and induce rotor-to-stator rubbing on the rotor, which makes the mass, gyroscopic and stiffness matrixes of the dynamic equation become time-varying and highly nonlinear, consequently leads to the system's complicated vibration. The dynamic analysis of the aero-engine rotor system is one essential requirement of the authorities and is vital to the aero-engine's safety. The paper aims at studying the dynamic responses of the complicated dual-rotor systems at instantaneous and windmilling statuses when FBO event occurs. The physical process and mechanical characteristics of the FBO event are described qualitatively, based on which the dynamic modeling for an aero-engine dual-rotor system is carried out considering several excitations caused by FBO. Meanwhile the transient response during the instantaneous status and steady-state response at the windmilling status are obtained. The results reveal that the sudden unbalance can induce impact load to the rotor, and lead to the sharp increase of the vibration amplitude and reaction force. The rub-impact will apply constraint effects on the rotor and restrict the transient vibration amplitude, while the inertia asymmetry has little influence on the transient response. When the rotor with huge unbalance operates at windmilling status, the rub-impact turns to be the main factor determining the rotor's dynamic behavior, and several potential motion states, such as instable dry whip, intermittent rubbing and synchronous full annular rubbing would happen on certain conditions.

Aeroelastic Analysis for Rotorcraft in Flight or in a Wind Tunnel

NASA Technical Reports Server (NTRS)

Johnson, W.

1977-01-01

An analytical model is developed for the aeroelastic behavior of a rotorcraft in flight or in a wind tunnel. A unified development is presented for a wide class of rotors, helicopters, and operating conditions. The equations of motion for the rotor are derived using an integral Newtonian method, which gives considerable physical insight into the blade inertial and aerodynamic forces. The rotor model includes coupled flap-lag bending and blade torsion degrees of freedom, and is applicable to articulated, hingeless, gimballed, and teetering rotors with an arbitrary number of blades. The aerodynamic model is valid for both high and low inflow, and for axial and nonaxial flight. The rotor rotational speed dynamics, including engine inertia and damping, and the perturbation inflow dynamics are included. For a rotor on a wind-tunnel support, a normal mode representation of the test module, strut, and balance system is used. The aeroelastic analysis for the rotorcraft in flight is applicable to a general two-rotor aircraft, including single main-rotor and tandem helicopter configurations, and side-by-side or tilting proprotor aircraft configurations.

Experimental quiet engine program

NASA Technical Reports Server (NTRS)

Cornell, W. G.

1975-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2013-04-01

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

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

NASA Technical Reports Server (NTRS)

Sitaram, N.; Lakshminarayana, B.

1983-01-01

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

Rotor blade assembly having internal loading features

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

Soloway, Daniel David

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

Engine Evaluation of Advanced Technology Control Components

DTIC Science & Technology

1976-08-01

producer turbine rotor blades. This is a very desirable control feature, because protecting turbine blades from overtemperature is particularly...centrifugal boost stage operating back to back on a common drive shaft that is direct driven through the alter- nator rotor shaft. The main stage is an...computation makes this simple dynamic pumping machine possible. The pen- alty of this simple design is lower overall efficiency as com- pared to a

Reduced-Order Blade Mistuning Analysis Techniques Developed for the Robust Design of Engine Rotors

NASA Technical Reports Server (NTRS)

Min, James B.

2004-01-01

The primary objective of this research program is to develop vibration analysis tools, design tools, and design strategies to significantly improve the safety and robustness of turbine engine rotors. Bladed disks in turbine engines always feature small, random blade-to-blade differences, or mistuning. Mistuning can lead to a dramatic increase in blade forced-response amplitudes and stresses. Ultimately, this results in high-cycle fatigue, which is a major safety and cost concern. In this research program, the necessary steps will be taken to transform a state-of-the-art vibration analysis tool, the Turbo-Reduce forced-response prediction code, into an effective design tool by enhancing and extending the underlying modeling and analysis methods. Furthermore, novel techniques will be developed to assess the safety of a given design. In particular, a procedure will be established for using eigenfrequency curve veerings to identify "danger zones" in the operating conditions--ranges of rotational speeds and engine orders in which there is a great risk that the rotor blades will suffer high stresses. This work also will aid statistical studies of the forced response by reducing the necessary number of simulations. Finally, new strategies for improving the design of rotors will be pursued. Several methods will be investigated, including the use of intentional mistuning patterns to mitigate the harmful effects of random mistuning, and the modification of disk stiffness to avoid reaching critical values of interblade coupling in the desired operating range. Recent research progress is summarized in the following paragraphs. First, significant progress was made in the development of the component mode mistuning (CMM) and static mode compensation (SMC) methods for reduced-order modeling of mistuned bladed disks (see the following figure). The CMM method has been formalized and extended to allow a general treatment of mistuning. In addition, CMM allows individual mode mistuning, which accounts for the realistic effects of local variations in blade properties that lead to different mistuning values for different mode types (e.g., mistuning of the first torsion mode versus the second flexural mode). The accuracy and efficiency of the CMM method and the corresponding Turbo-Reduce code were validated for an example finite element model of a bladed disk.

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

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

Design and development of an active Gurney flap for rotorcraft

NASA Astrophysics Data System (ADS)

Freire Gómez, Jon; Booker, Julian D.; Mellor, Phil H.

2013-03-01

The EU's Green Rotorcraft programme will develop an Active Gurney Flap (AGF) for a full-scale helicopter main rotor blade as part of its `smart adaptive rotor blade' technology demonstrators. AGFs can be utilized to provide a localized and variable lift enhancement on the rotor, enabling a redistribution of loading on the rotor blade around the rotor azimuth. Further advantages include the possibility of using AGFs to allow a rotor speed reduction, which subsequently provides acoustic benefits. Designed to be integrable into a commercial helicopter blade, and thereby capable of withstanding real in-flight centrifugal loading, blade vibrations and aerodynamic loads, the demonstrator is expected to achieve a high technology readiness level (TRL). The AGF will be validated initially by a constant blade section 2D wind tunnel test and latterly by full blade 3D whirl tower testing. This paper presents the methodology adopted for the AGF concept topology selection, based on a series of both qualitative and quantitative performance criteria. Two different AGF candidate mechanisms are compared, both powered by a small commercial electromagnetic actuator. In both topologies, the link between the actuator and the control surface consists of two rotating torque bars, pivoting on flexure bearings. This provides the required reliability and precision, while making the design virtually frictionless. The engineering analysis presented suggests that both candidates would perform satisfactorily in a 2D wind tunnel test, but that equally, both have design constraints which limit their potential to be further taken into a whirl tower test under full scale centrifugal and inertial loads.

Computer-Aided Engineering | Wind | NREL

Science.gov Websites

Computes coupled section properties of composite blades for beam-type models Inputs are the airfoil shape approach BModes Computes coupled mode shapes and frequencies of blades and towers Inputs are the boundary -Coordinate transformation Transforms the cumulative dynamics of spinning rotor blades into the non-rotating

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

NASA Astrophysics Data System (ADS)

Erler, Engin

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

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2009-01-01

A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the all-axial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio of 2.4%. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade span. The compressors conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow conditions, and velocity triangles for each stage.

Compressor Study to Meet Large Civil Tilt Rotor Engine Requirements

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2009-01-01

A vehicle concept study has been made to meet the requirements of the Large Civil Tilt Rotorcraft vehicle mission. A vehicle concept was determined, and a notional turboshaft engine system study was conducted. The engine study defined requirements for the major engine components, including the compressor. The compressor design-point goal was to deliver a pressure ratio of 31:1 at an inlet weight flow of 28.4 lbm/sec. To perform a conceptual design of two potential compressor configurations to meet the design requirement, a mean-line compressor flow analysis and design code were used. The first configuration is an eight-stage axial compressor. Some challenges of the all-axial compressor are the small blade spans of the rear-block stages being 0.28 in., resulting in the last-stage blade tip clearance-to-span ratio of 2.4 percent. The second configuration is a seven-stage axial compressor, with a centrifugal stage having a 0.28-in. impeller-exit blade span. The compressors conceptual designs helped estimate the flow path dimensions, rotor leading and trailing edge blade angles, flow conditions, and velocity triangles for each stage.

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.

Identification of Rotorcraft Structural Dynamics from Flight and Wind Tunnel Data

NASA Technical Reports Server (NTRS)

McKillip, Robert M., Jr.

1997-01-01

Excessive vibration remains one one of the most difficult problems that faces the helicopter industry today, affecting all production helicopters at some phase of their development. Vibrations in rotating structures may arise from external periodic dynamic airloads whose frequencies are are close to the natural frequencies of the rotating system itself. The goal for the structures engineer would thus be to design a structure as free from resonance effects as possible. In the case of a helicopter rotor blade these dynamic loads are a consequence of asymmetric airload distribution on the rotor blade in forward flight, leading to a rich collection of higher harmonic airloads that force rotor and airframe response. Accurate prediction of the dynamic characteristics of a helicopter rotor blade will provide the opportunity to affect in a positive manner noise intensity, vibration level, durability, reliability and operating costs by reducing objectionable frequencies or moving them to a different frequency range and thus providing us with a lower vibration rotor. In fact, the dynamic characteristics tend to define the operating limits of a rotorcraft. As computing power has increased greatly over the last decade, researchers and engineers have turned to analyzing the vibrational characteristics of aerospace structures at the design and development stage of the production of an aircraft. Modern rotor blade construction methods lead to products with low mass and low inherent damping so careful design and analysis is required to avoid resonance and an undesirable dynamic performance. In addition, accurate modal analysis is necessary for several current approaches in elastic system identification and active control.

Unsteady flow and dynamic response analyses for helicopter rotor blades

NASA Technical Reports Server (NTRS)

Bratanow, T.

1979-01-01

Research is presented on helicopter rotor blade vibration and on two and three dimensional analyses of unsteady incompressible viscous flow past oscillating helicopter rotor blades. A summary is presented of the two international research collaborations which resulted from the NASA project: the collaboration under the auspices of NATO between the University of Wisconsin-Milwaukee, University of Brussels, Belgium and the Aerodynamics Research Establishment in Goettingen, West Germany, and the collaboration under the auspices of the National Science Foundation between UWM and the University of Hamburg and the Ship Research Establishment in Hamburg, West Germany. A summary is given of the benefits from the NASA project to UWM, the College of Engineering and Applied Science, and the participants on the project.

Rotor blades for turbine engines

DOEpatents

Piersall, Matthew R; Potter, Brian D

2013-02-12

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

Comparison of Theoretically and Experimentally Determined Effects of Oxide Coatings Supplied by Fuel Additives on Uncooled Turbine-blade Temperature During Transient Turbojet-engine Operation

NASA Technical Reports Server (NTRS)

Schafer, Louis J; Stepka, Francis S; Brown, W Byron

1953-01-01

An analysis was made to permit the calculation of the effectiveness of oxide coatings in retarding the transient heat flow into turbine blades when the combustion gas temperature of a turbojet engine is suddenly changed. The analysis is checked with experimental data obtained from a turbojet engine whose blades were coated with two different coating materials (silicon dioxide and boric oxide) by adding silicone oil and tributyl borate to the engine fuel. The very thin coatings (approximately 0.001 in.) that formed on the blades produced a negligible effect on the turbine-blade transient temperature response. With the analysis discussed here, it was possible to predict the turbine rotor-blade temperature response with a maximum error of 40 F.

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.

Active Control of Jet Engine Inlet Flows

DTIC Science & Technology

2007-03-31

These S-shaped ducts do not provide a direct line of sight to the compressor blades , thus hiding the engine from incoming radar waves. Also, serpentine...circumferential distortion pattern acts as an unsteady forcing function, inducing blade vibration that can result in structural fatigue and failure 3. This...shortcoming occurs when the rotor blades pass through regions of reduced axial velocity (i.e., where the total pressure is low). In these areas, since the

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

Federal Register 2010, 2011, 2012, 2013, 2014

2010-07-01

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

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

NASA Technical Reports Server (NTRS)

Chaderjian, Neal M.

2017-01-01

High-resolution simulations of rotor/vortex-wake interaction for a UH60-A rotor under BVI and dynamic stallconditions were carried out with the OVERFLOW Navier-Stokes code.a. The normal force and pitching moment variation with azimuth angle were in good overall agreementwith flight-test data, similar to other CFD results reported in the literature.b. The wake-grid resolution did not have a significant effect on the rotor-blade airloads. This surprisingresult indicates that a wake grid spacing of (Delta)S=10% ctip is sufficient for engineering airloads predictionfor hover and forward flight. This assumes high-resolution body grids, high-order spatial accuracy, anda hybrid RANS/DDES turbulence model.c. Three-dimensional dynamic stall was found to occur due the presence of blade-tip vortices passing overa rotor blade on the retreating side. This changed the local airfoil angle of attack, causing stall, unlikethe 2D perspective of pure pitch oscillation of the local airfoil section.

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.

Research turbine for high-temperature core engine application. 2: Effect of rotor tip clearance on overall performance

NASA Technical Reports Server (NTRS)

Szanca, E. M.; Behning, F. P.; Schum, H. J.

1974-01-01

A 25.4-cm (10-in) tip diameter turbine was tested to determine the effect of rotor radial tip clearance on turbine overall performance. The test turbine was a half-scale model of a 50.8-cm-(20-in.-) diameter research turbine designed for high-temperature core engine application. The test turbine was fabricated with solid vanes and blades with no provision for cooling air and tested at much reduced inlet conditions. The tests were run at design speed over a range of pressure ratios for three different rotor clearances ranging from 2.3 to 6.7 percent of the annular blade passage height. The results obtained are compared to the results obtained with three other turbines of varying amounts of reaction.

Experimental evaluation of the effect of inlet distortion on compressor blade vibrations

NASA Technical Reports Server (NTRS)

Lubomski, J. F.

1979-01-01

Compressor rotor strain gage data from an engine test conducted with an inlet screen distortion were reduced and analyzed. These data are compared to data obtained from the same engine without inlet pressure distortion to determine the net effect of the distortion on the vibratory response of the compressor blades. The results obtained are presented.

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.

Blade retainer assembly

NASA Technical Reports Server (NTRS)

Luebering, G. W. (Inventor)

1977-01-01

A retaining assembly is provided for locking radially extending blades in a rotor disc associated with a gas turbine engine. The assembly includes a pair of spaced apart lugs axially extending from one side of the disc to form an access gap for insertion of a blade tang into a dovetail slot in the rotor disc. A pair of axially aligned inwardly facing recesses are disposed in the lugs. A retaining member resides in the recesses and extends across the gap to preclude egress of the blade tang from the dovetail slot. The retaining member includes at least one axially extending protrusion adapted to radially overlap and abuttingly engage a radially inwardly facing abutment surface on the lugs.

Rotor Re-Design for the SSME Fuel Flowmeter

NASA Technical Reports Server (NTRS)

Marcu, Bogdan

1999-01-01

The present report describes the process of redesigning a new rotor for the SSME Fuel Flowmeter. The new design addresses the specific requirement of a lower rotor speed which would allow the SSME operation at 1 15% rated power level without reaching a blade excitation by the wakes behind the hexagonal flow straightener upstream at frequencies close to the blade natural frequency. A series of calculations combining fleet flowmeters test data, airfoil fluid dynamics and CFD simulations of flow patterns behind the flowmeter's hexagonal straightener has led to a blade twist design alpha = alpha (radius) targeting a kf constant of 0.8256. The kf constant relates the fuel volume flow to the flowmeter rotor speed, for this particular value 17685 GPM at 3650 RPM. Based on this angle distribution, two actual blade designs were developed. A first design using the same blade airfoil as the original design targeted the new kf value only. A second design using a variable blade chord length and airfoil relative thickness targeted simultaneously the new kf value and an optimum blade design destined to provide smooth and stable operation and a significant increase in the blade natural frequency associated with the first bending mode, such that a comfortable margin could be obtained at 115% RPL. The second design is a result of a concurrent engineering process, during which several iterations were made in order to achieve a targeted blade natural frequency associated with the first bending mode of 1300 Hz. Water flow tests preliminary results indicate a kf value of 0.8179 for the f-irst design, which is within 1% of the target value. The second design rotor shows a natural frequency associated with the first bending mode of 1308 Hz, and a water-flow calibration constant of kf 0.8169.

Blading System and Method For Controlling Structural Vibrations

NASA Technical Reports Server (NTRS)

Nguyen, Nhan (Inventor)

2000-01-01

A new blading system for controlling the structural vibrations in axial-flow compressors, turbines, or fans, as in aircraft engines and like turbomachines including a stator disc and a rotor disc is presented. The rotor disc defines several radial hubs that retain the rotor blading systems. Each blading system includes a blade formed of an airfoil, and a root attachment which is dimensioned to fit within, and to engage a corresponding hub. Viscoelastic dampers are selectively applied to the outer surfaces of the root attachment on which compressive or shear forces are likely to develop, intermediate the root attachment and the hub, for compression therebetween upon rotation of the rotor disc, in order to dampen structural vibrations. One advantage presented by the viscoelastic dampers lies in its simplicity, efficiency, cost effectiveness, and its ability to be retrofitted into existing turbomachines with minor surface treatment of the root attachments. Furthermore, since the dampers are not exposed to the inflowing airstream, they do not affect the aerodynamic performance of the turbomachine. Another feature of the damping system is that it provides a significant source of damping to minimize destructive structural vibrations, thereby increasing the durability of the turbomachine, and reducing acoustic noise accompanying high amplitude vibrations.

Transonic Fan/Compressor Rotor Design Study. Volume 5

DTIC Science & Technology

1982-02-01

Fan Aircraft Engines Compressor Blade Thickness Rotor Camber Distribution Aerodesign Throat Margin Aerodynamics 20. ABStTRACT (Continue n reverse...Technology Branch FOR THE COMNANDER H. IV N BUS Director, Turbine Engine Division A If your address has changed, if you wish to be removed from our...ORGANIZATION NAME AND ADDRESS 10. PROGRAM ELEMENT. PROJECT. TASK General Electric Ctmpany AREA & WORK UNIT NUMBERS Aircraft Engine Business Group Project 2307

Acoustic Liner for Turbomachinery Applications

NASA Technical Reports Server (NTRS)

Huff, Dennis L.; Sutliff, Daniel L.; Jones, Michael G.; Hebsur, Mohan G.

2010-01-01

The purpose of this innovation is to reduce aircraft noise in the communities surrounding airports by significantly attenuating the noise generated by the turbomachinery, and enhancing safety by providing a containment barrier for a blade failure. Acoustic liners are used in today's turbofan engines to reduce noise. The amount of noise reduction from an acoustic liner is a function of the treatment area, the liner design, and the material properties, and limited by the constraints of the nacelle or casement design. It is desirable to increase the effective area of the acoustic treatment to increase noise suppression. Modern turbofan engines use wide-chord rotor blades, which means there is considerable treatment area available over the rotor tip. Turbofan engines require containment over the rotors for protection from blade failure. Traditional methods use a material wrap such as Kevlar integrated with rub strips and sometimes metal layers (sandwiches). It is possible to substitute the soft rub-strip material with an open-cell metallic foam that provides noise-reduction benefits and a sacrificial material in the first layer of the containment system. An open-cell foam was evaluated that behaves like a bulk acoustic liner, serves as a tip rub strip, and can be integrated with a rotor containment system. Foams can be integrated with the fan-containment system to provide sufficient safety margins and increased noise attenuation. The major innovation is the integration of the foam with the containment.

Preliminary Aerodynamic Investigation of Fan Rotor Blade Morphing

NASA Technical Reports Server (NTRS)

Tweedt, Daniel L.

2012-01-01

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

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.

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.

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.

Noise data from tests of a 1.83 meter (6-ft-) diameter variable-pitch 1.2-pressure-ratio fan (QF-9)

NASA Technical Reports Server (NTRS)

Glaser, F. W.; Wazyniak, J. A.; Friedman, R.

1975-01-01

Acoustic and aerodynamic data for a 1.83-meter (6-ft.) diameter fan suitable for a quiet engine for short-takeoff-and-landing (STOL) aircraft are documented. The QF-9 rotor blades had an adjustable pitch feature which provided a means for testing at several rotor blade setting angles, including one for reverse thrust. The fan stage incorporated features for low noise. Far-field noise around the fan was measured without acoustic suppression over a range of operating conditions for six different rotor blade setting angles in the forward thrust configuration, and for one in the reverse configuration. Complete results of one-third-octave band analysis of the data are presented in tabular form. Also included are power spectra, data referred to the source, and sideline perceived noise levels.

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.

Advances in Gammalloy Materials-Processes-Application Technology: Successes, Dilemmas, and Future

NASA Astrophysics Data System (ADS)

Kim, Young-Won; Kim, Sang-Lan

2018-04-01

For the last several years, gamma titanium aluminide ( γ-TiAl)-based alloys, called "gammalloys," in specific alloy-microstructure forms began to be implemented in civil aero-engines as cast or wrought low-pressure turbine (LPT) blades and in select ground vehicle engines as cast turbocharger rotors and wrought exhaust valves. Their operation temperatures are approximately up to 750°C for LPT blades and around 1000°C for turbocharger rotors. This article critically assesses current engineering gammalloys and their limitations and introduces eight strengthening pathways that can be adopted immediately for the development of advanced, higher temperature gammalloys. Intelligent integration of the pathways into the emerging application-specific research and development processes is emphasized as the key to the advancement of the gammalloy technology to the next higher engineering performance levels.

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.

Rotor fragment protection program: Statistics on aircraft gas turbine ngine rotor failures that occurred in U.S. commercial aviation during 1978

NASA Technical Reports Server (NTRS)

Delucia, R. A.; Salvino, J. T.

1981-01-01

This report presents statistical information relating to the number of gas turbine engine rotor failures which occurred in commercial aviation service use. The predominant failure involved blade fragments, 82.4 percent of which were contained. Although fewer rotor rim, disk, and seal failures occurred, 33.3%, 100% and 50% respectively were uncontained. Sixty-five percent of the 166 rotor failures occurred during the takeoff and climb stages of flight.

14 CFR 33.27 - Turbine, compressor, fan, and turbosupercharger rotors.

Code of Federal Regulations, 2011 CFR

2011-01-01

... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; General § 33.27... service. (c) The most critically stressed rotor component (except blades) of each turbine, compressor, and... 14 Aeronautics and Space 1 2011-01-01 2011-01-01 false Turbine, compressor, fan, and...

14 CFR 33.27 - Turbine, compressor, fan, and turbosupercharger rotors.

Code of Federal Regulations, 2010 CFR

2010-01-01

... TRANSPORTATION AIRCRAFT AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES Design and Construction; General § 33.27... service. (c) The most critically stressed rotor component (except blades) of each turbine, compressor, and... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Turbine, compressor, fan, and...

Tuned mass damper for integrally bladed turbine rotor

NASA Technical Reports Server (NTRS)

Marra, John J. (Inventor)

1994-01-01

The invention is directed to a damper ring for damping the natural vibration of the rotor blades of an integrally bladed rocket turbine rotor. The invention consists of an integral damper ring which is fixed to the underside of the rotor blade platform of a turbine rotor. The damper ring includes integral supports which extend radially outwardly therefrom. The supports are located adjacent to the base portion and directly under each blade of the rotor. Vibration damping is accomplished by action of tuned mass damper beams attached at each end to the supports. These beams vibrate at a predetermined frequency during operation. The vibration of the beams enforce a local node of zero vibratory amplitude at the interface between the supports and the beam. The vibration of the beams create forces upon the supports which forces are transmitted through the rotor blade mounting platform to the base of each rotor blade. When these forces attain a predetermined design frequency and magnitude and are directed to the base of the rotor blades, vibration of the rotor blades is effectively counteracted.

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.

Laser velocimeter measurements of the flowfield generated by an advanced counterrotating propeller

NASA Technical Reports Server (NTRS)

Podboy, Gary G.; Krupar, Martin J.

1989-01-01

Results are presented of an investigation to measure the flowfield generated by an advanced counterrotating pusher propeller model similar to the full-scale Unducted Fan demonstrator engine. A laser Doppler velocimeter was used to measure the velocity field in several planes normal to the centerline of the model at axial stations upstream and downstream of each rotor. During this investigation, blades of the F4/A4 type were installed on the model which was operating in a freestream Mach 0.72 regime, with the advance ratio of each rotor set at 2.80. The measured data indicate only a slight influence of the potential field of each front rotor blade on the flowfield upstream of the rotor. The data measured downstream of the front rotor characterize the tip vortices, vortex sheets and potential field nonuniformities generated by the front rotor. The unsteadiness of the flow in the rotating frame of reference of the aft rotor is also illustrated.

Blade loss transient dynamics analysis with flexible bladed disk

NASA Technical Reports Server (NTRS)

Gallardo, V. C.; Black, G.; Bach, L.; Cline, S.; Storace, A.

1983-01-01

The transient dynamic response of a flexible bladed disk on a flexible rotor in a two rotor system is formulated by modal synthesis and a Lagrangian approach. Only the nonequilibrated one diameter flexible mode is considered for the flexible bladed disk, while the two flexible rotors are represented by their normal modes. The flexible bladed disk motion is modeled as a combination of two one diameter standing waves, and is coupled inertially and gyroscopically to the flexible rotors. Application to a two rotor model shows that a flexible bladed disk on one rotor can be driven into resonance by an unbalance in the other rotor, and at a frequency equal to the difference in the rotor speeds.

Drive system for the retraction/extension of variable diameter rotor systems

NASA Technical Reports Server (NTRS)

Gmirya, Yuriy (Inventor)

2003-01-01

A drive system for a variable diameter rotor (VDR) system includes a plurality of rotor blade assemblies with inner and outer rotor blade segments. The outer blade segment being telescopically mounted to the inner blade segment. The VDR retraction/extension system includes a drive housing mounted at the root of each blade. The housing supports a spool assembly, a harmonic gear set and an electric motor. The spool assembly includes a pair of counter rotating spools each of which drive a respective cable which extends through the interior of the inboard rotor blade section and around a pulley mounted to the outboard rotor blade section. In operation, the electric motor drives the harmonic gear set which rotates the counter rotating spools. Rotation of the spools causes the cables to be wound onto or off their respective spool consequently effecting retraction/extension of the pulley and the attached outboard rotor blade section relative the inboard rotor blade section. As each blade drive system is independently driven by a separate electrical motor, each independent VDR blade assembly is independently positionable.

CSM solutions of rotating blade dynamics using integrating matrices

NASA Technical Reports Server (NTRS)

Lakin, William D.

1992-01-01

The dynamic behavior of flexible rotating beams continues to receive considerable research attention as it constitutes a fundamental problem in applied mechanics. Further, beams comprise parts of many rotating structures of engineering significance. A topic of particular interest at the present time involves the development of techniques for obtaining the behavior in both space and time of a rotor acted upon by a simple airload loading. Most current work on problems of this type use solution techniques based on normal modes. It is certainly true that normal modes cannot be disregarded, as knowledge of natural blade frequencies is always important. However, the present work has considered a computational structural mechanics (CSM) approach to rotor blade dynamics problems in which the physical properties of the rotor blade provide input for a direct numerical solution of the relevant boundary-and-initial-value problem. Analysis of the dynamics of a given rotor system may require solution of the governing equations over a long time interval corresponding to many revolutions of the loaded flexible blade. For this reason, most of the common techniques in computational mechanics, which treat the space-time behavior concurrently, cannot be applied to the rotor dynamics problem without a large expenditure of computational resources. By contrast, the integrating matrix technique of computational mechanics has the ability to consistently incorporate boundary conditions and 'remove' dependence on a space variable. For problems involving both space and time, this feature of the integrating matrix approach thus can generate a 'splitting' which forms the basis of an efficient CSM method for numerical solution of rotor dynamics problems.

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.

Energy efficient engine: Fan test hardware detailed design report

NASA Technical Reports Server (NTRS)

Sullivan, T. J.

1980-01-01

A single stage fan and quarter stage booster were designed for the energy efficient engine. The fan has an inlet radius ratio of 0.342 and a specific flow rate of 208.9 Kg/S sq m (42.8 lbm/sec sq ft). The fan rotor has 32 medium aspect ratio (2.597) titanium blades with a partspan shroud at 55% blade height. The design corrected fan tip speed is 411.5 M/S (1350 ft/sec). The quarter stage island splits the total fan flow with approximately 22% of the flow being supercharged by the quarter stage rotor. The fan bypass ratio is 6.8. The core flow total pressure ratio is 1.67 and the fan bypass pressure ratio is 1.65. The design details of the fan and booster blading, and the fan frame and static structure for the fan configuration are presented.

Investigation of rotor blade element airloads for a teetering rotor in the blade stall regime

NASA Technical Reports Server (NTRS)

Dadone, L. U.; Fukushima, T.

1974-01-01

A model of a teetering rotor was tested in a low speed wind tunnel. Blade element airloads measured on an articulated model rotor were compared with the teetering rotor and showed that the teetering rotor is subjected to less extensive flow separation. Retreating blade stall was studied. Results show that stall, under the influence of unsteady aerodynamic effects, consists of four separate stall events, each associated with a vortex shed from the leading edge and sweeping over the upper surface of the rotor blade. Current rotor performance prediction methodology was evaluated through computer simulation.

Wavelet-based identification of rotor blades in passage-through-resonance tests

NASA Astrophysics Data System (ADS)

Carassale, Luigi; Marrè-Brunenghi, Michela; Patrone, Stefano

2018-01-01

Turbine blades are critical components in turbo engines and their design process usually includes experimental tests in order to validate and/or update numerical models. These tests are generally carried out on full-scale rotors having some blades instrumented with strain gauges and usually involve a run-up or a run-down phase. The quantification of damping in these conditions is rather challenging for several reasons. In this work, we show through numerical simulations that the usual identification procedures lead to a systematic overestimation of damping due both to the finite sweep velocity, as well as to the variation of the blade natural frequencies with the rotation speed. To overcome these problems, an identification procedure based on the continuous wavelet transform is proposed and validated through numerical simulation.

49 CFR 830.5 - Immediate notification.

Code of Federal Regulations, 2011 CFR

2011-10-01

... any internal turbine engine component that results in the escape of debris other than out the exhaust... utilized. (8) Release of all or a portion of a propeller blade from an aircraft, excluding release caused... airspace. (11) Damage to helicopter tail or main rotor blades, including ground damage, that requires major...

49 CFR 830.5 - Immediate notification.

Code of Federal Regulations, 2014 CFR

2014-10-01

... any internal turbine engine component that results in the escape of debris other than out the exhaust... utilized. (8) Release of all or a portion of a propeller blade from an aircraft, excluding release caused... airspace. (11) Damage to helicopter tail or main rotor blades, including ground damage, that requires major...

49 CFR 830.5 - Immediate notification.

Code of Federal Regulations, 2012 CFR

2012-10-01

... any internal turbine engine component that results in the escape of debris other than out the exhaust... utilized. (8) Release of all or a portion of a propeller blade from an aircraft, excluding release caused... airspace. (11) Damage to helicopter tail or main rotor blades, including ground damage, that requires major...

49 CFR 830.5 - Immediate notification.

Code of Federal Regulations, 2013 CFR

2013-10-01

... any internal turbine engine component that results in the escape of debris other than out the exhaust... utilized. (8) Release of all or a portion of a propeller blade from an aircraft, excluding release caused... airspace. (11) Damage to helicopter tail or main rotor blades, including ground damage, that requires major...

On the impact of multi-axial stress states on trailing edge bondlines in wind turbine rotor blades

NASA Astrophysics Data System (ADS)

Noever Castelos, Pablo; Balzani, Claudio

2016-09-01

For a reliable design of wind turbine systems all of their components have to be designed to withstand the loads appearing in the turbine's lifetime. When performed in an integral manner this is called systems engineering, and is exceptionally important for components that have an impact on the entire wind turbine system, such as the rotor blade. Bondlines are crucial subcomponents of rotor blades, but they are not much recognized in the wind energy research community. However, a bondline failure can lead to the loss of a rotor blade, and potentially of the entire turbine, and is extraordinarily relevant to be treated with strong emphasis when designing a wind turbine. Modern wind turbine rotor blades with lengths of 80 m and more offer a degree of flexibility that has never been seen in wind energy technology before. Large deflections result in high strains in the adhesive connections, especially at the trailing edge. The latest edition of the DNV GL guideline from end of 2015 demands a three-dimensional stress analysis of bondlines, whereas before an isolated shear stress proof was sufficient. In order to quantify the lack of safety from older certification guidelines this paper studies the influence of multi-axial stress states on the ultimate and fatigue load resistance of trailing edge adhesive bonds. For this purpose, detailed finite element simulations of the IWES IWT-7.5-164 reference wind turbine blades are performed. Different yield criteria are evaluated for the prediction of failure and lifetime. The results show that the multi-axial stress state is governed by span-wise normal stresses. Those are evidently not captured in isolated shear stress proofs, yielding non-conservative estimates of lifetime and ultimate load resistance. This finding highlights the importance to include a three-dimensional stress state in the failure analysis of adhesive bonds in modern wind turbine rotor blades, and the necessity to perform a three-dimensional characterization of adhesive materials.

Comparison of composite rotor blade models: A coupled-beam analysis and an MSC/NASTRAN finite-element model

NASA Technical Reports Server (NTRS)

Hodges, Robert V.; Nixon, Mark W.; Rehfield, Lawrence W.

1987-01-01

A methodology was developed for the structural analysis of composite rotor blades. This coupled-beam analysis is relatively simple to use compared with alternative analysis techniques. The beam analysis was developed for thin-wall single-cell rotor structures and includes the effects of elastic coupling. This paper demonstrates the effectiveness of the new composite-beam analysis method through comparison of its results with those of an established baseline analysis technique. The baseline analysis is an MSC/NASTRAN finite-element model built up from anisotropic shell elements. Deformations are compared for three linear static load cases of centrifugal force at design rotor speed, applied torque, and lift for an ideal rotor in hover. A D-spar designed to twist under axial loading is the subject of the analysis. Results indicate the coupled-beam analysis is well within engineering accuracy.

Flapping inertia for selected rotor blades

NASA Technical Reports Server (NTRS)

Berry, John D.; May, Matthew J.

1991-01-01

Aerodynamics of helicopter rotor systems cannot be investigated without consideration for the dynamics of the rotor. One of the principal properties of the rotor which affects the rotor dynamics is the inertia of the rotor blade about its root attachment. Previous aerodynamic investigation have been performed on rotor blades with a variety of planforms to determine the performance differences due to blade planform. The blades tested for this investigation have been tested on the U.S. Army 2 meter rotor test system (2MRTS) in the NASA Langley 14 by 22 foot subsonic tunnel for hover performance. This investigation was intended to provide fundamental information on the flapping inertia of five rotor blades with differing planforms. The inertia of the bare cuff and the cuff with a blade extension were also measured for comparison with the inertia of the blades. Inertia was determined using a swing testing technique, using the period of oscillation to determine the effective flapping inertia. The effect of damping in the swing test was measured and described. A comparison of the flapping inertials for rectangular and tapered planform blades of approximately the same mass showed the tapered blades to have a lower inertia, as expected.

Durability of zirconia thermal-barrier ceramic coatings on air-cooled turbine blades in cyclic jet engine operation

NASA Technical Reports Server (NTRS)

Liebert, C. H.; Jacobs, R. E.; Stecura, S.; Morse, C. R.

1976-01-01

Thermal barrier ceramic coatings of stabilized zirconia over a bond coat of Ni Cr Al Y were tested for durability on air cooled turbine rotor blades in a research turbojet engine. Zirconia stabilized with either yttria, magnesia, or calcia was investigated. On the basis of durability and processing cost, the yttria stabilized zirconia was considered the best of the three coatings investigated.

Air cooling of disk of a solid integrally cast turbine rotor for an automotive gas turbine

NASA Technical Reports Server (NTRS)

Gladden, H. J.

1977-01-01

A thermal analysis is made of surface cooling of a solid, integrally cast turbine rotor disk for an automotive gas turbine engine. Air purge and impingement cooling schemes are considered and compared with an uncooled reference case. Substantial reductions in blade temperature are predicted with each of the cooling schemes studied. It is shown that air cooling can result in a substantial gain in the stress-rupture life of the blade. Alternatively, increases in the turbine inlet temperature are possible.

Prediction of Unshsrouded Rotor Blade Tip Heat Transfer

NASA Technical Reports Server (NTRS)

Ameri, A. A.; Steinthorsson, E.

1994-01-01

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

Psychoacoustic Testing of Modulated Blade Spacing for Main Rotors

NASA Technical Reports Server (NTRS)

Edwards, Bryan; Booth, Earl R., Jr. (Technical Monitor)

2002-01-01

Psychoacoustic testing of simulated helicopter main rotor noise is described, and the subjective results are presented. The objective of these tests was to evaluate the potential acoustic benefits of main rotors with modulated (uneven) blade spacing. Sound simulations were prepared for six main rotor configurations. A baseline 4-blade main rotor with regular blade spacing was based on the Bell Model 427 helicopter. A 5-blade main rotor with regular spacing was designed to approximate the performance of the 427, but at reduced tipspeed. Four modulated rotors - one with "optimum" spacing and three alternate configurations - were derived from the 5 bladed regular spacing rotor. The sounds were played to 2 subjects at a time, with care being taken in the speaker selection and placement to ensure that the sounds were identical for each subject. A total of 40 subjects participated. For each rotor configuration, the listeners were asked to evaluate the sounds in terms of noisiness. The test results indicate little to no "annoyance" benefit for the modulated blade spacing. In general, the subjects preferred the sound of the 5-blade regular spaced rotor over any of the modulated ones. A conclusion is that modulated blade spacing is not a promising design feature to reduce the annoyance for helicopter main rotors.

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

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

NASA Technical Reports Server (NTRS)

Sekula, Martin K; Wilbur, Matthew L.

2014-01-01

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

Stall induced instability of a teetered rotor

NASA Astrophysics Data System (ADS)

Glasgow, J. C.; Corrigan, R. D.

Recent tests on the 38m Mod-0 horizontal experimental wind turbine yielded quantitative information on stall induced instability of a teetered rotor. Tests were conducted on rotor blades with NACA 230 series and NACA 643-618 airfoils at low rotor speeds to produce high angles of attack at relatively low wind speeds and power levels. The behavior of the rotor shows good agreement with predicted rotor response based on blade angle of attack calculations and airfoil section properties. The untwisted blades with the 64 series airfoil sections had a slower rate of onset of rotor instability when compared with the twisted 230 series blades, but high teeter angles and teeter stop impacts were experienced with both rotors as wind speeds increased to produce high angles of attack on the outboard portion of the blade. The relative importance of blade twist and airfoil section stall characteristics on the rate of onset of rotor unstability with increasing wind speed was not established however. Blade pitch was shown to be effective in eliminating rotor instability at the expense of some loss in rotor performance near rated wind speed.

Projection Moire Interferometry for Rotorcraft Applications: Deformation Measurements of Active Twist Rotor Blades

NASA Technical Reports Server (NTRS)

Fleming, Gary A.; Soto, Hector L.; South, Bruce W.

2002-01-01

Projection Moire Interferometry (PMI) has been used during wind tunnel tests to obtain azimuthally dependent blade bending and twist measurements for a 4-bladed Active Twist Rotor (ATR) system in simulated forward flight. The ATR concept offers a means to reduce rotor vibratory loads and noise by using piezoelectric active fiber composite actuators embedded in the blade structure to twist each blade as they rotate throughout the rotor azimuth. The twist imparted on the blades for blade control causes significant changes in blade loading, resulting in complex blade deformation consisting of coupled bending and twist. Measurement of this blade deformation is critical in understanding the overall behavior of the ATR system and the physical mechanisms causing the reduction in rotor loads and noise. PMI is a non-contacting, video-based optical measurement technique capable of obtaining spatially continuous structural deformation measurements over the entire object surface within the PMI system field-of-view. When applied to rotorcraft testing, PMI can be used to measure the azimuth-dependent blade bending and twist along the full span of the rotor blade. This paper presents the PMI technique as applied to rotorcraft testing, and provides results obtained during the ATR tests demonstrating the PMI system performance. PMI measurements acquired at select blade actuation conditions generating minimum and maximum rotor loads are provided to explore the interrelationship between rotor loads, blade bending, and twist.

Wind-tunnel evaluation of an advanced main-rotor blade design for a utility-class helicopter

NASA Technical Reports Server (NTRS)

Yeager, William T., Jr.; Mantay, Wayne R.; Wilbur, Matthew L.; Cramer, Robert G., Jr.; Singleton, Jeffrey D.

1987-01-01

An investigation was conducted in the Langley Transonic Dynamics Tunnel to evaluate differences between an existing utility-class main-rotor blade and an advanced-design main-rotor blade. The two rotor blade designs were compared with regard to rotor performance oscillatory pitch-link loads, and 4-per-rev vertical fixed-system loads. Tests were conducted in hover and over a range of simulated full-scale gross weights and density altitude conditions at advance ratios from 0.15 to 0.40. Results indicate that the advanced blade design offers performance improvements over the baseline blade in both hover and forward flight. Pitch-link oscillatory loads for the baseline rotor were more sensitive to the test conditions than those of the advanced rotor. The 4-per-rev vertical fixed-system load produced by the advanced blade was larger than that produced by the baseline blade at all test conditions.

Gas-Dynamic Methods to Reduce Gas Flow Nonuniformity from the Annular Frames of Gas Turbine Engines

NASA Astrophysics Data System (ADS)

Kolmakova, D.; Popov, G.

2018-01-01

Gas flow nonuniformity is one of the main sources of rotor blade vibrations in the gas turbine engines. Usually, the flow circumferential nonuniformity occurs near the annular frames, located in the flow channel of the engine. This leads to the increased dynamic stresses in blades and consequently to the blade damage. The goal of the research was to find an acceptable method of reducing the level of gas flow nonuniformity. Two different methods were investigated during this research. Thus, this study gives the ideas about methods of improving the flow structure in gas turbine engine. Based on existing conditions (under development or existing engine) it allows the selection of the most suitable method for reducing gas flow nonuniformity.

Demonstration of an elastically coupled twist control concept for tilt rotor blade application

NASA Technical Reports Server (NTRS)

Lake, R. C.; Nixon, M. W.; Wilbur, M. L.; Singleton, J. D.; Mirick, P. H.

1994-01-01

The purpose of this Note is to present results from an analytic/experimental study that investigated the potential for passively changing blade twist through the use of extension-twist coupling. A set of composite model rotor blades was manufactured from existing blade molds for a low-twist metal helicopter rotor blade, with a view toward establishing a preliminary proof concept for extension-twist-coupled rotor blades. Data were obtained in hover for both a ballasted and unballasted blade configuration in sea-level atmospheric conditions. Test data were compared with results obtained from a geometrically nonlinear analysis of a detailed finite element model of the rotor blade developed in MSC/NASTRAN.

Diffusion bonded boron/aluminum spar-shell fan blade

NASA Technical Reports Server (NTRS)

Carlson, C. E. K.; Cutler, J. L.; Fisher, W. J.; Memmott, J. V. W.

1980-01-01

Design and process development tasks intended to demonstrate composite blade application in large high by-pass ratio turbofan engines are described. Studies on a 3.0 aspect radio space and shell construction fan blade indicate a potential weight savings for a first stage fan rotor of 39% when a hollow titanium spar is employed. An alternate design which featured substantial blade internal volume filled with titanium honeycomb inserts achieved a 14% potential weight savings over the B/M rotor system. This second configuration requires a smaller development effort and entails less risk to translate a design into a successful product. The feasibility of metal joining large subsonic spar and shell fan blades was demonstrated. Initial aluminum alloy screening indicates a distinct preference for AA6061 aluminum alloy for use as a joint material. The simulated airfoil pressings established the necessity of rigid air surfaces when joining materials of different compressive rigidities. The two aluminum alloy matrix choices both were successfully formed into blade shells.

High Bypass Turbofan Component Development. Phase II. Fan Detail Design.

DTIC Science & Technology

1979-12-01

Vane metal angles ........ ..................... ... 18 22 Vane conical airfoil sections ..... ............... ... 19 23 Principal blade stresses at...31.25 deg. The number of rotor airfoils is 20 while the stator has 42 vanes . The number of vanes and the vane - blade spacing were consequences of...effect of radius change are accounted for. Figure 16 shows the blade hub, mean, and tip conical airfoil sections in engine orientation. For

78 FR 37152 - Airworthiness Directives; Bell Helicopter Textron Canada Limited (Bell) Helicopters

Federal Register 2010, 2011, 2012, 2013, 2014

2013-06-20

... rotor (M/R) blade spar space to determine whether it is oversized and reidentifying the blade and... Blvd., Room 663, Fort Worth, Texas 76137. FOR FURTHER INFORMATION CONTACT: Sharon Miles, Aerospace... Management Group, FAA, may approve AMOCs for this AD. Send your proposal to: Sharon Miles, Aerospace Engineer...

Development and application of a method for predicting rotor free wake positions and resulting rotor blade air loads. Volume 1: Model and results

NASA Technical Reports Server (NTRS)

Sadler, S. G.

1971-01-01

Rotor wake geometries are predicted by a process similar to the startup of a rotor in a free stream. An array of discrete trailing and shed vortices is generated with vortex strengths corresponding to stepwise radial and azimuthal blade circulations. The array of shed and trailing vortices is limited to an arbitrary number of azimuthal steps behind each blade. The remainder of the wake model of each blade is an arbitrary number of trailing vortices. Vortex element end points were allowed to be transported by the resultant velocity of the free stream and vortex-induced velocities. Wake geometry, wake flow, and wake-induced velocity influence coefficients are generated by this program for use in the blade loads portion of the calculations. Blade loads computations include the effects of nonuniform inflow due to a free wake, nonlinear airfoil characteristics, and response of flexible blades to the applied loads. Computed wake flows and blade loads are compared with experimentally measured data. Predicted blade loads, response and shears and moments are obtained for a model rotor system having two independent rotors. The effects of advance ratio, vertical separation of rotors, different blade radius ratios, and different azimuthal spacing of the blades of one rotor with respect to the other are investigated.

Probabilistic evaluation of uncertainties and risks in aerospace components

NASA Technical Reports Server (NTRS)

Shah, A. R.; Shiao, M. C.; Nagpal, V. K.; Chamis, C. C.

1992-01-01

This paper summarizes a methodology developed at NASA Lewis Research Center which computationally simulates the structural, material, and load uncertainties associated with Space Shuttle Main Engine (SSME) components. The methodology was applied to evaluate the scatter in static, buckling, dynamic, fatigue, and damage behavior of the SSME turbo pump blade. Also calculated are the probability densities of typical critical blade responses, such as effective stress, natural frequency, damage initiation, most probable damage path, etc. Risk assessments were performed for different failure modes, and the effect of material degradation on the fatigue and damage behaviors of a blade were calculated using a multi-factor interaction equation. Failure probabilities for different fatigue cycles were computed and the uncertainties associated with damage initiation and damage propagation due to different load cycle were quantified. Evaluations on the effects of mistuned blades on a rotor were made; uncertainties in the excitation frequency were found to significantly amplify the blade responses of a mistuned rotor. The effects of the number of blades on a rotor were studied. The autocorrelation function of displacements and the probability density function of the first passage time for deterministic and random barriers for structures subjected to random processes also were computed. A brief discussion was included on the future direction of probabilistic structural analysis.

Vibration analysis of rotor blades with pendulum absorbers

NASA Technical Reports Server (NTRS)

Murthy, V. R.; Hammond, C. E.

1979-01-01

A comprehensive vibration analysis of rotor blades with spherical pendulum absorbers is presented. Linearized equations of motion for small oscillations about the steady-state deflection of a spherical pendulum on elastic rotor blades undergoing coupled flapwise bending, chordwise bending, and torsional vibrations are obtained. A transmission matrix formulation is given to determine the natural vibrational characteristics of rotor blades with spherical or simple flapping pendulum absorbers. The natural frequencies and mode shapes of a hingeless rotor blade with a spherical pendulum are computed.

Noninterference Systems Developed for Measuring and Monitoring Rotor Blade Vibrations

NASA Technical Reports Server (NTRS)

Kurkov, Anatole P.

2003-01-01

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

Evaluation of Rotor Structural and Aerodynamic Loads using Measured Blade Properties

NASA Technical Reports Server (NTRS)

Jung, Sung N.; You, Young-Hyun; Lau, Benton H.; Johnson, Wayne; Lim, Joon W.

2012-01-01

The structural properties of Higher harmonic Aeroacoustic Rotor Test (HART I) blades have been measured using the original set of blades tested in the wind tunnel in 1994. A comprehensive rotor dynamics analysis is performed to address the effect of the measured blade properties on airloads, blade motions, and structural loads of the rotor. The measurements include bending and torsion stiffness, geometric offsets, and mass and inertia properties of the blade. The measured properties are correlated against the estimated values obtained initially by the manufacturer of the blades. The previously estimated blade properties showed consistently higher stiffnesses, up to 30% for the flap bending in the blade inboard root section. The measured offset between the center of gravity and the elastic axis is larger by about 5% chord length, as compared with the estimated value. The comprehensive rotor dynamics analysis was carried out using the measured blade property set for HART I rotor with and without HHC (Higher Harmonic Control) pitch inputs. A significant improvement on blade motions and structural loads is obtained with the measured blade properties.

Effects of Unsteadiness Due to Wake Passing on Rotor Blade Heat Transfer

NASA Technical Reports Server (NTRS)

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

2007-01-01

14. ABSTRACT In a gas turbine engine, the turbine rotor blades are buffeted by the wakes of the vanes located upstream. There is a transient effect from the passing of wakes on the blade heat transfer. This transient effect has been computed for a representative rotor by introducing a wake upstream via an unsteady inlet flow boundary condition, or "gust" condition. Two cases of turbulent flow and laminar flow with Reynolds numbers of 385,000 and 385 respectively were considered. For the turbulent flow case a quasi-steady calculation was also performed. The variation in the unsteady heat transfer coefficient was found to be as high as 120 percent of the mean. For the turbulent flow case a quasisteady calculation was also performed. The time mean of the unsteady heat transfer, the mean of the quasi-steady variations and the steady results agree reasonably well on all blade locations except for the turbulent results which differ near the leading edge. The quasi-steady heat transfer results do not agree with the instantaneous unsteady results, although the time-mean values are similar.

An experimental study of the sensitivity of helicopter rotor blade tracking to root pitch adjustment in hover

NASA Technical Reports Server (NTRS)

Wilkie, W. Keats; Langston, Chester W.; Mirick, Paul H.; Singleton, Jeffrey D.; Wilbur, Matthew L.; Yeager, William T., Jr.

1991-01-01

The sensitivity of blade tracking in hover to variations in root pitch was examined for two rotor configurations. Tests were conducted using a four bladed articulated rotor mounted on the NASA-Army aeroelastic rotor experimental system (ARES). Two rotor configurations were tested: one consisting of a blade set with flexible fiberglass spars and one with stiffer (by a factor of five in flapwise and torsional stiffnesses) aluminum spars. Both blade sets were identical in planform and airfoil distribution and were untwisted. The two configurations were ballasted to the same Lock number so that a direct comparison of the tracking sensitivity to a gross change in blade stiffness could be made. Experimental results show no large differences between the two sets of blades in the sensitivity of the blade tracking to root pitch adjustments. However, a measurable reduction in intrack coning of the fiberglass spar blades with respect to the aluminum blades is noted at higher rotor thrust conditions.

Methods and apparatus for rotor blade ice detection

DOEpatents

LeMieux, David Lawrence

2006-08-08

A method for detecting ice on a wind turbine having a rotor and one or more rotor blades each having blade roots includes monitoring meteorological conditions relating to icing conditions and monitoring one or more physical characteristics of the wind turbine in operation that vary in accordance with at least one of the mass of the one or more rotor blades or a mass imbalance between the rotor blades. The method also includes using the one or more monitored physical characteristics to determine whether a blade mass anomaly exists, determining whether the monitored meteorological conditions are consistent with blade icing; and signaling an icing-related blade mass anomaly when a blade mass anomaly is determined to exist and the monitored meteorological conditions are determined to be consistent with icing.

Helicopter external noise prediction and reduction

NASA Astrophysics Data System (ADS)

Lewy, Serge

Helicopter external noise is a major challenge for the manufacturers, both in the civil domain and in the military domain. The strongest acoustic sources are due to the main rotor. Two flight conditions are analyzed in detail because radiated sound is then very loud and very impulsive: (1) high-speed flight, with large thickness and shear terms on the advancing blade side; and (2) descent flight, with blade-vortex interaction for certain rates of descent. In both cases, computational results were obtained and tests on new blade designs have been conducted in wind tunnels. These studies prove that large noise reduction can be achieved. It is shown in conclusion, however, that the other acoustic sources (tail rotor, turboshaft engines) must not be neglected to define a quiet helicopter.

Parametric Study and Design of Tab Shape for Improving Aerodynamic Performance of Rotor Blade

NASA Astrophysics Data System (ADS)

Han, Jaeseong; Kwon, Oh Joon

2018-04-01

In the present study, the parametric study was performed to analyze the effect of the tab on the aerodynamic performance and characteristics of rotor blades. Also, the tab shape was designed to improve the aerodynamic performance of rotor blades. A computational fluid dynamics solver based on three-dimensional Reynolds averaged Navier-Stokes equation using an unstructured mesh was used for the parametric study and the tab design. For airfoils, the effect of length and angle of a tab was studied on the aerodynamic characteristics of airfoils. In addition, including those parameters, the effect of a span of a tab was studied for rotor blades in hovering flight. The results of the parametric study were analyzed in terms of change of the aerodynamic performance and characteristics to understand the effect of a tab. Considering the analysis, the design of tab shape was conducted to improve the aerodynamic performance of rotor blades. The simply attached tab to trailing edge of the rotor blades increases the thrust of the rotor blades without significant changing of aerodynamic characteristics of the rotor blades in hovering and forward flight.

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.

Design and initial testing of a one-bladed 30-meter-diameter rotor on the NASA/DOE mod-O wind turbine

NASA Technical Reports Server (NTRS)

Corrigan, R. D.; Ensworth, C. B. F.

1986-01-01

The concept of a one-bladed horizontal-axis wind turbine has been of interest to wind turbine designers for many years. Many designs and economic analyses of one-bladed wind turbines have been undertaken by both United States and European wind energy groups. The analyses indicate significant economic advantages but at the same time, significant dynamic response concerns. In an effort to develop a broad data base on wind turbine design and operations, the NASA Wind Energy Project Office has tested a one-bladed rotor at the NASA/DOE Mod-O Wind Turbine Facility. This is the only known test on an intermediate-sized one-bladed rotor in the United States. The 15.2-meter-radius rotor consists of a tip-controlled blade and a counterweight assembly. A rigorous test series was conducted in the Fall of 1985 to collect data on rotor performance, drive train/generator dynamics, structural dynamics, and structural loads. This report includes background information on one-bladed rotor concepts, and Mod-O one-bladed rotor test configuration, supporting design analysis, the Mod-O one-blade rotor test plan, and preliminary test results.

Smart helicopter rotor with active blade tips

NASA Astrophysics Data System (ADS)

Bernhard, Andreas Paul Friedrich

2000-10-01

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

Aerodynamic optimization of wind turbine rotor using CFD/AD method

NASA Astrophysics Data System (ADS)

Cao, Jiufa; Zhu, Weijun; Wang, Tongguang; Ke, Shitang

2018-05-01

The current work describes a novel technique for wind turbine rotor optimization. The aerodynamic design and optimization of wind turbine rotor can be achieved with different methods, such as the semi-empirical engineering methods and more accurate computational fluid dynamic (CFD) method. The CFD method often provides more detailed aerodynamics features during the design process. However, high computational cost limits the application, especially for rotor optimization purpose. In this paper, a CFD-based actuator disc (AD) model is used to represent turbulent flow over a wind turbine rotor. The rotor is modeled as a permeable disc of equivalent area where the forces from the blades are distributed on the circular disc. The AD model is coupled with a Reynolds Averaged Navier-Stokes (RANS) solver such that the thrust and power are simulated. The design variables are the shape parameters comprising the chord, the twist and the relative thickness of the wind turbine rotor blade. The comparative aerodynamic performance is analyzed between the original and optimized reference wind turbine rotor. The results showed that the optimization framework can be effectively and accurately utilized in enhancing the aerodynamic performance of the wind turbine rotor.

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

NASA Technical Reports Server (NTRS)

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

2013-01-01

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

Life and Utilization Criteria Identification in Design (LUCID). Volume 1

DTIC Science & Technology

1981-10-01

stator, seal /spacer, etc. weights are added to these rotor weights in estimating module weights. Weights of other engine modules (combustor, augmentor...of turbine airfoil/platform cooling air and disk cooling/ seal leakage air), number of vanes and blades for the single stage high-pressure turbine, and...subroutines include hubs, shafts, seals and spacers in estimating rotor weights. Module weight is estimated by adding case and stator weights to the rotor

The effect of helicopter main rotor blade phasing and spacing on performance, blade loads, and acoustics

NASA Technical Reports Server (NTRS)

Gangwani, S. T.

1976-01-01

The performance, blade loads, and acoustic characteristics of a variable geometry rotor (VGR) system in forward flight and in a pullup maneuver were determined by the use of existing analytical programs. The investigation considered the independent effects of vertical separation of two three-bladed rotor systems as well as the effects of azimuthal spacing between the blades of the two rotors. The computations were done to determine the effects of these parameters on the performance, blade loads, and acoustic characteristics at two advance ratios in steady-state level flight and for two different g pullups at one advance ratio. To evaluate the potential benefits of the VGR concept in forward flight and pullup maneuvers, the results were compared as to performance, oscillatory blade loadings, vibratory forces transmitted to the fixed fuselage, and the rotor noise characteristics of the various VGR configurations with those of the conventional six-bladed rotor system.

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.

Aeroelastic considerations for torsionally soft rotors

NASA Technical Reports Server (NTRS)

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

1986-01-01

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

14 CFR 33.83 - Vibration test.

Code of Federal Regulations, 2010 CFR

2010-01-01

... 14 Aeronautics and Space 1 2010-01-01 2010-01-01 false Vibration test. 33.83 Section 33.83... STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.83 Vibration test. (a) Each engine... experience, analysis, and component test and shall address, as a minimum, blades, vanes, rotor discs, spacers...

14 CFR 33.83 - Vibration test.

Code of Federal Regulations, 2014 CFR

2014-01-01

... 14 Aeronautics and Space 1 2014-01-01 2014-01-01 false Vibration test. 33.83 Section 33.83... STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.83 Vibration test. (a) Each engine... experience, analysis, and component test and shall address, as a minimum, blades, vanes, rotor discs, spacers...

14 CFR 33.83 - Vibration test.

Code of Federal Regulations, 2013 CFR

2013-01-01

... 14 Aeronautics and Space 1 2013-01-01 2013-01-01 false Vibration test. 33.83 Section 33.83... STANDARDS: AIRCRAFT ENGINES Block Tests; Turbine Aircraft Engines § 33.83 Vibration test. (a) Each engine... experience, analysis, and component test and shall address, as a minimum, blades, vanes, rotor discs, spacers...

Minimum weight design of helicopter rotor blades with frequency constraints

NASA Technical Reports Server (NTRS)

Chattopadhyay, Aditi; Walsh, Joanne L.

1989-01-01

The minimum weight design of helicopter rotor blades subject to constraints on fundamental coupled flap-lag natural frequencies has been studied in this paper. A constraint has also been imposed on the minimum value of the blade autorotational inertia to ensure that the blade has sufficient inertia to autorotate in case of an engine failure. The program CAMRAD has been used for the blade modal analysis and the program CONMIN has been used for the optimization. In addition, a linear approximation analysis involving Taylor series expansion has been used to reduce the analysis effort. The procedure contains a sensitivity analysis which consists of analytical derivatives of the objective function and the autorotational inertia constraint and central finite difference derivatives of the frequency constraints. Optimum designs have been obtained for blades in vacuum with both rectangular and tapered box beam structures. Design variables include taper ratio, nonstructural segment weights and box beam dimensions. The paper shows that even when starting with an acceptable baseline design, a significant amount of weight reduction is possible while satisfying all the constraints for blades with rectangular and tapered box beams.

Aeronautical Engineering. A Continuing Bibliography with Indexes

DTIC Science & Technology

1987-09-01

engines 482 01 AERONAUTICS (GENERAL) i-10 aircraft equipped with turbine engine ...rate adaptive control with applications to lateral Statistics on aircraft gas turbine engine rotor failures Unified model for the calculation of blade ...PUMPS p 527 A87-35669 to test data for a composite prop-tan model Gas turbine combustor and engine augmentor tube GENERAL AVIATION AIRCRAFT

Active Blade Vibration Control Being Developed and Tested

NASA Technical Reports Server (NTRS)

Johnson, Dexter

2003-01-01

Gas turbine engines are currently being designed to have increased performance, lower weight and manufacturing costs, and higher reliability. Consequently, turbomachinery components, such as turbine and compressor blades, have designs that are susceptible to new vibration problems and eventual in-service failure due to high-cycle fatigue. To address this problem, researchers at the NASA Glenn Research Center are developing and testing innovative active blade vibration control concepts. Preliminary results of using an active blade vibration control system, involving a rotor supported by an active magnetic bearing in Glenn's Dynamic Spin Rig, indicate promising results (see the photograph). Active blade vibration control was achieved using feedback of blade strain gauge signals within the magnetic bearing control loop. The vibration amplitude was reduced substantially (see the graphs). Also, vibration amplitude amplification was demonstrated; this could be used to enhance structural mode identification, if desired. These results were for a nonrotating two-bladed disk. Tests for rotating blades are planned. Current and future active blade vibration control research is planned to use a fully magnetically suspended rotor and smart materials. For the fully magnetically suspended rotor work, three magnetic bearings (two radial and one axial) will be used as actuators instead of one magnetic bearing. This will allow additional degrees of freedom to be used for control. For the smart materials work, control effectors located on and off the blade will be considered. Piezoelectric materials will be considered for on-the-blade actuation, and actuator placement on a stator vane, or other nearby structure, will be investigated for off-the-blade actuation. Initial work will focus on determining the feasibility of these methods by performing basic analysis and simple experiments involving feedback control.

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

Code of Federal Regulations, 2010 CFR

2010-01-01

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

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

Code of Federal Regulations, 2010 CFR

2010-01-01

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

The start-up of a gas turbine engine using compressed air tangentially fed onto the blades of the basic turbine

NASA Technical Reports Server (NTRS)

Slobodyanyuk, L. K.; Dayneko, V. I.

1983-01-01

The use of compressed air was suggested to increase the reliability and motor lifetime of a gas turbine engine. Experiments were carried out and the results are shown in the form of the variation in circumferential force as a function of the entry angle of the working jet onto the turbine blade. The described start-up method is recommended for use with massive rotors.

Corner flow control in high through-flow axial commercial fan/booster using blade 3-D optimization

NASA Astrophysics Data System (ADS)

Zhu, Fang; Jin, Donghai; Gui, Xingmin

2012-02-01

This study is aimed at using blade 3-D optimization to control corner flows in the high through-flow fan/booster of a high bypass ratio commercial turbofan engine. Two kinds of blade 3-D optimization, end-bending and bow, are focused on. On account of the respective operation mode and environment, the approach to 3-D aerodynamic modeling of rotor blades is different from stator vanes. Based on the understanding of the mechanism of the corner flow and the consideration of intensity problem for rotors, this paper uses a variety of blade 3-D optimization approaches, such as loading distribution optimization, perturbation of departure angles and stacking-axis manipulation, which are suitable for rotors and stators respectively. The obtained 3-D blades and vanes can improve the corner flow features by end-bending and bow effects. The results of this study show that flows in corners of the fan/booster, such as the fan hub region, the tip and hub of the vanes of the booster, are very complex and dominated by 3-D effects. The secondary flows there are found to have a strong detrimental effect on the compressor performance. The effects of both end-bending and bow can improve the flow separation in corners, but the specific ways they work and application scope are somewhat different. Redesigning the blades via blade 3-D optimization to control the corner flow has effectively reduced the loss generation and improved the stall margin by a large amount.

Cost/benefit analysis of advanced material technologies for small aircraft turbine engines

NASA Technical Reports Server (NTRS)

Comey, D. H.

1977-01-01

Cost/benefit studies were conducted on ten advanced material technologies applicable to small aircraft gas turbine engines to be produced in the 1985 time frame. The cost/benefit studies were applied to a two engine, business-type jet aircraft in the 6800- to 9100-Kg (15,000- to 20,000-lb) gross weight class. The new material technologies are intended to provide improvements in the areas of high-pressure turbine rotor components, high-pressure turbine rotor components, high-pressure turbine stator airfoils, and static structural components. The cost/benefit of each technology is presented in terms of relative value, which is defined as a change in life cycle cost times probability of success divided by development cost. Technologies showing the most promising cost/benefits based on relative value are uncooled single crystal MAR-M 247 turbine blades, cooled DS MAR-M 247 turbine blades, and cooled ODS 'M'CrAl laminate turbine stator vanes.

Advanced Gas Turbine (AGT) powertrain system development for automotive applications

NASA Technical Reports Server (NTRS)

1984-01-01

Rotor dynamic instability investigations were conducted. Forward ball bearing hydraulic mount configurations were tested with little effect. Trial assembly of S/N 002 ceramic engine was initiated. Impeller design activities were completed on the straight line element (SLE) blade definition to address near-net-shape powder metal die forging. Performance characteristics of the Baseline Test 2A impeller were closely preserved. The modified blading design has been released for tooling procurement. Developmental testing of the diffusion flame combustor (DFC) for initial use in the S/N 002 2100 F ceramic structures engine was completed. A natural gas slave preheater was designed and fabricated. Preliminary regenerator static seal rig testing showed a significant reduction in leakage and sensitivity to stack height. Ceramic screening tests were completed and two complete sets of ceramic static structures were qualified for engine testing. Efforts on rotor dynamics development to resolve subsynchronous motion were continued.

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.

Rotor Hover Performance and Flowfield Measurements with Untwisted and Highly-Twisted Blades

NASA Technical Reports Server (NTRS)

Ramasamy, Manikandan; Gold, Nili P.; Bhagwat, Mahendra J.

2010-01-01

The flowfield and performance characteristics of highly-twisted blades were analyzed at various thrust conditions to improve the fundamental understanding relating the wake effects on rotor performance. Similar measurements made using untwisted blades served as the baseline case. Twisted blades are known to give better hover performance than untwisted blades at high thrust coefficients typical of those found in full-scale rotors. However, the present experiments were conducted at sufficiently low thrust (beginning from zero thrust), where the untwisted blades showed identical, if not better, performance when compared with the highly-twisted blades. The flowfield measurements showed some key wake differences between the two rotors, as well. These observations when combined with simple blade element momentum theory (also called annular disk momentum theory) helped further the understanding of rotor performance characteristics.

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

DOEpatents

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

2006-10-10

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

Turbine and method for reducing shock losses in a turbine

DOEpatents

Ristau, Neil

2015-09-01

A turbine includes a rotor and a casing that circumferentially surrounds at least a portion of the rotor. The rotor and the casing at least partially define a gas path through the turbine. A last stage of rotating blades is circumferentially arranged around the rotor and includes a downstream swept portion radially outward from the rotor. A method for reducing shock losses in a turbine includes removing a last stage of rotating blades circumferentially arranged around a rotor and replacing the last stage of rotating blades with rotating blades having a downstream swept portion radially outward from the rotor.

BLADED IMPELLER FOR TURBOBLOWERS

DOEpatents

Baumann, K.

1949-10-01

A means is given of holding open-sided impeller blades in a turbo-rotor. Two half blades, with dovetail roots of sufficient weight to contain the center of gravity, are fitted into slots cut in the rotor so as to form the desired angle between the blade faces. The adjoining edges of the half blades are welded to form one solid blade that is securely locked an the rotor. This design permits the manufacture of a V shaped impeller blade without the need of machining the entire V shaped contour from a single blank, and furthermore provides excellent locking characteristics for attachment to the rotor.

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.

Thermal stresses investigation of a gas turbine blade

NASA Astrophysics Data System (ADS)

Gowreesh, S.; Pravin, V. K.; Rajagopal, K.; Veena, P. H.

2012-06-01

The analysis of structural and thermal stress values that are produced while the turbine is operating are the key factors of study while designing the next generation gas turbines. The present study examines structural, thermal, modal analysis of the first stage rotor blade of a two stage gas turbine. The design features of the turbine segment of the gas turbine have been taken from the preliminary design of a power turbine for maximization of an existing turbojet engine with optimized dump gap of the combustion chamber, since the allowable temperature on the turbine blade dependents on the hot gas temperatures from the combustion chamber. In the present paper simplified 3-D Finite Element models are developed with governing boundary conditions and solved using the commercial FEA software ANSYS. As the temperature has a significant effect on the overall stress on the rotor blades, a detail study on mechanical and thermal stresses are estimated and evaluated with the experimental values.

Exo-Skeletal Engine: Novel Engine Concept

NASA Technical Reports Server (NTRS)

Chamis, Cristos C.; Blankson, Isaiah M.

2004-01-01

The exo-skeletal engine concept represents a new radical engine technology with the potential to substantially revolutionize engine design. It is an all-composite drum-rotor engine in which conventionally heavy shafts and discs are eliminated and are replaced by rotating casings that support the blades in spanwise compression. Thus the rotating blades are in compression rather than tension. The resulting open channel at the engine centerline has immense potential for jet noise reduction and can also accommodate an inner combined-cycle thruster such as a ramjet. The exo-skeletal engine is described in some detail with respect to geometry, components, and potential benefits. Initial evaluations and results for drum rotors, bearings, and weights are summarized. Component configuration, assembly plan, and potential fabrication processes are also identified. A finite element model of the assembled engine and its major components is described. Preliminary results obtained thus far show at least a 30-percent reduction of engine weight and about a 10-dB noise reduction, compared with a baseline conventional high-bypass-ratio engine. Potential benefits in all aspects of this engine technology are identified and tabulated. Quantitative assessments of potential benefits are in progress.

Acoustic testing of a 1.5 pressure ratio low tip speed fan with a serrated rotor (QEP fan B scale model). [reduction of engine noise

NASA Technical Reports Server (NTRS)

Kazin, S. B.; Paas, J. E.; Minzner, W. R.

1973-01-01

A scale model of the bypass flow region of a 1.5 pressure ratio, single stage, low tip speed fan was tested with a serrated rotor leading edge to determine its effects on noise generation. The serrated rotor was produced by cutting teeth into the leading edge of the nominal rotor blades. The effects of speed and exhaust nozzle area on the scale models noise characteristics were investigated with both the nominal rotor and serrated rotor. Acoustic results indicate the serrations reduced front quadrant PNL's at takeoff power. In particular, the 200 foot (61.0 m) sideline noise was reduced from 3 to 4 PNdb at 40 deg for nominal and large nozzle operation. However, the rear quadrant maximum sideline PNL's were increased 1.5 to 3 PNdb at approach thust and up to 2 PNdb at takeoff thust with these serrated rotor blades. The configuration with the serrated rotor produced the lowest maximum 200 foot (61.0 m) sideline PNL for any given thust when the large nozzle (116% of design area) was employed.

Investigation of Rotor Performance and Loads of a UH-60A Individual Blade Control System

DTIC Science & Technology

2010-05-01

the pitch link for each rotor blade with an actuator so that the blade root pitch angles could be changed independently. This design was previously...with the ultimate goal of providing the technology for timely and cost-effective design and development of new rotors. Analytical studies on IBC [8...rotor with coincident flap and lag articulation provided at the blade root by elastomeric bearings. This bearing, through the rotor spindle , also

Effects of Rotor Blade Scaling on High-Pressure Turbine Unsteady Loading

NASA Astrophysics Data System (ADS)

Lastiwka, Derek; Chang, Dongil; Tavoularis, Stavros

2013-03-01

The present work is a study of the effects of rotor blade scaling of a single-stage high pressure turbine on the time-averaged turbine performance and on parameters that influence vibratory stresses on the rotor blades and stator vanes. Three configurations have been considered: a reference case with 36 rotor blades and 24 stator vanes, a case with blades upscaled by 12.5%, and a case with blades downscaled by 10%. The present results demonstrate that blade scaling effects were essentially negligible on the time-averaged turbine performance, but measurable on the unsteady surface pressure fluctuations, which were intensified as blade size was increased. In contrast, blade torque fluctuations increased significantly as blade size decreased. Blade scaling effects were also measurable on the vanes.

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).

A prescribed wake rotor inflow and flow field prediction analysis, user's manual and technical approach

NASA Technical Reports Server (NTRS)

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

1982-01-01

A user's manual is provided which includes the technical approach for the Prescribed Wake Rotor Inflow and Flow Field Prediction Analysis. The analysis is used to provide the rotor wake induced velocities at the rotor blades for use in blade airloads and response analyses and to provide induced velocities at arbitrary field points such as at a tail surface. This analysis calculates the distribution of rotor wake induced velocities based on a prescribed wake model. Section operating conditions are prescribed from blade motion and controls determined by a separate blade response analysis. The analysis represents each blade by a segmented lifting line, and the rotor wake by discrete segmented trailing vortex filaments. Blade loading and circulation distributions are calculated based on blade element strip theory including the local induced velocity predicted by the numerical integration of the Biot-Savart Law applied to the vortex wake model.

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.

Boundary Layer Transition Detection on a Rotor Blade Using Rotating Mirror Thermography

NASA Technical Reports Server (NTRS)

Heineck, James T.; Schuelein, Erich; Raffel, Markus

2014-01-01

Laminar-to-turbulent transition on a rotor blade in hover has been imaged using an area-scan infrared camera. A new method for tracking a blade using a rotating mirror was employed. The mirror axis of rotation roughly corresponded to the rotor axis of rotation and the mirror rotational frequency is 1/2 that of the rotor. This permitted the use of cameras whose integration time was too long to prevent image blur due to the motion of the blade. This article will show the use of this method for a rotor blade at different collective pitch angles.

Blade attachment assembly

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

Garcia-Crespo, Andres Jose; Delvaux, John McConnell; Miller, Diane Patricia

An assembly and method for affixing a turbomachine rotor blade to a rotor wheel are disclosed. In an embodiment, an adaptor member is provided disposed between the blade and the rotor wheel, the adaptor member including an adaptor attachment slot that is complementary to the blade attachment member, and an adaptor attachment member that is complementary to the rotor wheel attachment slot. A coverplate is provided, having a coverplate attachment member that is complementary to the rotor wheel attachment slot, and a hook for engaging the adaptor member. When assembled, the coverplate member matingly engages with the adaptor member, andmore » retains the blade in the adaptor member, and the assembly in the rotor wheel.« less

Vibration analysis of rotor blades with an attached concentrated mass

NASA Technical Reports Server (NTRS)

Murthy, V. R.; Barna, P. S.

1977-01-01

The effect of an attached concentrated mass on the dynamics of helicopter rotor blades is determined. The point transmission matrix method was used to define, through three completely automated computer programs, the natural vibrational characteristics (natural frequencies and mode shapes) of rotor blades. The problems of coupled flapwise bending, chordwise bending, and torsional vibration of a twisted nonuniform blade and its special subcase pure torsional vibration are discussed. The orthogonality relations that exist between the natural modes of rotor blades with an attached concentrated mass are derived. The effect of pitch, rotation, and point mass parameters on the collective, cyclic, scissor, and pure torsional modes of a seesaw rotor blade is determined.

Daniel Laird | NREL

Science.gov Websites

the high-fidelity modeling, wind plant controls, and rotor dynamics focus areas. Prior to joining NREL , composite materials, and blade reliability. Education M.S. and Ph.D. in Mechanical Engineering, University

Stainless-Steel-Foam Structures Evaluated for Fan and Rotor Blades

NASA Technical Reports Server (NTRS)

Lerch, Bradley A.; Raj, Sai V.; Ghosn, Louis J.; Hebsur, Mohan G.; Cosgriff, Laura M.; Min, James B.; Holland, Frederic A., Jr.

2005-01-01

The goal of this project is to use a sandwich structure design, consisting of two stainlesssteel face sheets and a stainless-steel-foam core, to fabricate engine fan and propeller blades. Current fan blades are constructed either of polymer matrix composites (PMCs) or hollow titanium alloys. The PMC blades are expensive and have poor impact resistance on their leading edges, thereby requiring a metallic leading edge to satisfy the Federal Aviation Administration s impact requirements relating to bird strikes. Hollow titanium blades cost more to fabricate because of the intrinsically difficult fabrication issues associated with titanium alloys. However, both these current concepts produce acceptable lightweight fan blades.

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

NASA Technical Reports Server (NTRS)

Wheatley, John B; Hood, Manley J

1935-01-01

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

Aeroelastic considerations for torsionally soft rotors

NASA Technical Reports Server (NTRS)

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

1985-01-01

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

Test Operations Procedure (TOP) 01-2-603 Rotorcraft Laboratory Vibration Test Schedules

DTIC Science & Technology

2017-06-12

for all rotary wing aircraft platforms. Tonal amplitudes are tabular based solely on engine revolutions per minute (RPM) and blade count. (4...Power Spectral Density (PSD) format with superimposed sinusoidal components that are associated with the rotor speeds and blade count of each...harmonics are not limited to the 3rd harmonic of the blade passage as in MIL-STD- TOP 01-2-603 12 June 2017 5 810. In addition, attempts were

Performance Limiting Flow Processes in High-State Loading High-Mach Number Compressors

DTIC Science & Technology

2008-03-13

the Doctoral Thesis Committee of the doctoral student. 3 3.0 Technical Background A strong incentive exists to reduce airfoil count in aircraft engine ...Advanced Turbine Engine ). A basic constraint on blade reduction is seen from the Euler turbine equation, which shows that, although a design can be carried...on the vane to rotor blade ratio of 8:11). Within the MSU Turbo code, specifying a small number of time steps requires more iteration at each time

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

NASA Technical Reports Server (NTRS)

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

1972-01-01

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

The investigation of a variable camber blade lift control for helicopter rotor systems

NASA Technical Reports Server (NTRS)

Awani, A. O.

1982-01-01

A new rotor configuration called the variable camber rotor was investigated numerically for its potential to reduce helicopter control loads and improve hover performance. This rotor differs from a conventional rotor in that it incorporates a deflectable 50% chord trailing edge flap to control rotor lift, and a non-feathering (fixed) forward portion. Lift control is achieved by linking the blade flap to a conventional swashplate mechanism; therefore, it is pilot action to the flap deflection that controls rotor lift and tip path plane tilt. This report presents the aerodynamic characteristics of the flapped and unflapped airfoils, evaluations of aerodynamics techniques to minimize flap hinge moment, comparative hover rotor performance and the physical concepts of the blade motion and rotor control. All the results presented herein are based on numerical analyses. The assessment of payoff for the total configuration in comparison with a conventional blade, having the same physical characteristics as an H-34 helicopter rotor blade was examined for hover only.

Development of an active twist rotor blade with distributed actuation and orthotropic material

NASA Astrophysics Data System (ADS)

Wierach, Peter; Riemenschneider, Johannes; Keye, Stefan

2005-05-01

Individual blade control (IBC) as well as higher harmonic control (HHC) for helicopter rotors promises to be a method to increase flight performance and to reduce vibration and noise. For those controls, an additional twist actuation of the rotor blade is needed. The developed concept comprises the implementation of distributed piezoelectric actuation into the rotor blade skin. In order to maximize the twist within given constraints, as torsional rigidity and given actuator design, the concept takes advantage of an orthotropic rotor blade skin. That way, a combination of shear actuation with orthotropic coupling generates more twist than each one of these effects alone. Previous approaches with distributed actuation used actuators operating in +/-45° direction with quasi-isotropic composites. A FE-Model of the blade was developed and validated using a simplified demonstrator. The objective of this study was to identify the effects of various geometric and material parameters to optimize the active twist performance of the blades. The whole development was embedded in an iterative process followed by an objective assessment. For this purpose a detailed structural model on the basis of the BO105 model rotor blade was developed, to predict the performance with respect to rotor dynamics, stability, aerodynamics and acoustics. Rotor dynamic simulations provided an initial overview of the active twist rotor performance. In comparison to the BO105 baseline rotor a noise reduction of 3 dB was predicted for an active twist of 0.8° at the blade tip. Additionally, a power reduction of 2.3% at 87m/s based on a 2.5 to BO105 was computed. A demonstrator blade with a rotor radius of 2m has been designed and manufactured. This blade will be tested to prove, that the calculated maximum twist can also be achieved under centrifugal loads.

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

NASA Technical Reports Server (NTRS)

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

1971-01-01

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

Minimum weight design of rectangular and tapered helicopter rotor blades with frequency constraints

NASA Technical Reports Server (NTRS)

Chattopadhyay, Aditi; Walsh, Joanne L.

1988-01-01

The minimum weight design of a helicopter rotor blade subject to constraints on coupled flap-lag natural frequencies has been studied. A constraint has also been imposed on the minimum value of the autorotational inertia of the blade in order to ensure that it has sufficient inertia to autorotate in the case of engine failure. The program CAMRAD is used for the blade modal analysis and CONMIN is used for the optimization. In addition, a linear approximation analysis involving Taylor series expansion has been used to reduce the analysis effort. The procedure contains a sensitivity analysis which consists of analytical derivatives of the objective function and the autorotational inertia constraint and central finite difference derivatives of the frequency constraints. Optimum designs have been obtained for both rectangular and tapered blades. Design variables include taper ratio, segment weights, and box beam dimensions. It is shown that even when starting with an acceptable baseline design, a significant amount of weight reduction is possible while satisfying all the constraints for both rectangular and tapered blades.

Minimum weight design of rectangular and tapered helicopter rotor blades with frequency constraints

NASA Technical Reports Server (NTRS)

Chattopadhyay, Aditi; Walsh, Joanne L.

1988-01-01

The minimum weight design of a helicopter rotor blade subject to constraints on coupled flap-lag natural frequencies has been studied. A constraint has also been imposed on the minimum value of the autorotational inertia of the blade in order to ensure that it has sufficient inertia to aurorotate in the case of engine failure. The program CAMRAD is used for the blade modal analysis and CONMIN is used for the optimization. In addition, a linear approximation analysis involving Taylor series expansion has been used to reduce the analysis effort. The procedure contains a sensitivity analysis which consists of analytical derivatives of the objective function and the autorotational inertia constraint and central finite difference derivatives of the frequency constraints. Optimum designs have been obtained for both rectangular and tapered blades. Design variables include taper ratio, segment weights, and box beam dimensions. It is shown that even when starting with an acceptable baseline design, a significant amount of weight reduction is possible while satisfying all the constraints for both rectangular and tapered blades.

Structural analysis of wind turbine rotors for NSF-NASA Mod-0 wind power system

NASA Technical Reports Server (NTRS)

Spera, D. A.

1976-01-01

Preliminary estimates are presented of vibratory loads and stresses in hingeless and teetering rotors for the proposed NSF-NASA Mod-0 wind power system. Preliminary blade design utilizes a tapered tubular aluminum spar which supports nonstructural aluminum ribs and skin and is joined to the rotor hub by a steel shank tube. Stresses in the shank of the blade are calculated for static, rated, and overload operating conditions. Blade vibrations were limited to the fundamental flapping modes, which were elastic cantilever bending for hingeless rotor blades and rigid-body rotation for teetering rotor blades. The MOSTAB-C computer code was used to calculate aerodynamic and mechanical loads. The teetering rotor has substantial advantages over the hingeless rotor with respect to shank stresses, fatigue life, and tower loading. The hingeless rotor analyzed does not appear to be structurally stable during overloads.

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

NASA Technical Reports Server (NTRS)

Owczarek, J. A.

1983-01-01

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

Design of helicopter rotor blades for optimum dynamic characteristics

NASA Technical Reports Server (NTRS)

Peters, D. A.; Ko, T.; Korn, A.; Rossow, M. P.

1984-01-01

The optimal design of helicopter rotor blades is addressed. The forced response of an initial (i.e., non-optimized) blade to those of a final (optimized) blade are compared. Response of starting design and optimal designs for varying forcing frequencies, blade response to harmonics of rotor speed, and derivation of mass and stiffness matrices or functions of natural frequencies are discussed.

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.

Gas turbine engine

DOEpatents

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

2016-03-08

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

Stability Analysis of the Slowed-Rotor Compound Helicopter Configuration

NASA Technical Reports Server (NTRS)

Johnson, Wayne; Floros, Matthew W.

2004-01-01

The stability and control of rotors at high advance ratio are considered. Teetering, articulated, gimbaled, and rigid hub types are considered for a compound helicopter (rotor and fixed wing). Stability predictions obtained using an analytical rigid flapping blade analysis, a rigid blade CAMRAD II model, and an elastic blade CAMRAD II model are compared. For the flapping blade analysis, the teetering rotor is the most stable, 5howing no instabilities up to an advance ratio of 3 and a Lock number of 18. With an elastic blade model, the teetering rotor is unstable at an advance ratio of 1.5. Analysis of the trim controls and blade flapping shows that for small positive collective pitch, trim can be maintained without excessive control input or flapping angles.

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

NASA Technical Reports Server (NTRS)

Schum, Harold J.; Whitney, Warren J.

1949-01-01

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

Predesign study for a modern 4-bladed rotor for RSRA

NASA Technical Reports Server (NTRS)

Davis, S. J.

1981-01-01

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

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 of an Advanced Wood Composite Rotor and Development of Wood Composite Blade Technology

NASA Technical Reports Server (NTRS)

Stroebel, Thomas; Dechow, Curtis; Zuteck, Michael

1984-01-01

In support of a program to advance wood composite wind turbine blade technology, a design was completed for a prototype, 90-foot diameter, two-bladed, one-piece rotor, with all wood/epoxy composite structure. The rotor was sized for compatibility with a generator having a maximum power rating of 4000 kilowatts. Innovative features of the rotor include: a teetering hub to minimize the effects of gust loads, untwisted blades to promote rotor power control through stall, joining of blades to the hub structure via an adhesive bonded structural joint, and a blade structural design which was simplified relative to earlier efforts. The prototype rotor was designed to allow flexibility for configuring the rotor upwind or downwind of the tower, for evaluating various types of teeter dampers and/or elastomeric stops, and with variable delta-three angle settings of the teeter shaft axis. The prototype rotor was also designed with provisions for installing pressure tap and angle of attack instrumentation in one blade. A production version rotor cost analysis was conducted. Included in the program were efforts directed at developing advanced load take-off stud designs for subsequent evaluation testing by NASA, development of aerodynamic tip brake concepts, exploratory testing of a wood/epoxy/graphite concept, and compression testing of wood/epoxy laminate, with scarf-jointed plies.

Tail Rotor Airfoils Stabilize Helicopters, Reduce Noise

NASA Technical Reports Server (NTRS)

2010-01-01

Founded by former Ames Research Center engineer Jim Van Horn, Van Horn Aviation of Tempe, Arizona, built upon a Langley Research Center airfoil design to create a high performance aftermarket tail rotor for the popular Bell 206 helicopter. The highly durable rotor has a lifetime twice that of the original equipment manufacturer blade, reduces noise by 40 percent, and displays enhanced performance at high altitudes. These improvements benefit helicopter performance for law enforcement, military training, wildfire and pipeline patrols, and emergency medical services.

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.

Ballistic and Cyclic Rig Testing of Braided Composite Fan Case Structures

NASA Technical Reports Server (NTRS)

Watson, William R.; Roberts, Gary D.; Pereira, J. Michael; Braley, Michael S.

2015-01-01

FAA fan blade-out certification testing on turbofan engines occurs very late in an engine's development program and is very costly. It is of utmost importance to approach the FAA Certification engine test with a high degree of confidence that the containment structure will not only contain the high-energy debris, but that it will also withstand the cyclic loads that occur with engine spooldown and continued rotation as the non-running engine maintains a low rotor RPM due to forced airflow as the engine-out aircraft returns to an airport. Accurate rig testing is needed for predicting and understanding material behavior of the fan case structure during all phases of this fan blade-out event.

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.

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

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

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.

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

Blade Displacement Measurements of the Full-Scale UH-60A Airloads Rotor

NASA Technical Reports Server (NTRS)

Barrows, Danny A.; Burner, Alpheus W.; Abrego, Anita I.; Olson, Lawrence E.

2011-01-01

Blade displacement measurements were acquired during a wind tunnel test of the full-scale UH-60A Airloads rotor. The test was conducted in the 40- by 80-Foot Wind Tunnel of the National Full-Scale Aerodynamics Complex at NASA Ames Research Center. Multi-camera photogrammetry was used to measure the blade displacements of the four-bladed rotor. These measurements encompass a range of test conditions that include advance ratios from 0.15 to unique slowed-rotor simulations as high as 1.0, thrust coefficient to rotor solidity ratios from 0.01 to 0.13, and rotor shaft angles from -10.0 to 8.0 degrees. The objective of these measurements is to provide a benchmark blade displacement database to be utilized in the development and validation of rotorcraft computational tools. The methodology, system development, measurement techniques, and preliminary sample blade displacement measurements are presented.

Aeroelastic Stability of Modern Bearingless Rotors: A Parametric Investigation

NASA Technical Reports Server (NTRS)

Nguyen, Khanh Q.

1994-01-01

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

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.

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.

Nonlinear static and dynamic analysis of beam structures using fully intrinsic equations

NASA Astrophysics Data System (ADS)

Sotoudeh, Zahra

2011-07-01

Beams are structural members with one dimension much larger than the other two. Examples of beams include propeller blades, helicopter rotor blades, and high aspect-ratio aircraft wings in aerospace engineering; shafts and wind turbine blades in mechanical engineering; towers, highways and bridges in civil engineering; and DNA modeling in biomedical engineering. Beam analysis includes two sets of equations: a generally linear two-dimensional problem over the cross-sectional plane and a nonlinear, global one-dimensional analysis. This research work deals with a relatively new set of equations for one-dimensional beam analysis, namely the so-called fully intrinsic equations. Fully intrinsic equations comprise a set of geometrically exact, nonlinear, first-order partial differential equations that is suitable for analyzing initially curved and twisted anisotropic beams. A fully intrinsic formulation is devoid of displacement and rotation variables, making it especially attractive because of the absence of singularities, infinite-degree nonlinearities, and other undesirable features associated with finite rotation variables. In spite of the advantages of these equations, using them with certain boundary conditions presents significant challenges. This research work will take a broad look at these challenges of modeling various boundary conditions when using the fully intrinsic equations. Hopefully it will clear the path for wider and easier use of the fully intrinsic equations in future research. This work also includes application of fully intrinsic equations in structural analysis of joined-wing aircraft, different rotor blade configuration and LCO analysis of HALE aircraft.

Preliminary Airworthiness Evaluation AH-1S Helicopter with OGEE Tip Shape Rotor Blades

DTIC Science & Technology

1980-05-01

ENGINEER PROJECT PILOT HENRY ARNAIZ PROJECT ENGINEER DTIC MAY 1980 ELECTEV SEP 2 I8 Approved for public release; distribution unlimited. A UNITED STATES...compressibility effects between flights. 7. Airspeed and altitude were obtained from a boom-mounted pitot -static probe. Corrections for position error

Summary of Full-Scale Blade Displacement Measurements of the UH- 60A Airloads Rotor

NASA Technical Reports Server (NTRS)

Abrego, Anita I.; Meyn, Larry; Burner, Alpheus W.; Barrows, Danny A.

2016-01-01

Blade displacement measurements using multi-camera photogrammetry techniques were acquired for a full-scale UH-60A rotor, tested in the National Full-Scale Aerodynamic Complex 40-Foot by 80-Foot Wind Tunnel. The measurements, acquired over the full rotor azimuth, encompass a range of test conditions that include advance ratios from 0.15 to 1.0, thrust coefficient to rotor solidity ratios from 0.01 to 0.13, and rotor shaft angles from -10.0 to 8.0 degrees. The objective was to measure the blade displacements and deformations of the four rotor blades and provide a benchmark blade displacement database to be utilized in the development and validation of rotorcraft prediction techniques. An overview of the blade displacement measurement methodology, system development, and data analysis techniques are presented. Sample results based on the final set of camera calibrations, data reduction procedures and estimated corrections that account for registration errors due to blade elasticity are shown. Differences in blade root pitch, flap and lag between the previously reported results and the current results are small. However, even small changes in estimated root flap and pitch can lead to significant differences in the blade elasticity values.

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

NASA Technical Reports Server (NTRS)

Brent, J. A.; Clemmons, D.

1972-01-01

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

Hover performance tests of baseline metal and Advanced Technology Blade (ATB) rotor systems for the XV-15 tilt rotor aircraft

NASA Technical Reports Server (NTRS)

Bartie, K.; Alexander, H.; Mcveigh, M.; Lamon, S.; Bishop, H.

1986-01-01

Rotor hover performance data were obtained for two full-scale rotor systems designed for the XV-15 Tilt Rotor Research Aircraft. One rotor employed the rectangular planform metal blades (rotor solidity = 0.089) which were used on the initial flight configuration of the XV-15. The second rotor configuration examined the nonlinear taper, composite-construction, Advanced Technology Blade (ATB), (rotor solidity = 0.10) designed to replace the metal blades on the XV-15. Variations of the baseline ATB tip and cuff shapes were also tested. A new six-component rotor force and moment balance designed to obtain highly accurate data over a broad range of thrust and torque conditions is described. The test data are presented in nondimensional coefficient form for the performance results, and in dimensional form for the steady and alternating loads. Some wake and acoustic data are also shown.

Effects of blade-to-blade dissimilarities on rotor-body lead-lag dynamics

NASA Technical Reports Server (NTRS)

Mcnulty, M. J.

1986-01-01

Small blade-to-blade property differences are investigated to determine their effects on the behavior of a simple rotor-body system. An analytical approach is used which emphasizes the significance of these effects from the experimental point of view. It is found that the primary effect of blade-to-blade dissimilarities is the appearance of additional peaks in the frequency spectrum which are separated from the convention response modes by multiples of the rotor speed. These additional responses are potential experimental problems because when they occur near a mode of interest they act as contaminant frequencies which can make damping measurements difficult. The effects of increased rotor-body coupling and a rotor shaft degree of freedom act to improve the situation by altering the frequency separation of the modes.

Stability Analysis of the Slowed-Rotor Compound Helicopter Configuration

NASA Technical Reports Server (NTRS)

Floros, Matthew W.; Johnson, Wayne

2007-01-01

The stability and control of rotors at high advance ratio are considered. Teetering, articulated, gimbaled, and rigid hub types are considered for a compound helicopter (rotor and fixed wing). Stability predictions obtained using an analytical rigid flapping blade analysis, a rigid blade CAMRAD II model, and an elastic blade CAMRAD II model are compared. For the flapping blade analysis, the teetering rotor is the most stable, showing no instabilities up to an advance ratio of 3 and a Lock number of 18. A notional elastic blade model of a teetering rotor is unstable at an advance ratio of 1.5, independent of pitch frequency. Analysis of the trim controls and blade flapping shows that for small positive collective pitch, trim can be maintained without excessive control input or flapping angles.

Noise reduction experience at Hughes Helicopter, Inc.

NASA Astrophysics Data System (ADS)

Janakiram, D. S.

1982-07-01

Noise reduction is mostly limited to light helicopters whose noise signature is dominated by their tail rotors. It is primarily hardware oriented. Well known noise reduction techniques such as reduction of rotor speeds with an accompanying increase in solidity to maintain performance, engine noise reduction with the use of exhaust mufflers, and acoustic blanketing of transmission and engine compartment are used. The concept of blade phasing as a means of reducing tail rotor noise is also used. Engine noise (exhaust noise), power train noise and airframe noise becomes important at low rotor tip speeds and means must be found to reduce these noise sources if further noise reductions are desired. The use of a special test rig aids in isolating the various noise sources and arriving at the penalties (performance or payload) involved in quieting them. Significant noise reduction are achieved for the light helicopter with minimum performance or weight penalties because of the dominance of a single noise source (the tail rotor).

Noise reduction experience at Hughes Helicopter, Inc.

NASA Technical Reports Server (NTRS)

Janakiram, D. S.

1982-01-01

Noise reduction is mostly limited to light helicopters whose noise signature is dominated by their tail rotors. It is primarily hardware oriented. Well known noise reduction techniques such as reduction of rotor speeds with an accompanying increase in solidity to maintain performance, engine noise reduction with the use of exhaust mufflers, and acoustic blanketing of transmission and engine compartment are used. The concept of blade phasing as a means of reducing tail rotor noise is also used. Engine noise (exhaust noise), power train noise and airframe noise becomes important at low rotor tip speeds and means must be found to reduce these noise sources if further noise reductions are desired. The use of a special test rig aids in isolating the various noise sources and arriving at the penalties (performance or payload) involved in quieting them. Significant noise reduction are achieved for the light helicopter with minimum performance or weight penalties because of the dominance of a single noise source (the tail rotor).

Turbine rotor disk health monitoring assessment based on sensor technology and spin tests data.

PubMed

Abdul-Aziz, Ali; Woike, Mark

2013-01-01

The paper focuses on presenting data obtained from spin test experiments of a turbine engine like rotor disk and assessing their correlation to the development of a structural health monitoring and fault detection system. The data were obtained under various operating conditions such as the rotor disk being artificially induced with and without a notch and rotated at a rotational speed of up to 10,000 rpm under balanced and imbalanced state. The data collected included blade tip clearance, blade tip timing measurements, and shaft displacements. Two different sensor technologies were employed in the testing: microwave and capacitive sensors, respectively. The experimental tests were conducted at the NASA Glenn Research Center's Rotordynamics Laboratory using a high precision spin system. Disk flaw observations and related assessments from the collected data for both sensors are reported and discussed.

100-kW hingeless metal wind turbine blade design, analysis and fabrication

NASA Technical Reports Server (NTRS)

Donham, R. E.; Schmidt, J.; Linscott, B. S.

1975-01-01

The design, fabrication and analysis of aluminum wind turbine rotor blades is discussed. The blades are designed to meet criteria established for a 100-kilowatt wind turbine generator operating between 8 and 60-mile-per-hour speeds at 40 revolutions per minute. The design wind speed is 18 miles per hour. Two rotor blades are used on a new facility which includes a hingeless hub and its shaft, gearbox, generator and tower. Experience shows that, for stopped rotors, safe wind speeds are strongly dependent on blade torsional and bending rigidities which the basic D spar structural blade design provides. The 0.25-inch-thick nose skin is brake/bump formed to provide the basic 'D' spar structure for the tapered, twisted blades. Adequate margins for flutter and divergence are predicted from the use of existing, correlated stopped rotor and helicopter rotor analysis programs.

A Review of Sparsity-Based Methods for Analysing Radar Returns from Helicopter Rotor Blades

DTIC Science & Technology

2016-09-01

UNCLASSIFIED A Review of Sparsity-Based Methods for Analysing Radar Returns from Helicopter Rotor Blades Ngoc Hung Nguyen 1, Hai-Tan Tran 2, Kutluyıl...TR–3292 ABSTRACT Radar imaging of rotating blade -like objects, such as helicopter rotors, using narrowband radar has lately been of significant...Methods for Analysing Radar Returns from Helicopter Rotor Blades Executive Summary Signal analysis and radar imaging of fast-rotating objects such as

Integration of dynamic, aerodynamic, and structural optimization of helicopter rotor blades

NASA Technical Reports Server (NTRS)

Peters, David A.

1991-01-01

Summarized here is the first six years of research into the integration of structural, dynamic, and aerodynamic considerations in the design-optimization process for rotor blades. Specifically discussed here is the application of design optimization techniques for helicopter rotor blades. The reduction of vibratory shears and moments at the blade root, aeroelastic stability of the rotor, optimum airframe design, and an efficient procedure for calculating system sensitivities with respect to the design variables used are discussed.

Advanced optical blade tip clearance measurement system

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

A Novel Numerical Approach for Generation and Propagation of Rotor-Stator Interaction Noise

NASA Astrophysics Data System (ADS)

Patel, Krishna

As turbofan engine designs move towards bypass ratios ≥12 and corresponding low pressure ratios, fan rotor blade tip Mach numbers are reduced, leading to rotor-stator interaction becoming an important contributor to tonal fan noise. For future aircraft configurations employing boundary layer ingestion, non-uniform flow enters the fan. The impact of such non-uniform flows on the generation and propagation of rotor-stator interaction tones has yet to be assessed. In this thesis, a novel approach is proposed to numerically predict the generation and propagation of rotor-stator interaction noise with distorted inflow. The approach enables a 42% reduction in computational cost compared to traditional approaches employing a sliding interface between the rotor and stator. Such an interface may distort rotor wakes and can cause non-physical acoustic wave reflections if time steps are not sufficiently small. Computational costs are reduced by modelling the rotor using distributed, volumetric body forces. This eliminates the need for a sliding interface and thus allows a larger time step size. The force model responds to local flow conditions and thus can capture the effects of long-wavelength flow distortions. Since interaction noise is generated by the incidence of the rotor wakes onto the stator vanes, the key challenge is to produce the wakes using a body force field since the rotor blades are not directly modelled. It is shown that such an approach can produce wakes by concentrating the viscous forces along streamtubes in the last 15% chord. The new approach to rotor wake generation is assessed on the GE R4 fan from NASA's Source Diagnostic Test, for which the computed overall aerodynamic performance matches the experiment to within 1%. The rotor blade wakes are generated with widths in excellent agreement and depths in fair agreement with the experiment. An assessment of modal sound power levels computed in the exhaust duct indicates that this approach can be used for predicting downstream propagating interaction noise.

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.

Reusable rocket engine turbopump condition monitoring

NASA Technical Reports Server (NTRS)

Hampson, M. E.; Barkhoudarian, S.

1985-01-01

Significant improvements in engine readiness with attendant reductions in maintenance costs and turnaround times can be achieved with an engine condition monitoring system (CMS). The CMS provides real time health status of critical engine components, without disassembly, through component monitoring with advanced sensor technologies. Three technologies were selected to monitor the rotor bearings and turbine blades: the isotope wear detector and fiber optic deflectometer (bearings), and the fiber optic pyrometer (blades). Signal processing algorithms were evaluated and ranked for their utility in providing useful component health data to unskilled maintenance personnel. Design modifications to current configuration Space Shuttle Main Engine (SSME) high pressure turbopumps and the MK48-F turbopump were developed to incorporate the sensors.

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.

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.

Turbine seal assembly

DOEpatents

Little, David A.

2013-04-16

A seal assembly that limits gas leakage from a hot gas path to one or more disc cavities in a turbine engine. The seal assembly includes a seal apparatus that limits gas leakage from the hot gas path to a respective one of the disc cavities. The seal apparatus comprises a plurality of blade members rotatable with a blade structure. The blade members are associated with the blade structure and extend toward adjacent stationary components. Each blade member includes a leading edge and a trailing edge, the leading edge of each blade member being located circumferentially in front of the blade member's corresponding trailing edge in a direction of rotation of the turbine rotor. The blade members are arranged such that a space having a component in a circumferential direction is defined between adjacent circumferentially spaced blade members.

Data collection and analysis software development for rotor dynamics testing in spin laboratory

NASA Astrophysics Data System (ADS)

Abdul-Aziz, Ali; Arble, Daniel; Woike, Mark

2017-04-01

Gas turbine engine components undergo high rotational loading another complex environmental conditions. Such operating environment leads these components to experience damages and cracks that can cause catastrophic failure during flights. There are traditional crack detections and health monitoring methodologies currently being used which rely on periodic routine maintenances, nondestructive inspections that often times involve engine and components dis-assemblies. These methods do not also offer adequate information about the faults, especially, if these faults at subsurface or not clearly evident. At NASA Glenn research center, the rotor dynamics laboratory is presently involved in developing newer techniques that are highly dependent on sensor technology to enable health monitoring and prediction of damage and cracks in rotor disks. These approaches are noninvasive and relatively economical. Spin tests are performed using a subscale test article mimicking turbine rotor disk undergoing rotational load. Non-contact instruments such as capacitive and microwave sensors are used to measure the blade tip gap displacement and blade vibrations characteristics in an attempt develop a physics based model to assess/predict the faults in the rotor disk. Data collection is a major component in this experimental-analytical procedure and as a result, an upgrade to an older version of the data acquisition software which is based on LabVIEW program has been implemented to support efficiently running tests and analyze the results. Outcomes obtained from the tests data and related experimental and analytical rotor dynamics modeling including key features of the updated software are presented and discussed.

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

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

Wind energy converter with high-speed vertical axis rotor and straight rotor blades

NASA Astrophysics Data System (ADS)

Zelck, G.

1982-11-01

Complete documents for a wind energy converter with a vertical axis rotor and straight blades (H-rotor) were developed. The 2 blade rotor with rigid and rectangular air foils in wooden construction reaches the nominal output of 75 KVA from 11,4 m/sec. wind velocity onwards. The development activities are supported by wind tunnel and component tests. The final design selected was based upon previous development work. Trade offs show that the design is more advantageous compared to other designs. The use of wood as a material for the rotary and horizontal blade supports gives positive result.

Approximate analysis of containment/deflection ring responses to engine rotor fragment impact.

NASA Technical Reports Server (NTRS)

Wu, R. W.-H.; Witmer, E. A.

1973-01-01

The transient responses of containment and/or deflection rings to impact from an engine rotor-blade fragment are analyzed. Energy and momentum considerations are employed in an approximate analysis to predict the collision-induced velocities which are imparted to the fragment and to the affected ring segment. This collision analysis is combined with the spatial finite-element representation of the ring and a temporal finite-difference solution procedure to predict the resulting large transient elastic-plastic deformations of containment/deflection rings. Some comparisons with experimental data are given.

A Study on Aircraft Structure and Jet Engine

NASA Astrophysics Data System (ADS)

Park, Gil Moon; Park, Hwan Kyu; Kim, Jong Il; Kim, Jin Won; Kim, Jin Heung; Lee, Moo Seok; Chung, Nak Kyu

1985-12-01

The one of critical factor in gas turbine engine performance is high turbine inlet gas temperature. Therefore, the turbine rotor has so many problems which must be considered such as the turbine blade cooling, thermal stress of turbine disk due to severe temperature gradient, turbine rotor tip clearance, under the high operation temperature. The purpose of this study is to provide the temperature distribution and heat flux in turbine disk which is required to considered premensioned problem by the Finite Difference Method and the Finite Element Methods on the steady state condition.

JT15D simulated flight data evaluation

NASA Technical Reports Server (NTRS)

Holm, R. G.

1984-01-01

The noise characteristics of the JT15D turbofan engine was analyzed with the objectives of: (1) assessing the state-of-art ability to simulate flight acoustic data using test results acquired in wind tunnel and outdoor (turbulence controlled) environments; and (2) predicting the farfield noise directivity of the blade passage frequency (BPF) tonal components using results from rotor blade mounted dynamic pressure instrumentation. Engine rotor tip speeds at subsonic, transonic, and supersonic conditions were evaluated. The ability to simulate flight results was generally within 2-3 dB for both outdoor and wind tunnel acoustic results. Some differences did occur in the broadband noise level and in the multiple-pure-tone harmonics at supersonic tip speeds. The prediction of blade passage frequency tone directivity from dynamic pressure measurements was accomplished for the three tip speed conditions. Predictions were made of the random and periodic components of the tone directivity. The technique for estimating the random tone component used hot wire data to establish a correlation between dynamic pressure and turbulence intensity. This prediction overestimated the tone level by typically 10 dB with the greatest overestimates occurring at supersonic conditions.

Analytical prediction of the unsteady lift on a rotor caused by downstream struts

NASA Technical Reports Server (NTRS)

Taylor, A. C., III; Ng, W. F.

1987-01-01

A two-dimensional, inviscid, incompressible procedure is presented for predicting the unsteady lift on turbomachinery blades caused by the upstream potential disturbance of downstream flow obstructions. Using the Douglas-Neumann singularity superposition potential flow computer program to model the downstream flow obstructions, classical equations of thin airfoil theory are then employed, to compute the unsteady lift on the upstream rotor blades. The method is applied to a particular geometry which consists of a rotor, a downstream stator, and downstream struts which support the engine casing. Very good agreement between the Douglas-Neumann program and experimental measurements was obtained for the downstream stator-strut flow field. The calculations for the unsteady lift due to the struts were in good agreement with the experiments in showing that the unsteady lift due to the struts decays exponentially with increased axial separation of the rotor and the struts. An application of the method showed that for a given axial spacing between the rotor and the strut, strut-induced unsteady lift is a very weak function of the axial or circumferential position of the stator.

Wind turbine rotor hub and teeter joint

DOEpatents

Coleman, Clint; Kurth, William T.; Jankowski, Joseph

1994-10-11

A rotor hub is provided for coupling a wind turbine rotor blade and a shaft. The hub has a yoke with a body which is connected to the shaft, and extension portions which are connected to teeter bearing blocks, each of which has an aperture. The blocks are connected to a saddle which envelops the rotor blade by one or two shafts which pass through the apertures in the bearing blocks. The saddle and blade are separated by a rubber interface which provides for distribution of stress over a larger portion of the blade. Two teeter control mechanisms, which may include hydraulic pistons and springs, are connected to the rotor blade and to the yoke at extension portions. These control mechanisms provide end-of-stroke damping, braking, and stiffness based on the teeter angle and speed of the blade.

Data Reduction Procedures for Laser Velocimeter Measurements in Turbomachinery Rotors

NASA Technical Reports Server (NTRS)

Lepicovsky, Jan

1994-01-01

Blade-to-blade velocity distributions based on laser velocimeter data acquired in compressor or fan rotors are increasingly used as benchmark data for the verification and calibration of turbomachinery computational fluid dynamics (CFD) codes. Using laser Doppler velocimeter (LDV) data for this purpose, however, must be done cautiously. Aside from the still not fully resolved issue of the seed particle response in complex flowfields, there is an important inherent difference between CFD predictions and LDV blade-to-blade velocity distributions. CFD codes calculate velocity fields for an idealized rotor passage. LDV data, on the other hand, stem from the actual geometry of all blade channels in a rotor. The geometry often varies from channel to channel as a result of manufacturing tolerances, assembly tolerances, and incurred operational damage or changes in the rotor individual blades.

Fatigue Lifespan of Engine Box Influenced by Fan Blade Out

NASA Astrophysics Data System (ADS)

Qiu, Ju; Shi, Jingwei; Su, Huaizhong; Zhang, Jinling; Feng, Juan; Shi, Qian; Tian, Xiaoyu

2017-11-01

This provides precious experience and reliable reference data for future design. This paper introduces the analysis process of Fan-blade-out, and considers the effect of windmill load on the fatigue lifespan of the case. According to Extended Operations (ETOPS) in the airworthiness regulations, the fatigue crack of it is analyzed by the unbalanced rotor load, during FBO. Compared with the lifespan in normal work of the engine, this research provides valuable design experience and reliable reference data for the case design in the near future.

Multicyclic control for helicopters - Research in progress at Ames Research Center

NASA Technical Reports Server (NTRS)

Mccloud, J. L., III

1980-01-01

The term multicyclic control describes a blade pitch control technique used by helicopter designers to alleviate vibration in rotorcraft. Because rotor-induced vibrations are periodic, a multicyclic system, synchronized to the main rotor's azimuth position, is suitable. Many types of rotors - ranging from the jet-flap and circulation-control rotors to the conventional full-blade feathering rotors - have utilized multicyclic control. Multicyclic control systems may be designed to reduce blade-bending stresses, to reduce rotor-induced vibration, and to improve rotor performance. Rotor types are reviewed, primarily to highlight their differences. The increased use of composites in blade construction is seen to indicate that vibration alleviation will be the prime focus of multicyclic control. Adaptive feedback control systems, which also incorporate gust alleviation, are considered to be the ultimate application of multicyclic control.

Development of an Active Twist Rotor for Wind: Tunnel Testing (NLPN97-310

NASA Technical Reports Server (NTRS)

Cesnik, Carlos E. S.; Shin, SangJoon; Hagood, Nesbitt W., IV

1998-01-01

The development of the Active Twist Rotor prototype blade for hub vibration and noise reduction studies is presented in this report. Details of the modeling, design, and manufacturing are explored. The rotor blade is integrally twisted by direct strain actuation. This is accomplished by distributing embedded piezoelectric fiber composites along the span of the blade. The development of the analysis framework for this type of active blade is presented. The requirements for the prototype blade, along with the final design results are also presented. A detail discussion on the manufacturing aspects of the prototype blade is described. Experimental structural characteristics of the prototype blade compare well with design goals, and preliminary bench actuation tests show lower performance than originally predicted. Electrical difficulties with the actuators are also discussed. The presented prototype blade is leading to a complete fully articulated four-blade active twist rotor system for future wind tunnel tests.

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.

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.

Main rotor-body action for virtual blades model

NASA Astrophysics Data System (ADS)

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

2018-06-01

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

A Comprehensive Solution of the Problems of Ensuring the Strength of Gas Turbine Engine Compressor at the Design Stage

NASA Astrophysics Data System (ADS)

Vedeneev, V. V.; Kolotnikov, M. E.; Mossakovskii, P. A.; Kostyreva, L. A.; Abdukhakimov, F. A.; Makarov, P. V.; Pyhalov, A. A.; Dudaev, M. A.

2018-01-01

In this paper we present a complex numerical workflow for analysis of blade flutter and high-amplitude resonant oscillations, impenetrability of casing if the blade is broken off, and the rotor reaction to the blade detachment and following misbalance, with the assessment of a safe flight possibility at the auto-rotation regime. All the methods used are carefully verified by numerical convergence study and correlations with experiments. The use of the workflow developed significantly improves the efficiency of the design process of modern jet engine compressors. It ensures a significant reduction of time and cost of the compressor design with the required level of strength and durability.

Design, analysis, optimization and control of rotor tip flows

NASA Astrophysics Data System (ADS)

Maesschalck, Cis Guy M. De

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

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.

Adaptor assembly for coupling turbine blades to rotor disks

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

Garcia-Crespo, Andres Jose; Delvaux, John McConnell

2014-09-23

An adaptor assembly for coupling a blade root of a turbine blade to a root slot of a rotor disk is described. The adaptor assembly includes a turbine blade having a blade root and an adaptor body having an adaptor root. The adaptor body defines a slot having an open end configured to receive the blade root of the turbine blade such that the adaptor root of the adaptor body and the blade root of the turbine blade are adjacent to one another when the blade root of the turbine blade is positioned within the slot. Both the adaptor rootmore » of the adaptor body and the blade root of the turbine blade are configured to be received within the root slot of the rotor disk.« less

Rotor blade boundary layer measurement hardware feasibility demonstration

NASA Technical Reports Server (NTRS)

Clark, D. R.; Lawton, T. D.

1972-01-01

A traverse mechanism which allows the measurement of the three dimensional boundary layers on a helicopter rotor blade has been built and tested on a full scale rotor to full scale conditions producing centrifugal accelerations in excess of 400 g and Mach numbers of 0.6 and above. Boundary layer velocity profiles have been measured over a range of rotor speeds and blade collective pitch angles. A pressure scanning switch and transducer were also tested on the full scale rotor and found to be insensitive to centrifugal effects within the normal main rotor operating range. The demonstration of the capability to measure boundary layer behavior on helicopter rotor blades represents the first step toward obtaining, in the rotating system, data of a quality comparable to that already existing for flows in the fixed system.

Welding blades to rotors

NASA Technical Reports Server (NTRS)

Hoklo, K. H.; Moore, T. J. (Inventor)

1973-01-01

A process is described to form T-joints between dissimilar thickness parts by magnetic force upset welding. This type of resistance welding is used to join compressor and turbine parts which thereby reduces the weight and cost of jet engines.

Calculated effects of turbine rotor-blade cooling-air flow, altitude, and compressor bleed point on performance of a turbojet engine

NASA Technical Reports Server (NTRS)

Arne, Vernon L; Nachtigall, Alfred J

1951-01-01

Effects of air-cooling turbine rotor blades on performance of a turbojet engine were calculated for a range of altitudes from sea level to 40,000 feet and a range of coolant flows up to 3 percent of compressor air flow, for two conditions of coolant bleed from the compressor. Bleeding at required coolant pressure resulted in a sea-level thrust reduction approximately twice the percentage coolant flow and in an increase in specific fuel consumption approximately equal to percentage coolant flow. For any fixed value of coolant flow ratio the percentage thrust reduction and percentage increase in specific fuel consumption decreased with altitude. Bleeding coolant at the compressor discharge resulted in an additional 1 percent loss in performance at sea level and in smaller increase in loss of performance at higher altitudes.

Self-Tuning Impact Damper for Rotating Blades

NASA Technical Reports Server (NTRS)

Pufy, Kirsten P. (Inventor); Brown, Gerald V. (Inventor); Bagley, Ronald L. (Inventor)

2004-01-01

A self-tuning impact damper is disclosed that absorbs and dissipates vibration energy in the blades of rotors in compressors and/or turbines thereby dramatically extending their service life and operational readiness. The self-tuning impact damper uses the rotor speed to tune the resonant frequency of a rattling mass to an engine order excitation frequency. The rating mass dissipates energy through collisions between the rattling mass and the walls of a cavity of the self-tuning impact damper, as well as though friction between the rattling mass and the base of the cavity. In one embodiment, the self-tuning impact damper has a ball-in-trough configuration with tire ball serving as the rattling mass.

77 FR 46185 - United States v. United Technologies Corporation and Goodrich Corporation; Proposed Final...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-08-02

... for generating power for all the in-flight systems that run on electricity, including pumping breathable air into the fuselage, operating the lights, and running the navigation and communication... turning a propeller blade on a turboprop engine, a rotor shaft on a turboshaft engine, or a fan in front...

Structural Evaluation of Exo-Skeletal Engine Fan Blades

NASA Technical Reports Server (NTRS)

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

2003-01-01

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

The influence of engine/transmission/governor on tilting proprotor aircraft dynamics

NASA Technical Reports Server (NTRS)

Johnson, W.

1975-01-01

An analytical model is developed for the dynamics of a tilting proprotor aircraft engine and drive train, including a rotor speed governor and interconnect shaft. The dynamic stability of a proprotor and cantilever wing is calculated, including the engine-transmission-governor model. It is concluded that the rotor behaves much as if windmilling as far as its dynamic behavior is concerned, with some influence of the turboshaft engine inertia and damping. The interconnect shaft has a significant influence on the antisymmetric dynamics of proprotor aircraft. The proprotor aerodynamics model is extended to include reverse flow, and a refinement on the method used to calculate the kinematic pitch-bending coupling of the blade is developed.

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

NASA Technical Reports Server (NTRS)

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

2001-01-01

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

The elaboration of a new family of helicopter blade profiles

NASA Technical Reports Server (NTRS)

Thibert, J. J.

1981-01-01

An airfoil family of helicopter rotor blades was designed. Three airfoils with thickness to chord ratios of 12, 9, and 7% were designed. Their improved performance in two dimensional rotor mockup wind tunnel tests led to testing of the tapered blades on four bladed rotors in a wind tunnel and flight tests on the Dauphin series of helicopters, confirming the expected gains.

In-plane inertial coupling in tuned and severely mistuned bladed disks

NASA Technical Reports Server (NTRS)

Crawley, E. F.

1982-01-01

A model has been developed and verified for blade-disk-shaft coupling in rotors due to the in-plane rigid body modes of the disk. An analytic model has been developed which couples the in-plane rigid body modes of the disk on an elastic shaft with the blade bending modes. Bench resonance test were carried out on the M.I.T. Compressor Rotor, typical of research rotors with flexible blades and a thick rigid disk. When the rotor was carefully tuned, the structural coupling of the blades by the disks was confined to zero and one nodal diameter modes, whose modal frequencies were greater than the blade cantilever frequency. In the case of the tuned rotor, and in two cases where severe mistuning was intentionally introduced, agreement between the predicted and observed natural frequencies is excellent. The analytic model was then extended to include the effects of constant angular rotation of the disk.

Stiffness Characteristics of Composite Rotor Blades With Elastic Couplings

NASA Technical Reports Server (NTRS)

Piatak, David J.; Nixon, Mark W.; Kosmatka, John B.

1997-01-01

Recent studies on rotor aeroelastic response and stability have shown the beneficial effects of incorporating elastic couplings in composite rotor blades. However, none of these studies have clearly identified elastic coupling limits and the effects of elastic couplings on classical beam stiffnesses of representative rotor blades. Knowledge of these limits and effects would greatly enhance future aeroelastic studies involving composite rotor blades. The present study addresses these voids and provides a preliminary design database for investigators who may wish to study the effects of elastic couplings on representative blade designs. The results of the present study should provide a basis for estimating the potential benefits associated with incorporating elastic couplings without the need for first designing a blade cross section and then performing a cross-section analysis to obtain the required beam section properties as is customary in the usual one-dimensional beam-type approach.

Single-crystal-material-based induced-shear actuation for vibration reduction of helicopters with composite rotor system

NASA Astrophysics Data System (ADS)

Pawar, Prashant M.; Jung, Sung Nam

2008-12-01

In this study, an assessment is made for the helicopter vibration reduction of composite rotor blades using an active twist control concept. Special focus is given to the feasibility of implementing the benefits of the shear actuation mechanism along with elastic couplings of composite blades for achieving maximum vibration reduction. The governing equations of motion for composite rotor blades with surface bonded piezoceramic actuators are obtained using Hamilton's principle. The equations are then solved for dynamic response using finite element discretization in the spatial and time domains. A time domain unsteady aerodynamic theory with free wake model is used to obtain the airloads. A newly developed single-crystal piezoceramic material is introduced as an actuator material to exploit its superior shear actuation authority. Seven rotor blades with different elastic couplings representing stiffness properties similar to stiff-in-plane rotor blades are used to investigate the hub vibration characteristics. The rotor blades are modeled as a box beam with actuator layers bonded on the outer surface of the top and bottom of the box section. Numerical results show that a notable vibration reduction can be achieved for all the combinations of composite rotor blades. This investigation also brings out the effect of different elastic couplings on various vibration-reduction-related parameters which could be useful for the optimal design of composite helicopter blades.

Perturbation solutions for the influence of forward flight on helicopter rotor flapping stability

NASA Technical Reports Server (NTRS)

Johnson, W.

1974-01-01

The stability of the flapping motion of a helicopter rotor blade in forward flight is investigated, using a perturbation technique which gives analytic expressions for the eigenvalues, including the influence of the periodic aerodynamic forces in forward flight. The perturbation solutions are based on small advance ratio (the ratio of the helicopter forward speed to the rotor tip speed). The rotor configurations considered are a single, independent blade; a teetering rotor; a gimballed rotor with three, four, and five or more blades; and a rotor with N independent blades. The constant coefficient approximation with the equations and degrees of freedom in the nonrotating frame represents the flap dynamic reasonably well for the lower frequency modes, although it cannot, of course, be completely correct. The transfer function of the rotor flap response to sinusoidal pitch input is examined, as an alternative to the eigenvalues as a representation of the dynamic characteristics of the flap motion.

Hover Testing of the NASA/Army/MIT Active Twist Rotor Prototype Blade

NASA Technical Reports Server (NTRS)

Wilbur, Matthew L.; Yeager, William T., Jr.; Wilkie, W. Keats; Cesnik, Carlos E. S.; Shin, Sangloon

2000-01-01

Helicopter rotor individual blade control promises to provide a mechanism for increased rotor performance and reduced rotorcraft vibrations and noise. Active material methods, such as piezoelectrically actuated trailing-edge flaps and strain-induced rotor blade twisting, provide a means of accomplishing individual blade control without the need for hydraulic power in the rotating system. Recent studies have indicated that controlled strain induced blade twisting can be attained using piezoelectric active fiber composite technology. In order to validate these findings experimentally, a cooperative effort between NASA Langley Research Center, the Army Research Laboratory, and the MIT Active Materials and Structures Laboratory has been developed. As a result of this collaboration an aeroelastically-scaled active-twist model rotor blade has been designed and fabricated for testing in the heavy gas environment of the Langley Transonic Dynamics Tunnel (TDT). The results of hover tests of the active-twist prototype blade are presented in this paper. Comparisons with applicable analytical predictions of active-twist frequency response in hovering flight are also presented.

Reliability analysis of forty-five strain-gage systems mounted on the first fan stage of a YF-100 engine

NASA Technical Reports Server (NTRS)

Holanda, R.; Frause, L. M.

1977-01-01

The reliability of 45 state-of-the-art strain gage systems under full scale engine testing was investigated. The flame spray process was used to install 23 systems on the first fan rotor of a YF-100 engine; the others were epoxy cemented. A total of 56 percent of the systems failed in 11 hours of engine operation. Flame spray system failures were primarily due to high gage resistance, probably caused by high stress levels. Epoxy system failures were principally erosion failures, but only on the concave side of the blade. Lead-wire failures between the blade-to-disk jump and the control room could not be analyzed.

Turbine Rotor Disk Health Monitoring Assessment Based on Sensor Technology and Spin Tests Data

PubMed Central

2013-01-01

The paper focuses on presenting data obtained from spin test experiments of a turbine engine like rotor disk and assessing their correlation to the development of a structural health monitoring and fault detection system. The data were obtained under various operating conditions such as the rotor disk being artificially induced with and without a notch and rotated at a rotational speed of up to 10,000 rpm under balanced and imbalanced state. The data collected included blade tip clearance, blade tip timing measurements, and shaft displacements. Two different sensor technologies were employed in the testing: microwave and capacitive sensors, respectively. The experimental tests were conducted at the NASA Glenn Research Center's Rotordynamics Laboratory using a high precision spin system. Disk flaw observations and related assessments from the collected data for both sensors are reported and discussed. PMID:23844396

Core compressor exit stage study. Volume 1: Blading design. [turbofan engines

NASA Technical Reports Server (NTRS)

Wisler, D. C.

1977-01-01

A baseline compressor test stage was designed as well as a candidate rotor and two candidate stators that have the potential of reducing endwall losses relative to the baseline stage. These test stages are typical of those required in the rear stages of advanced, highly-loaded core compressors. The baseline Stage A is a low-speed model of Stage 7 of the 10 stage AMAC compressor. Candidate Rotor B uses a type of meanline in the tip region that unloads the leading edge and loads the trailing edge relative to the baseline Rotor A design. Candidate Stator B embodies twist gradients in the endwall region. Candidate Stator C embodies airfoil sections near the endwalls that have reduced trailing edge loading relative to Stator A. Tests will be conducted using four identical stages of blading so that the designs described will operate in a true multistage environment.

Blade Displacement Measurement Technique Applied to a Full-Scale Rotor Test

NASA Technical Reports Server (NTRS)

Abrego, Anita I.; Olson, Lawrence E.; Romander, Ethan A.; Barrows, Danny A.; Burner, Alpheus W.

2012-01-01

Blade displacement measurements using multi-camera photogrammetry were acquired during the full-scale wind tunnel test of the UH-60A Airloads rotor, conducted in the National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel. The objectives were to measure the blade displacement and deformation of the four rotor blades as they rotated through the entire rotor azimuth. These measurements are expected to provide a unique dataset to aid in the development and validation of rotorcraft prediction techniques. They are used to resolve the blade shape and position, including pitch, flap, lag and elastic deformation. Photogrammetric data encompass advance ratios from 0.15 to slowed rotor simulations of 1.0, thrust coefficient to rotor solidity ratios from 0.01 to 0.13, and rotor shaft angles from -10.0 to 8.0 degrees. An overview of the blade displacement measurement methodology and system development, descriptions of image processing, uncertainty considerations, preliminary results covering static and moderate advance ratio test conditions and future considerations are presented. Comparisons of experimental and computational results for a moderate advance ratio forward flight condition show good trend agreements, but also indicate significant mean discrepancies in lag and elastic twist. Blade displacement pitch measurements agree well with both the wind tunnel commanded and measured values.

The effect of blade pitch in the rotor hydrodynamics of a cross-flow turbine

NASA Astrophysics Data System (ADS)

Somoano, Miguel; Huera-Huarte, Francisco

2016-11-01

In this work we will show how the hydrodynamics of the rotor of a straight-bladed Cross-Flow Turbine (CFT) are affected by the Tip Speed Ratio (TSR), and the blade pitch angle imposed to the rotor. The CFT model used in experiments consists of a three-bladed (NACA-0015) vertical axis turbine with a chord (c) to rotor diameter (D) ratio of 0.16. Planar Digital Particle Image Velocimetry (DPIV) was used, with the laser sheet aiming at the mid-span of the blades, illuminating the inner part of the rotor and the near wake of the turbine. Tests were made by forcing the rotation of the turbine with a DC motor, which provided precise control of the TSR, while being towed in a still-water tank at a constant Reynolds number of 61000. A range of TSRs from 0.7 to 2.3 were covered for different blade pitches, ranging from 8° toe-in to 16° toe-out. The interaction between the blades in the rotor will be discussed by examining dimensionless phase-averaged vorticity fields in the inner part of the rotor and mean velocity fields in the near wake of the turbine. Supported by the Spanish Ministry of Economy and Competitiveness, Grant BES-2013-065366 and project DPI2015-71645-P.

Propeller blade retention system

NASA Technical Reports Server (NTRS)

Elston, III, Sidney B. (Inventor); Simon, III, Victor H. (Inventor); Tseng, Wu-Yang (Inventor); Butler, Lawrence (Inventor)

1993-01-01

The invention concerns the mounting of propeller blades to a ring-shaped rotor. The blades are of the variable pitch type, and the shank of each blade extends through a respective hole in the rotor. Each hole contains an annular shelf which is fastened to the wall of the hole and surrounds each shank. Each shank bears a pair of bearing races which sandwich the annular shelf in order to connect the blade to the rotor. Bearing rollers are positioned between the annular shelf and the bearing races.

Recent radial turbine research at the NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Rohlik, H. E.; Kofskey, M. G.

1971-01-01

The high efficiencies of small radial turbines led to their application in space power systems and numerous APU and shaft power engines. Experimental and analytical work associated with these systems included examination of blade-shroud clearance, blade loading, and exit diffuser design. Results indicate high efficiency over a wide range of specific speed and also insensitivity to clearance and blade loading in the radial part of the rotor. The exit diffuser investigation indicated that a conventional conical outer wall may not provide the velocity variation consistent with minimum overall diffuser loss.

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

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.

Performance ‘S’ Type Savonius Wind Turbine with Variation of Fin Addition on Blade

NASA Astrophysics Data System (ADS)

Pamungkas, S. F.; Wijayanto, D. S.; Saputro, H.; Widiastuti, I.

2018-01-01

Wind power has been receiving attention as the new energy resource in addressing the ecological problems of burning fossil fuels. Savonius wind rotor is a vertical axis wind turbines (VAWT) which has relatively simple structure and low operating speed. These characteristics make it suitable for areas with low average wind speed as in Indonesia. To identify the performance of Savonius rotor in generating electrical energy, this research experimentally studied the effect of fin addition for the ‘S’ shape of Savonius VAWT. The fin is added to fill the space in the blade in directing the wind flow. This rotor has two turbine blades, a rotor diameter of 1.1 m and rotor height of 1.4 m, used pulley transmission system with 1:4.2 multiplication ratio, and used a generator type PMG 200 W. The research was conducted during dry season by measuring the wind speed in the afternoon. The average wind speed in the area is 2.3 m/s with the maximum of 4.5 m/s. It was found that additional fin significantly increase the ability of Savonius rotor VAWT to generate electrical energy shown by increasing of electrical power. The highest power generated is 13.40 Watt at a wind speed of 4.5 m/s by adding 1 (one) fin in the blade. It increased by 22.71% from the rotor blade with no additional fin. However, increasing number of fins in the blade was not linearly increase the electrical power generated. The wind rotor blade with 4 additional fins is indicated has the lowest performance, generating only 10.80 Watt electrical power, accounted lower than the one generated by no fin-rotor blade. By knowing the effect of the rotor shape, the rotor dimension, the addition of fin, transmission, and generator used, it is possible to determine alternative geometry design in increasing the electrical power generated by Savonius wind turbine.

Macroscopic balance model for wave rotors

NASA Technical Reports Server (NTRS)

Welch, Gerard E.

1996-01-01

A mathematical model for multi-port wave rotors is described. The wave processes that effect energy exchange within the rotor passage are modeled using one-dimensional gas dynamics. Macroscopic mass and energy balances relate volume-averaged thermodynamic properties in the rotor passage control volume to the mass, momentum, and energy fluxes at the ports. Loss models account for entropy production in boundary layers and in separating flows caused by blade-blockage, incidence, and gradual opening and closing of rotor passages. The mathematical model provides a basis for predicting design-point wave rotor performance, port timing, and machine size. Model predictions are evaluated through comparisons with CFD calculations and three-port wave rotor experimental data. A four-port wave rotor design example is provided to demonstrate model applicability. The modeling approach is amenable to wave rotor optimization studies and rapid assessment of the trade-offs associated with integrating wave rotors into gas turbine engine systems.

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

NASA Technical Reports Server (NTRS)

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

2004-01-01

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

Compressible, unsteady lifting-surface theory for a helicopter rotor in forward flight

NASA Technical Reports Server (NTRS)

Runyan, H. L.; Tai, H.

1985-01-01

A lifting-surface theory has been developed for a helicopter rotor in forward flight for compressible and incompressible flow. The method utilizes the concept of the linearized acceleration potential and makes use of the doublet lattice procedure. Calculations demonstrating the application of the method are given in terms of the lift distribution on a one-bladed rotor, a two-bladed rotor, and a rotor with swept-forward and swept-back tips. Also, the lift on a rotor vibrating in a pitching mode at 4 per revolution is given. Compressibility effects and interference effects for a two-bladed rotor are discussed.

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.

Prediction of SA 349/2 GV blade loads in high speed flight using several rotor analyses

NASA Technical Reports Server (NTRS)

Gaubert, Michel; Yamauchi, Gloria K.

1987-01-01

The influence of blade dynamics, dynamic stall, and transonic aerodynamics on the predictions of rotor loads in high-speed flight are presented. Data were obtained from an Aerospatiale Gazelle SA 349/2 helicopter with three Grande Vitesse blades. Several analyses are used for this investigation. First, blade dynamics effects on the correlation are studied using three rotor analyses which differ mainly in the method of calculating the blade elastic response. Next, an ONERA dynamic stall model is used to predict retreating blade stall. Finally, advancing blade aerodynamic loads are calculated using a NASA-developed rotorcraft analysis coupled with two transonic finite-difference analyses.

Wind tunnel investigation of rotor lift and propulsive force at high speed: Data analysis

NASA Technical Reports Server (NTRS)

Mchugh, F.; Clark, R.; Soloman, M.

1977-01-01

The basic test data obtained during the lift-propulsive force limit wind tunnel test conducted on a scale model CH-47b rotor are analyzed. Included are the rotor control positions, blade loads and six components of rotor force and moment, corrected for hub tares. Performance and blade loads are presented as the rotor lift limit is approached at fixed levels of rotor propulsive force coefficients and rotor tip speeds. Performance and blade load trends are documented for fixed levels of rotor lift coefficient as propulsive force is increased to the maximum obtainable by the model rotor. Test data is also included that defines the effect of stall proximity on rotor control power. The basic test data plots are presented in volumes 2 and 3.

Wireless Inductive Power Device Suppresses Blade Vibrations

NASA Technical Reports Server (NTRS)

Morrison, Carlos R.; Provenza, Andrew J.; Choi, Benjamin B.; Bakhle, Milind A.; Min, James B.; Stefko, George L.; Duffy, Kirsten P.; Fougers, Alan J.

2011-01-01

Vibration in turbomachinery can cause blade failures and leads to the use of heavier, thicker blades that result in lower aerodynamic efficiency and increased noise. Metal and/or composite fatigue in the blades of jet engines has resulted in blade destruction and loss of lives. Techniques for suppressing low-frequency blade vibration, such as gtuned circuit resistive dissipation of vibratory energy, h or simply "passive damping," can require electronics incorporating coils of unwieldy dimensions and adding unwanted weight to the rotor. Other approaches, using vibration-dampening devices or damping material, could add undesirable weight to the blades or hub, making them less efficient. A wireless inductive power device (WIPD) was designed, fabricated, and developed for use in the NASA Glenn's "Dynamic Spin Rig" (DSR) facility. The DSR is used to simulate the functionality of turbomachinery. The relatively small and lightweight device [10 lb (approx.=4.5 kg)] replaces the existing venerable and bulky slip-ring. The goal is the eventual integration of this technology into actual turbomachinery such as jet engines or electric power generators, wherein the device will facilitate the suppression of potentially destructive vibrations in fan blades. This technology obviates slip rings, which require cooling and can prove unreliable or be problematic over time. The WIPD consists of two parts: a remote element, which is positioned on the rotor and provides up to 100 W of electrical power to thin, lightweight piezoelectric patches strategically placed on/in fan blades; and a stationary base unit that wirelessly communicates with the remote unit. The base unit supplies inductive power, and also acts as an input and output corridor for wireless measurement, and active control command to the remote unit. Efficient engine operation necessitates minimal disturbance to the gas flow across the turbine blades in any effort to moderate blade vibration. This innovation makes it possible to moderate vibration on or in turbomachinery blades by providing 100 W of wireless electrical power and actuation control to thin, lightweight vibration-suppressing piezoelectric patches (eight actuation and eight sensor patches in this prototype, for a total of 16 channels) positioned strategically on the surface of, or within, titanium fan blades, or embedded in composite fan blades. This approach moves significantly closer to the ultimate integration of "active" vibration suppression technology into jet engines and other turbomachinery devices such as turbine electrical generators used in the power industry. The novel feature of this device is in its utilization of wireless technology to simultaneously sense and actively control vibration in rotating or stationary turbomachinery blades using piezoelectric patches. In the past, wireless technology was used solely for sensing and diagnostics. This technology, however, will accomplish much more, in terms of simultaneously sensing, suppressing blade vibration, and making it possible for detailed study of vibration impact in turbomachinery blades.

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.

Extended aeroelastic analysis for helicopter rotors with prescribed hub motion and blade appended penduluum vibration absorbers

NASA Technical Reports Server (NTRS)

Bielawa, R. L.

1984-01-01

The mathematical development for the expanded capabilities of the G400 rotor aeroelastic analysis was examined. The G400PA expanded analysis simulates the dynamics of all conventional rotors, blade pendulum vibration absorbers, and the higher harmonic excitations resulting from prescribed vibratory hub motions and higher harmonic blade pitch control. The methodology for modeling the unsteady stalled airloads of two dimensional airfoils is discussed. Formulations for calculating the rotor impedance matrix appropriate to the higher harmonic blade excitations are outlined. This impedance matrix, and the associated vibratory hub loads, are the rotor dynamic characteristic elements for use in the simplified coupled rotor/fuselage vibration analysis (SIMVIB). Updates to the development of the original G400 theory, program documentation, user instructions and information are presented.

Investigation of the Effect of Blade Sweep on Rotor Vibratory Loads

NASA Technical Reports Server (NTRS)

Tarzanin, F. J., Jr.; Vlaminck, R. R.

1983-01-01

The effect of helicopter rotor blade planform sweep on rotor vibratory hub, blade, and control system loads has been analytically investigated. The importance of sweep angle, sweep initiation radius, flap bending stiffness and torsion bending stiffness is discussed. The mechanism by which sweep influences the vibratory hub loads is investigated.

Composite rotor blades for large wind energy installations

NASA Technical Reports Server (NTRS)

Kussmann, A.; Molly, J.; Muser, D.

1980-01-01

The design of large wind power systems in Germany is reviewed with attention given to elaboration of the total wind energy system, aerodynamic design of the rotor blade, and wind loading effects. Particular consideration is given to the development of composite glass fiber/plastic or carbon fiber/plastic rotor blades for such installations.

Blade loss transient dynamic analysis of turbomachinery

NASA Technical Reports Server (NTRS)

Stallone, M. J.; Gallardo, V.; Storace, A. F.; Bach, L. J.; Black, G.; Gaffney, E. F.

1982-01-01

This paper reports on work completed to develop an analytical method for predicting the transient non-linear response of a complete aircraft engine system due to the loss of a fan blade, and to validate the analysis by comparing the results against actual blade loss test data. The solution, which is based on the component element method, accounts for rotor-to-casing rubs, high damping and rapid deceleration rates associated with the blade loss event. A comparison of test results and predicted response show good agreement except for an initial overshoot spike not observed in test. The method is effective for analysis of large systems.

Development of a Highly Loaded Rotor Blade for Steam Turbines

NASA Astrophysics Data System (ADS)

Segawa, Kiyoshi; Shikano, Yoshio; Tsubouchi, Kuniyoshi; Shibashita, Naoaki

Turbine manufacturers have been concerned about efficient utilization of limited energy resources and prevention of environmental pollution. For steam turbine power plants, a higher efficiency gain is necessary to reduce the fuel consumption rate. Blade configurations have been studied for reductions of profile loss and endwall loss that lead to decreased steam turbine internal efficiency, by applying recent aerodynamic technologies based on advanced numerical analysis methods. This paper discusses increase of pitch-chord ratio by 14% (reduction of rotor blade numbers by 14%) and increased blade aerodynamic loading without deterioration of performance. A new rotor cascade is found which improves blade performance, especially at the root section where the reduction in the energy loss coefficient is about 40%. This rotor blade also provides lower manufacturing cost.

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.

Flight Test Evaluation of a Nonlinear Hub Spring on a UH-1H Helicopter.

DTIC Science & Technology

1981-04-01

APPLIED TECHNOLOGY LABORATORY POSITION STATEMENT This report documents the engineering analysis, development , arnd flight test of a non- linger hub...order to develop a design criteria to ensure that mast loads can be sustained during in-flight flapping stop contact. In addition, a comparison of the...LIST OF ILLUSTRATIONS Figure Page 1 Rotor blade-element aerodynamic coefficients used in ARHF01 .................................. 18 2 Rotor model on

Reconstruction method for running shape of rotor blade considering nonlinear stiffness and loads

NASA Astrophysics Data System (ADS)

Wang, Yongliang; Kang, Da; Zhong, Jingjun

2017-10-01

The aerodynamic and centrifugal loads acting on the rotating blade make the blade configuration deformed comparing to its shape at rest. Accurate prediction of the running blade configuration plays a significant role in examining and analyzing turbomachinery performance. Considering nonlinear stiffness and loads, a reconstruction method is presented to address transformation of a rotating blade from cold to hot state. When calculating blade deformations, the blade stiffness and load conditions are updated simultaneously as blade shape varies. The reconstruction procedure is iterated till a converged hot blade shape is obtained. This method has been employed to determine the operating blade shapes of a test rotor blade and the Stage 37 rotor blade. The calculated results are compared with the experiments. The results show that the proposed method used for blade operating shape prediction is effective. The studies also show that this method can improve precision of finite element analysis and aerodynamic performance analysis.

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.

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

NASA Technical Reports Server (NTRS)

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

1983-01-01

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

Aerodynamic Characteristics of SC1095 and SC1094 R8 Airfoils

DTIC Science & Technology

2003-12-01

Development, and Engineering Command Ames Research Center Moffett Field, California December 2003 National Aeronautics and Space Administration Ames...60A ROTOR BLADE AND AIRFOILS ................................................................................... 2 EVALUATION OF SECTION CHARACTERISTICS...Characteristics of SC1095 and SC1094 R8 Airfoils WILLIAM G. BOUSMAN Aeroflightdynamics Directorate U.S. Army Research, Development, and Engineering Command Ames

Design, fabrication, and test of a composite material wind turbine rotor blade

NASA Technical Reports Server (NTRS)

Griffee, D. G., Jr.; Gustafson, R. E.; More, E. R.

1977-01-01

The aerodynamic design, structural design, fabrication, and structural testing is described for a 60 foot long filament wound, fiberglass/epoxy resin matrix wind turbine rotor blade for a 125 foot diameter, 100 kW wind energy conversion system. One blade was fabricated which met all aerodynamic shape requirements and was structurally capable of operating under all specified design conditions. The feasibility of filament winding large rotor blades was demonstrated.

Electronic Warfare and Radar Systems Engineering Handbook

DTIC Science & Technology

1999-04-01

EWRL Electronic Warfare Reprogrammable ELNOT Emitter Library Notation Library (USN) EM Electromagnetic EWSI EW Systems Integration E-Mail Electronic...ram air turbine (RAT) propellers used to power aircraft pods, helicopter rotor blades, and protruding surfaces of automobile hubcaps will all provide

Rotating Shake Test and Modal Analysis of a Model Helicopter Rotor Blade

NASA Technical Reports Server (NTRS)

Wilkie, W. Keats; Mirick, Paul H.; Langston, Chester W.

1997-01-01

Rotating blade frequencies for a model generic helicopter rotor blade mounted on an articulated hub were experimentally determined. Testing was conducted using the Aeroelastic Rotor Experimental System (ARES) testbed in the Helicopter Hover Facility (HBF) at Langley Research Center. The measured data were compared to pretest analytical predictions of the rotating blade frequencies made using the MSC/NASTRAN finite-element computer code. The MSC/NASTRAN solution sequences used to analyze the model were modified to account for differential stiffening effects caused by the centrifugal force acting on the blade and rotating system dynamic effects. The correlation of the MSC/NASTRAN-derived frequencies with the experimental data is, in general, very good although discrepancies in the blade torsional frequency trends and magnitudes were observed. The procedures necessary to perform a rotating system modal analysis of a helicopter rotor blade with MSC/NASTRAN are outlined, and complete sample data deck listings are provided.

Numerical simulation of a hovering rotor using embedded grids

NASA Technical Reports Server (NTRS)

Duque, Earl-Peter N.; Srinivasan, Ganapathi R.

1992-01-01

The flow field for a rotor blade in hover was computed by numerically solving the compressible thin-layer Navier-Stokes equations on embedded grids. In this work, three embedded grids were used to discretize the flow field - one for the rotor blade and two to convect the rotor wake. The computations were performed at two hovering test conditions, for a two-bladed rectangular rotor of aspect ratio six. The results compare fairly with experiment and illustrates the use of embedded grids in solving helicopter type flow fields.

Rotorcraft Aeromechanics Branch Home Page on the World Wide Web

NASA Technical Reports Server (NTRS)

Peterson, Randall L.; Warmbrodt, William (Technical Monitor)

1996-01-01

The tilt rotor aircraft holds great promise for improving air travel in the future. It's benefits include vertical take off and landing combined with airspeeds comparable to propeller driven aircraft. However, the noise from a tilt rotor during approach to a landing is potentially a significant barrier to widespread acceptance of these aircraft. This approach noise is primarily caused by Blade Vortex Interactions (BVI), which are created when the blade passes near or through the vortex trailed by preceding blades. The XV- 15 Aeroacoustic test will measure the noise from a tilt rotor during descent conditions and demonstrate several possible techniques to reduce the noise. The XV- 15 Aeroacoustic test at NASA Ames Research Center will measure acoustics and performance for a full-scale XV-15 rotor. A single XV-15 rotor will be mounted on the Ames Rotor Test Apparatus (RTA) in the 80- by 120-Foot Wind Tunnel. The test will be conducted in helicopter mode with forward flight speeds up to 100 knots and tip path plane angles up to +/- 15 degrees. These operating conditions correspond to a wide range of tilt rotor descent and transition to forward flight cases. Rotor performance measurements will be made with the RTA rotor balance, while acoustic measurements will be made using an acoustic traverse and four fixed microphones. The acoustic traverse will provide limited directionality measurements on the advancing side of the rotor, where BVI noise is expected to be the highest. Baseline acoustics and performance measurements for the three-bladed rotor will be obtained over the entire test envelope. Acoustic measurements will also be obtained for correlation with the XV-15 aircraft Inflight Rotor Aeroacoustic Program (IRAP) recently conducted by Ames. Several techniques will be studied in an attempt to reduce the highest measured BVI noise conditions. The first of these techniques will use sub-wings mounted on the blade tips. These subwings are expected to alter the size, strength, and location of the tip vortex, therefore changing the BVI acoustics of the rotor. The subwings are approximately 20% of the blade chord and increase the rotor radius by about 3 percent. Four different subwing configurations will be tested, including square tipped subwings with different angles of incidence. The ability of active controls to reduce BVI acoustics will also be assessed. The dynamic control system of the RTA will be used to implement open- and closed-loop active control techniques, including individual blade control. Open-loop testing will be conducted using a personal computer based, automated, real-time data acquisition system. This system features combined automated output of open loop control signals and automated data acquisition of the resulting test data. A final technique to alter the noise of the rotor will be examined. This will involve changing the number of blades from three to four. A four-bladed rotor hub has been fabricated on which the XV-15 blades will be mounted. While the solidity of the rotor will increase, much useful information can be gained by examining the changes in the thrust and RPM with four blades.

Preliminary structural design of composite main rotor blades for minimum weight

NASA Technical Reports Server (NTRS)

Nixon, Mark W.

1987-01-01

A methodology is developed to perform minimum weight structural design for composite or metallic main rotor blades subject to aerodynamic performance, material strength, autorotation, and frequency constraints. The constraints and load cases are developed such that the final preliminary rotor design will satisfy U.S. Army military specifications, as well as take advantage of the versatility of composite materials. A minimum weight design is first developed subject to satisfying the aerodynamic performance, strength, and autorotation constraints for all static load cases. The minimum weight design is then dynamically tuned to avoid resonant frequencies occurring at the design rotor speed. With this methodology, three rotor blade designs were developed based on the geometry of the UH-60A Black Hawk titanium-spar rotor blade. The first design is of a single titanium-spar cross section, which is compared with the UH-60A Black Hawk rotor blade. The second and third designs use single and multiple graphite/epoxy-spar cross sections. These are compared with the titanium-spar design to demonstrate weight savings from use of this design methodology in conjunction with advanced composite materials.

Application of Computer Simulation to Identify Erosion Resistance of Materials of Wet-steam Turbine Blades

NASA Astrophysics Data System (ADS)

Korostelyov, D. A.; Dergachyov, K. V.

2017-10-01

A problem of identifying the efficiency of using materials, coatings, linings and solderings of wet-steam turbine rotor blades by means of computer simulation is considered. Numerical experiments to define erosion resistance of materials of wet-steam turbine blades are described. Kinetic curves for erosion area and weight of the worn rotor blade material of turbines K-300-240 LMP and atomic icebreaker “Lenin” have been defined. The conclusion about the effectiveness of using different erosion-resistant materials and protection configuration of rotor blades is also made.

Growian rotor blades: Production development, construction and test

NASA Technical Reports Server (NTRS)

Thiele, H. M.

1984-01-01

Development and construction of three 50 m rotor blades for a 3 MW wind turbine are described. A hybrid concept was chosen, i.e., a load carrying inflexible steel spar and a glass fiber reinforced plastic skin. A test blade was constructed and static loading tests, dynamic vibration tests and fatigue tests on critical welds as well as at the connection between spar and blade skin were performed. All test results show good accordance with calculated values, and were taken into consideration during the construction of two rotor blades.

Some issues on modeling atmospheric turbulence experienced by helicopter rotor blades

NASA Technical Reports Server (NTRS)

Costello, Mark; Gaonkar, G. H.; Prasad, J. V. R.; Schrage, D. P.

1992-01-01

The atmospheric turbulence velocities seen by nonrotating aircraft components and rotating blades can be substantially different. The differences are due to the spatial motion of the rotor blades, which move fore and aft through the gust waves. Body-fixed atmospheric turbulence refers to the actual atmospheric turbulence experienced by a point fixed on a nonrotating aircraft component such as the aircraft's center of gravity or the rotor hub, while blade-fixed atmospheric turbulence refers to the atmospheric turbulence experienced by an element of the rotating rotor blade. An example is presented, which, though overly simplified, shows important differences between blade- and body-fixed rotorcraft atmospheric turbulence models. All of the information necessary to develop the dynamic equations describing the atmospheric turbulence velocity field experienced by an aircraft is contained in the atmospheric turbulence velocity correlation matrix. It is for this reason that a generalized formulation of the correlation matrix describing atmospheric turbulence that a rotating blade encounters is developed. From this correlation matrix, earlier treated cases restricted to a rotor flying straight and level directly into the mean wind can be recovered as special cases.

Anisotropic piezoelectric twist actuation of helicopter rotor blades: Aeroelastic analysis and design optimization

NASA Astrophysics Data System (ADS)

Wilkie, William Keats

1997-12-01

An aeroelastic model suitable for control law and preliminary structural design of composite helicopter rotor blades incorporating embedded anisotropic piezoelectric actuator laminae is developed. The aeroelasticity model consists of a linear, nonuniform beam representation of the blade structure, including linear piezoelectric actuation terms, coupled with a nonlinear, finite-state unsteady aerodynamics model. A Galerkin procedure and numerical integration in the time domain are used to obtain a soluti An aeroelastic model suitable for control law and preliminary structural design of composite helicopter rotor blades incorporating embedded anisotropic piezoelectric actuator laminae is developed. The aeroelasticity model consists of a linear, nonuniform beam representation of the blade structure, including linear piezoelectric actuation terms, coupled with a nonlinear, finite-state unsteady aerodynamics model. A Galerkin procedure and numerical integration in the time domain are used to obtain amited additional piezoelectric material mass, it is shown that blade twist actuation approaches which exploit in-plane piezoelectric free-stain anisotropies are capable of producing amplitudes of oscillatory blade twisting sufficient for rotor vibration reduction applications. The second study examines the effectiveness of using embedded piezoelectric actuator laminae to alleviate vibratory loads due to retreating blade stall. A 10 to 15 percent improvement in dynamic stall limited forward flight speed, and a 5 percent improvement in stall limited rotor thrust were numerically demonstrated for the active twist rotor blade relative to a conventional blade design. The active twist blades are also demonstrated to be more susceptible than the conventional blades to dynamic stall induced vibratory loads when not operating with twist actuation. This is the result of designing the active twist blades with low torsional stiffness in order to maximize piezoelectric twist authority. Determining the optimum tradeoff between blade torsional stiffness and piezoelectric twist actuation authority is the subject of the third study. For this investigation, a linearized hovering-flight eigenvalue analysis is developed. Linear optimal control theory is then utilized to develop an optimum active twist blade design in terms of reducing structural energy and control effort cost. The forward flight vibratory loads characteristics of the torsional stiffness optimized active twist blade are then examined using the nonlinear, forward flight aeroelastic analysis. The optimized active twist rotor blade is shown to have improved passive and active vibratory loads characteristics relative to the baseline active twist blades.

Application of system identification to analytic rotor modeling from simulated and wind tunnel dynamic test data, part 2

NASA Technical Reports Server (NTRS)

Hohenemser, K. H.; Banerjee, D.

1977-01-01

An introduction to aircraft state and parameter identification methods is presented. A simplified form of the maximum likelihood method is selected to extract analytical aeroelastic rotor models from simulated and dynamic wind tunnel test results for accelerated cyclic pitch stirring excitation. The dynamic inflow characteristics for forward flight conditions from the blade flapping responses without direct inflow measurements were examined. The rotor blades are essentially rigid for inplane bending and for torsion within the frequency range of study, but flexible in out-of-plane bending. Reverse flow effects are considered for high rotor advance ratios. Two inflow models are studied; the first is based on an equivalent blade Lock number, the second is based on a time delayed momentum inflow. In addition to the inflow parameters, basic rotor parameters like the blade natural frequency and the actual blade Lock number are identified together with measurement bias values. The effect of the theoretical dynamic inflow on the rotor eigenvalues is evaluated.

Performance and loads data from an outdoor hover test of a Lynx tail rotor

NASA Technical Reports Server (NTRS)

Signor, David B.; Yamauchi, Gloria K.; Smith, Charles A.; Hagen, Martin J.

1989-01-01

A Lynx tail rotor was tested in hover at the Outdoor Aerodynamic Research Facility at NASA Ames Research Center. The test objectives were to measure the isolated rotor performance to provide a baseline for subsequent testing, and to operate the rotor throughout the speed and collective envelope before testing in the NFAC 40- by 80-Foot Wind Tunnel. Rotor forces and blade bending moments were measured at ambient wind conditions from zero to 6.23 m/sec. The test envelope was limited to rotor speeds of 1550 to 1850 rpm and minus 13 deg to plus 20 deg of blade collective pitch. The isolated rotor performance and blade loads data are presented.

Genetic fuzzy system for online structural health monitoring of composite helicopter rotor blades

NASA Astrophysics Data System (ADS)

Pawar, Prashant M.; Ganguli, Ranjan

2007-07-01

A structural health monitoring (SHM) methodology is developed for composite rotor blades. An aeroelastic analysis of composite rotor blades based on the finite element method in space and time and with implanted matrix cracking and debonding/delamination damage is used to obtain measurable system parameters such as blade response, loads and strains. A rotor blade with a two-cell airfoil section and [0/±45/90]s family of laminates is used for numerical simulations. The model based measurements are contaminated with noise to simulate real data. Genetic fuzzy systems (GFS) are developed for global online damage detection using displacement and force-based measurement deviations between damaged and undamaged conditions and for local online damage detection using strains. It is observed that the success rate of the GFS depends on number of measurements, type of measurements and training and testing noise level. The GFS work quite well with noisy data and is recommended for online SHM of composite helicopter rotor blades.

Process Modeling of Composite Materials for Wind-Turbine Rotor Blades: Experiments and Numerical Modeling

PubMed Central

Wieland, Birgit; Ropte, Sven

2017-01-01

The production of rotor blades for wind turbines is still a predominantly manual process. Process simulation is an adequate way of improving blade quality without a significant increase in production costs. This paper introduces a module for tolerance simulation for rotor-blade production processes. The investigation focuses on the simulation of temperature distribution for one-sided, self-heated tooling and thick laminates. Experimental data from rotor-blade production and down-scaled laboratory tests are presented. Based on influencing factors that are identified, a physical model is created and implemented as a simulation. This provides an opportunity to simulate temperature and cure-degree distribution for two-dimensional cross sections. The aim of this simulation is to support production processes. Hence, it is modelled as an in situ simulation with direct input of temperature data and real-time capability. A monolithic part of the rotor blade, the main girder, is used as an example for presenting the results. PMID:28981458

Process Modeling of Composite Materials for Wind-Turbine Rotor Blades: Experiments and Numerical Modeling.

PubMed

Wieland, Birgit; Ropte, Sven

2017-10-05

The production of rotor blades for wind turbines is still a predominantly manual process. Process simulation is an adequate way of improving blade quality without a significant increase in production costs. This paper introduces a module for tolerance simulation for rotor-blade production processes. The investigation focuses on the simulation of temperature distribution for one-sided, self-heated tooling and thick laminates. Experimental data from rotor-blade production and down-scaled laboratory tests are presented. Based on influencing factors that are identified, a physical model is created and implemented as a simulation. This provides an opportunity to simulate temperature and cure-degree distribution for two-dimensional cross sections. The aim of this simulation is to support production processes. Hence, it is modelled as an in situ simulation with direct input of temperature data and real-time capability. A monolithic part of the rotor blade, the main girder, is used as an example for presenting the results.

77 FR 33422 - Utility Scale Wind Towers From the People's Republic of China: Preliminary Affirmative...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-06-06

... support the nacelle and rotor blades in a wind turbine with a minimum rated electrical power generation... joined with non-subject merchandise, such as nacelles or rotor blades, and whether or not they have... are nacelles and rotor blades, regardless of whether they are attached to the wind tower. Also...

Design of helicopter rotor blades for optimum dynamic characteristics

NASA Technical Reports Server (NTRS)

Peters, D. A.; Ko, T.; Korn, A.; Rossow, M. P.

1985-01-01

The mass and stiffness distributions for helicopter rotor blades are tailored in such a way to give a predetermined placement of blade natural frequencies. The optimal design is pursued with respect of minimum weight, sufficient inertia, and reasonable dynamic characteristics. Finite element techniques are used as a tool. Rotor types include hingeless, articulated, and teetering.

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

NASA Technical Reports Server (NTRS)

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

1983-01-01

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

Analysis and Testing of a Composite Fuselage Shield for Open Rotor Engine Blade-Out Protection

NASA Technical Reports Server (NTRS)

Pereira, J. Michael; Emmerling, William; Seng, Silvia; Frankenberger, Charles; Ruggeri, Charles R.; Revilock, Duane M.; Carney, Kelly S.

2015-01-01

The Federal Aviation Administration is working with the European Aviation Safety Agency to determine the certification base for proposed new engines that would not have a containment structure on large commercial aircraft. Equivalent safety to the current fleet is desired by the regulators, which means that loss of a single fan blade will not cause hazard to the Aircraft. The NASA Glenn Research Center and The Naval Air Warfare Center (NAWC), China Lake, collaborated with the FAA Aircraft Catastrophic Failure Prevention Program to design and test lightweight composite shields for protection of the aircraft passengers and critical systems from a released blade that could impact the fuselage. In the test, two composite blades were pyrotechnically released from a running engine, each impacting a composite shield with a different thickness. The thinner shield was penetrated by the blade and the thicker shield prevented penetration. This was consistent with pre-test predictions. This paper documents the live fire test from the full scale rig at NAWC China Lake and describes the damage to the shields as well as instrumentation results.

Effects of rotor model degradation on the accuracy of rotorcraft real time simulation

NASA Technical Reports Server (NTRS)

Houck, J. A.; Bowles, R. L.

1976-01-01

The effects are studied of degrading a rotating blade element rotor mathematical model to meet various real-time simulation requirements of rotorcraft. Three methods of degradation were studied: reduction of number of blades, reduction of number of blade segments, and increasing the integration interval, which has the corresponding effect of increasing blade azimuthal advance angle. The three degradation methods were studied through static trim comparisons, total rotor force and moment comparisons, single blade force and moment comparisons over one complete revolution, and total vehicle dynamic response comparisons. Recommendations are made concerning model degradation which should serve as a guide for future users of this mathematical model, and in general, they are in order of minimum impact on model validity: (1) reduction of number of blade segments, (2) reduction of number of blades, and (3) increase of integration interval and azimuthal advance angle. Extreme limits are specified beyond which the rotating blade element rotor mathematical model should not be used.

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

NASA Astrophysics Data System (ADS)

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

2014-04-01

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

An advanced stochastic model for threshold crossing studies of rotor blade vibrations.

NASA Technical Reports Server (NTRS)

Gaonkar, G. H.; Hohenemser, K. H.

1972-01-01

A stochastic model to analyze turbulence-excited rotor blade vibrations, previously described by Gaonkar et al. (1971), is generalized to include nonuniformity of the atmospheric turbulence velocity across the rotor disk in the longitudinal direction. The results of the presented analysis suggest that the nonuniformity of the vertical turbulence over the rotor disk is of little influence on the random blade flapping response, at least as far as longitudinal nonuniformity is concerned.

Preliminary Investigation of a Gas Turbine with Sillimanite Ceramic Rotor Blades

DTIC Science & Technology

1947-07-01

1399 s PRELIMINARY INVESTIGATION OF A GAS TUR81NE WITH SILLIMMWTE CEIUUMIC ROTOR BIXDES By Frederick J. Hartwig, Bob W. Sheflin and Robert J. Jones...SILLIWITE CERAMIC ROTOR BLADES 13yFrederick J. Hartwig, Bob W. ShefMm and Robert J. Jones SUMMARY A gas turbine with rotor blades of a sillimanite...to rotating turbine components. Invostlgations have shown that advantageous appli- cations of ceramics to stressed components of gas turbines can be

Performance Data from a Wind-Tunnel Test of Two Main-rotor Blade Designs for a Utility-Class Helicopter

NASA Technical Reports Server (NTRS)

Singleton, Jeffrey D.; Yeager, William T., Jr.; Wilbur, Matthew L.

1990-01-01

An investigation was conducted in the NASA Langley Transonic Dynamics Tunnel to evaluate an advanced main rotor designed for use on a utility class helicopter, specifically the U.S. Army UH-60A Blackhawk. This rotor design incorporated advanced twist, airfoil cross sections, and geometric planform. For evaluation purposes, the current UH-60A main rotor was also tested and is referred to as the baseline blade set. A total of four blade sets were tested. One set of both the baseline and the advanced rotors were dynamically scaled to represent a full scale helicopter rotor blade design. The remaining advanced and baseline blade sets were not dynamically scaled so as to isolate the effects of structural elasticity. The investigation was conducted in hover and at rotor advance ratios ranging from 0.15 to 0.4 at a range of nominal test medium densities from 0.00238 to 0.009 slugs/cu ft. This range of densities, coupled with varying rotor lift and propulsive force, allowed for the simulation of several vehicle gross weight and density altitude combinations. Performance data are presented for all blade sets without analysis; however, cross referencing of data with flight condition may be useful to the analyst for validating aeroelastic theories and design methodologies as well as for evaluating advanced design parameters.

Flapping response characteristics of hingeless rotor blades by a gereralized harmonic balance method

NASA Technical Reports Server (NTRS)

Peters, D. A.; Ormiston, R. A.

1975-01-01

Linearized equations of motion for the flapping response of flexible rotor blades in forward flight are derived in terms of generalized coordinates. The equations are solved using a matrix form of the method of linear harmonic balance, yielding response derivatives for each harmonic of the blade deformations and of the hub forces and moments. Numerical results and approximate closed-form expressions for rotor derivatives are used to illustrate the relationships between rotor parameters, modeling assumptions, and rotor response characteristics. Finally, basic hingeless rotor response derivatives are presented in tabular and graphical form for a wide range of configuration parameters and operating conditions.

Digital computer program DF1758 fully coupled natural frequencies and mode shapes of a helicopter rotor blade

NASA Technical Reports Server (NTRS)

Bennett, R. L.

1975-01-01

The analytical techniques and computer program developed in the fully-coupled rotor vibration study are described. The rotor blade natural frequency and mode shape analysis was implemented in a digital computer program designated DF1758. The program computes collective, cyclic, and scissor modes for a single blade within a specified range of frequency for specified values of rotor RPM and collective angle. The analysis includes effects of blade twist, cg offset from reference axis, and shear center offset from reference axis. Coupled inplane, out-of-plane, and torsional vibrations are considered. Normalized displacements, shear forces and moments may be printed out and Calcomp plots of natural frequencies as a function of rotor RPM may be produced.

Applying Dynamic Wake Models to Induced Power Calculations for an Optimum Rotor

DTIC Science & Technology

2009-08-01

versions being special cases of the general one. Although the rotor blade may be moving at transonic speeds near the tip, the rotor wake is...The effect of a finite number of blades incurs an additional loss in wake energy due to the individual vortex sheets from each blade . In 1929... blades . Up to this point, previous developments have been able to achieve the full description of the wake in all ranges of flight regime

Transonic flow analysis for rotors. Part 1: Three-dimensional quasi-steady, full-potential calculation

NASA Technical Reports Server (NTRS)

Chang, I. C.

1984-01-01

A new computer program is presented for calculating the quasi-steady transonic flow past a helicopter rotor blade in hover as well as in forward flight. The program is based on the full potential equations in a blade attached frame of reference and is capable of treating a very general class of rotor blade geometries. Computed results show good agreement with available experimental data for both straight and swept tip blade geometries.

Remote-Controlled Rotorcraft Blade Vibration and Modal Analysis at Low Frequencies

DTIC Science & Technology

2016-02-01

modal analysis, remote-controlled helicopter , remote-controlled rotorcraft, HUMS for rotorcraft 16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF...Experimental Setup 1 4. Results 4 4.1 Rotor Blade Acceleration 4 4.2 Modal Analysis: Using an Impact Hammer 7 4.3 Dynamic Response Revisited 8 5... Rotor blade response to shaker outputting 1-V sine wave at 100 Hz ....5 Fig. 6 Rotor blade response to shaker outputting 1-V sine sweep from 20- to 100

Design of helicopter rotor blades with actuators made of a piezomacrofiber composite

NASA Astrophysics Data System (ADS)

Glukhikh, S.; Barkanov, E.; Kovalev, A.; Masarati, P.; Morandini, M.; Riemenschneider, J.; Wierach, P.

2008-01-01

For reducing the vibration and noise of helicopter rotor blades, the method of their controlled twisting by using built-in deformation actuators is employed. In this paper, the influence of various design parameters of the blades, including the location of actuators made of a piezomacrofiber material, on the twist angle is evaluated. The results of a parametric analysis performed allowed us to refine the statement of an optimization problem for the rotor blades.

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.

Aerodynamic design of a rotor blade for minimum noise radiation

NASA Technical Reports Server (NTRS)

Karamcheti, K.; Yu, Y. H.

1974-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1972-01-01

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

Comprehensive helicopter rotor instrumentation - A retrofit approach using miniature transducers

NASA Technical Reports Server (NTRS)

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

1992-01-01

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

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.

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

NASA Technical Reports Server (NTRS)

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

1997-01-01

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

Apparatus and method for ultrasonic reconstruction and testing of a turbine rotor blade attachment structure

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

Sabourin, P.F.

1995-04-25

An apparatus and method for ultrasonic reconstruction and testing of a non-visible turbine rotor blade attachment structure is described. The method of the invention includes positioning transducers at a first location to obtain slot region scan data corresponding to a slot region of the non-visible turbine rotor blade attachment structure, and positioning transducers at a second location to obtain straddle-mount region scan data corresponding to a straddle-mount region of the non-visible turbine rotor blade attachment structure. The shape of the non-visible turbine rotor blade attachment structure is reconstructed from the slot region scan data and the straddle-mount region scan datamore » to form reconstruction data. The reconstruction data is used to select test scan positions for ultrasonic testing. Ultrasonic testing is then performed at the selected test scan positions. 11 figs.« less

Permeability and Strength Measurements on Sintered, Porous, Hollow Turbine Blades Made by the American Electro Metal Corporation under Office of Naval Research Contract N-ONR-295 (01)

NASA Technical Reports Server (NTRS)

Richards, Hadley T.; Livingood, N.B.

1954-01-01

An experimental investigation was made to determine the permeability and strength characteristics of a number of sintered, porous, hollow turbine rotor blades and to determine the effectiveness of the blade fabrication method on permeability control. The test blades were fabricated by the American Electro Metal Corporation under a contract with the Office of Naval Research, Department of the Navy, and were submitted to the NACA for testing. Of the 22 test blades submitted, ten were sintered but not coined, five were sintered and coined, and seven were sintered and not coined but contained perforated reinforcements integral with the blade shells. Representative samples of each group of blades were tested. Large variations in permeability in both chordwise and spanwise directions were found. Local deviations as large as 155 to -85 percent from prescribed values were found in chordwise permeability. Only one blade, an uncoined one, had a chordwise permeability variations which reasonably approached that specified. Even for this blade, local deviations exceeded 10 percent. Spanwise permeability, specified to be held constant, varied as much as 50 percent from root to tip for both an uncoined and a coined blade. Previous NACA analyses have shown that in order to maintain proper control of blade wall temperatures, permeability variations must not exceed plus or minus 10 percent. Satisfactory control of permeability in either the chordwise or the spanwise direction was not achieved in the blades tested. Spin tests made at room temperature for six blades revealed the highest material rupture strength to be 8926 pounds per square inch. This value is about one third the strength required for rotor blades in present-day turbojet engines. The lowest value of blade strength was 1436 pounds per square inch.

Fan Noise Source Diagnostic Test: LDV Measured Flow Field Results

NASA Technical Reports Server (NTRS)

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

2003-01-01

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

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.

Lightweight engine containment. [Kevlar shielding

NASA Technical Reports Server (NTRS)

Weaver, A. T.

1977-01-01

Kevlar fabric styles and weaves were studied, as well as methods of application for advanced gas turbine engines. The Kevlar material was subjected to high speed impacts by simple projectiles fired from a rifle, as well as more complex shapes such as fan blades released from gas turbine rotors in a spin pit. Just contained data was developed for a variety of weave and/or application techniques, and a comparative containment weight efficiency was established for Kevlar containment applications. The data generated during these tests is being incorporated into an analytical design system so that blade containment trade-off studies between Kevlar and metal case engine structures can be made. Laboratory tests and engine environment tests were performed to determine the survivability of Kevlar in a gas turbine environment.

Investigation of Maximum Blade Loading Capability of Lift-Offset Rotors

NASA Technical Reports Server (NTRS)

Yeo, Hyeonsoo; Johnson, Wayne

2013-01-01

Maximum blade loading capability of a coaxial, lift-offset rotor is investigated using a rotorcraft configuration designed in the context of short-haul, medium-size civil and military missions. The aircraft was sized for a 6600-lb payload and a range of 300 nm. The rotor planform and twist were optimized for hover and cruise performance. For the present rotor performance calculations, the collective pitch angle is progressively increased up to and through stall with the shaft angle set to zero. The effects of lift offset on rotor lift, power, controls, and blade airloads and structural loads are examined. The maximum lift capability of the coaxial rotor increases as lift offset increases and extends well beyond the McHugh lift boundary as the lift potential of the advancing blades are fully realized. A parametric study is conducted to examine the differences between the present coaxial rotor and the McHugh rotor in terms of maximum lift capabilities and to identify important design parameters that define the maximum lift capability of the rotor. The effects of lift offset on rotor blade airloads and structural loads are also investigated. Flap bending moment increases substantially as lift offset increases to carry the hub roll moment even at low collective values. The magnitude of flap bending moment is dictated by the lift-offset value (hub roll moment) but is less sensitive to collective and speed.

Integrated Turbine Tip Clearance and Gas Turbine Engine Simulation

NASA Technical Reports Server (NTRS)

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

2016-01-01

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

Subsonic/transonic stall flutter investigation of a rotating rig

NASA Technical Reports Server (NTRS)

Jutras, R. R.; Fost, R. B.; Chi, R. M.; Beacher, B. F.

1981-01-01

Stall flutter is investigated by obtaining detailed quantitative steady and aerodynamic and aeromechanical measurements in a typical fan rotor. The experimental investigation is made with a 31.3 percent scale model of the Quiet Engine Program Fan C rotor system. Both subsonic/transonic (torsional mode) flutter and supersonic (flexural) flutter are investigated. Extensive steady and unsteady data on the blade deformations and aerodynamic properties surrounding the rotor are acquired while operating in both the steady and flutter modes. Analysis of this data shows that while there may be more than one traveling wave present during flutter, they are all forward traveling waves.

Numerical Study of Stratified Charge Combustion in Wave Rotors

NASA Technical Reports Server (NTRS)

Nalim, M. Razi

1997-01-01

A wave rotor may be used as a pressure-gain combustor effecting non-steady flow, and intermittent, confined combustion to enhance gas turbine engine performance. It will be more compact and probably lighter than an equivalent pressure-exchange wave rotor, yet will have similar thermodynamic and mechanical characteristics. Because the allowable turbine blade temperature limits overall fuel/air ratio to sub-flammable values, premixed stratification techniques are necessary to burn hydrocarbon fuels in small engines with compressor discharge temperature well below autoignition conditions. One-dimensional, unsteady numerical simulations of stratified-charge combustion are performed using an eddy-diffusivity turbulence model and a simple reaction model incorporating a flammability limit temperature. For good combustion efficiency, a stratification strategy is developed which concentrates fuel at the leading and trailing edges of the inlet port. Rotor and exhaust temperature profiles and performance predictions are presented at three representative operating conditions of the engine: full design load, 40% load, and idle. The results indicate that peak local gas temperatures will result in excessive temperatures within the rotor housing unless additional cooling methods are used. The rotor itself will have acceptable temperatures, but the pattern factor presented to the turbine may be of concern, depending on exhaust duct design and duct-rotor interaction.

Development of Virtual Blade Model for Modelling Helicopter Rotor Downwash in OpenFOAM

DTIC Science & Technology

2013-12-01

UNCLASSIFIED Development of Virtual Blade Model for Modelling Helicopter Rotor Downwash in OpenFOAM Stefano Wahono Aerospace...Georgia Institute of Technology. The OpenFOAM predicted result was also shown to compare favourably with ANSYS Fluent predictions. RELEASE...UNCLASSIFIED Development of Virtual Blade Model for Modelling Helicopter Rotor Downwash in OpenFOAM Executive Summary The Infrared

Adhesive quality inspection of wind rotor blades using thermography

NASA Astrophysics Data System (ADS)

Li, Xiaoli; Sun, Jiangang; Shen, Jingling; Wang, Xun; Zhang, Cunlin; Zhao, Yuejin

2018-04-01

Wind power is playing an increasingly important role in ensuring electrical safety for human beings. Because wind rotor blades are getting larger and larger in order to harvest wind energy more efficiently, there is a growing demand for nondestructive testing. Due to the glue structure of rotor blades, adhesive quality evaluation is needed. In this study, three adhesive samples with a wall thickness of 13mm, 28mm or 31mm were each designed with a different adhesive situation. The transmission thermography was applied to inspect the samples. The results illustrate that this method is effective to inspect adhesive quality of wind rotor blades.

Fundamental Understanding of Rotor Aeromechanics at High Advance Ratio Through Wind Tunnel Testing

NASA Astrophysics Data System (ADS)

Berry, Benjamin

The purpose of this research is to further the understanding of rotor aeromechanics at advance ratios (mu) beyond the maximum of 0.5 (ratio of forward airspeed to rotor tip speed) for conventional helicopters. High advance ratio rotors have applications in high speed compound helicopters. In addition to one or more conventional main rotors, these aircraft employ either thrust compounding (propellers), lift compounding (fixed-wings), or both. An articulated 4-bladed model rotor was constructed, instrumented, and tested up to a maximum advance ratio of mu=1.6 in the Glenn L. Martin Wind Tunnel at the University of Maryland. The data set includes steady and unsteady rotor hub forces and moments, blade structural loads, blade flapping angles, swashplate control angles, and unsteady blade pressures. A collective-thrust control reversal--where increasing collective pitch results in lower rotor thrust--was observed and is a unique phenomenon to the high advance ratio flight regime. The thrust reversal is explained in a physical manner as well as through an analytical formulation. The requirements for the occurrence of the thrust reversal are enumerated. The effects of rotor geometry design on the thrust reversal onset are explored through the formulation and compared to the measured data. Reverse-flow dynamic stall was observed to extend the the lifting capability of the edgewise rotor well beyond the expected static stall behavior of the airfoil sections. Through embedded unsteady blade surface pressure transducers, the normal force, pitching moment, and shed dynamic stall vortex time histories at a blade section in strong reverse flow were analyzed. Favorable comparisons with published 2-D pitching airfoil reverse flow dynamic stall data indicate that the 3-D stall environment can likely be predicted using models developed from such 2-D experiments. Vibratory hub loads were observed to increase with advance ratio. Maximum amplitude was observed near mu=1, with a reduction in vibratory loads at higher advance ratios. Blade load 4/rev harmonics dominated due to operation near a 4/rev fanplot crossing of the 2nd flap bending mode natural frequency. Oscillatory loads sharply increase in the presence of retreating blade reverse flow dynamic stall, and are evident in blade torsion, pitch link, and hub load measurements. The blades exhibited torsion moment vibrations at the frequency of the 1st torsion mode in response to the reverse flow pitching moment loading.

Effect of helicopter blade dynamics on blade aerodynamic and structural loads

NASA Technical Reports Server (NTRS)

Heffernan, Ruth M.

1987-01-01

The effect of rotor blade dynamics on aerodynamic and structural loads is examined for a conventional, main-rotor helicopter using a comprehensive rotorcraft analysis (CAMRAD) and flight-test data. The impact of blade dynamics on blade section lift-coefficient time histories is studied by comparing predictions from a rigid-blade analysis and an elastic-blade analysis with helicopter flight test data. The elastic blade analysis better predicts high-frequency behavior of section lift. In addition, components of the blade angle of attack such as elastic blade twist, blade flap rate, blade slope velocity, and inflow are examined as a function of blade mode. Elastic blade motion changed blade angle of attack by a few tenths of a degree, and up to the sixth rotor harmonic. A similar study of the influence of blade dynamics on bending and torsion moments was also conducted. A correlation study comparing predictions from several elastic-blade analyses with flight-test data revealed that an elastic-blade model consisting of only three elastic bending modes (first and second flap and first lag), and two elastic torsion modes was sufficient for good correlation.

Summary of design and blade-element performance data for 12 axial-flow pump rotor configurations

NASA Technical Reports Server (NTRS)

Miller, M. J.; Okiishi, T. H.; Serovy, G. K.; Sandercock, D. M.; Britsch, W. R.

1973-01-01

A collection of noncavitating blade-element performance data for 12 axial-flow pump rotor configurations is presented in tabular form. Rotor design philosophy, test apparatus and procedure, and data reduction and evaluation are discussed. A data storage and recall computer program is described. All but one of the rotor configurations considered were composed of double-circular-arc blade sections and were designed for high inlet relative flow angles. Hub-tip radius ranged from 0.40 to 0.90.

Aerodynamics of small-scale vertical-axis wind turbines

NASA Astrophysics Data System (ADS)

Paraschivoiu, I.; Desy, P.

1985-12-01

The purpose of this work is to study the influence of various rotor parameters on the aerodynamic performance of a small-scale Darrieus wind turbine. To do this, a straight-bladed Darrieus rotor is calculated by using the double-multiple-streamtube model including the streamtube expansion effects through the rotor (CARDAAX computer code) and the dynamicstall effects. The straight-bladed Darrieus turbine is as expected more efficient with respect the curved-bladed rotor but for a given solidity is operates at higher wind speeds.

Helicopter Maritime Environment Trainer: Maintenance Manual (Simulateur D’Entrainement Virtuel pour Helicoptere Maritime: Manual D’Entretien)

DTIC Science & Technology

2011-06-01

rotor blades. This increases or decreases the angle of attack of all the blades simultaneously and, consequently, the tilt or vertical thrust...is the primary horizontal control for the main rotor. Directional control is accomplished by tilting the main rotor that produces a directional...thrust in that direction. The rotor is tilted by changing the pitch of each blade individually as it makes a complete rotation. The cyclic pitch change

Performance of Axial-Flow Supersonic Compressor of XJ55-FF-1 Turbojet Engine. III; Over-All Performance of Compressor

NASA Technical Reports Server (NTRS)

Hartmann, Melvin J.; Tysl, Edward R.

1949-01-01

An investigation was conducted to determine the performance characteristics of the rotor and inlet guide vanes used in the axial-flow supersonic compressor of the XJ55-FF-1 turbojet engine. Outlet stators used in the engine were omitted to facilitate study of the supersonic rotor. The extent of the deviation from design performance indicates that the design-shock configuration was not obtained. A maximum pressure ratio of 2.26 was obtained at an equivalent tip speed of 1614 feet per second and an adiabatic efficiency of 0.61. The maximum efficiency obtained was 0.79 at an equivalent tip speed of 801 feet per second and a pressure ratio of 1.29. The performance obtained was considerably below design performance. The effective aerodynamic forces encountered appeared to be large enough to cause considerable damage to the thin aluminum leading edges of the rotor blades.

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.

The prediction of rotor rotational noise using measured fluctuating blade loads

NASA Technical Reports Server (NTRS)

Hosier, R. N.; Pegg, R. J.; Ramakrishnan, R.

1974-01-01

In tests conducted at the NASA Langley Research Center Helicopter Rotor Test Facility, simultaneous measurements of the high-frequency fluctuating aerodynamic blade loads and far-field radiated noise were made on a full-scale, nontranslating rotor system. After their characteristics were determined, the measured blade loads were used in an existing theory to predict the far-field rotational noise. A comparison of the calculated and measured rotational noise is presented with specific attention given to the effect of blade loading coefficients, chordwise loading distributions, blade loading phases, and observer azimuthal position on the predictions.

Computer program for design of two-dimensional supersonic turbine rotor blades with boundary-layer correction

NASA Technical Reports Server (NTRS)

Goldman, L. J.; Scullin, V. J.

1971-01-01

A FORTRAN 4 computer program for the design of two-dimensional supersonic rotor blade sections corrected for boundary-layer displacement thickness is presented. The ideal rotor is designed by the method of characteristics to produce vortex flow within the blade passage. The boundary-layer parameters are calculated by Cohen and Reshotoko's method for laminar flow and Sasman and Cresci's method for turbulent flow. The program input consists essentially of the blade surface Mach number distribution and total flow conditions. The primary output is the corrected blade profile and the boundary-layer parameters.

Probabilistic analysis of bladed turbine disks and the effect of mistuning

NASA Technical Reports Server (NTRS)

Shah, A. R.; Nagpal, V. K.; Chamis, Christos C.

1990-01-01

Probabilistic assessment of the maximum blade response on a mistuned rotor disk is performed using the computer code NESSUS. The uncertainties in natural frequency, excitation frequency, amplitude of excitation and damping are included to obtain the cumulative distribution function (CDF) of blade responses. Advanced mean value first order analysis is used to compute CDF. The sensitivities of different random variables are identified. Effect of the number of blades on a rotor on mistuning is evaluated. It is shown that the uncertainties associated with the forcing function parameters have significant effect on the response distribution of the bladed rotor.

Probabilistic analysis of bladed turbine disks and the effect of mistuning

NASA Technical Reports Server (NTRS)

Shah, Ashwin; Nagpal, V. K.; Chamis, C. C.

1990-01-01

Probabilistic assessment of the maximum blade response on a mistuned rotor disk is performed using the computer code NESSUS. The uncertainties in natural frequency, excitation frequency, amplitude of excitation and damping have been included to obtain the cumulative distribution function (CDF) of blade responses. Advanced mean value first order analysis is used to compute CDF. The sensitivities of different random variables are identified. Effect of the number of blades on a rotor on mistuning is evaluated. It is shown that the uncertainties associated with the forcing function parameters have significant effect on the response distribution of the bladed rotor.

Effectiveness enhancement of a cycloidal wind turbine by individual active control of blade motion

NASA Astrophysics Data System (ADS)

Hwang, In Seong; Lee, Yun Han; Kim, Seung Jo

2007-04-01

In this paper, a research for the effectiveness enhancement of a Cycloidal Wind Turbine by individual active control of blade motion is described. To improve the performance of the power generation system, which consists of several straight blades rotating about axis in parallel direction, the cycloidal blade system and the individual active blade control method are adopted. It has advantages comparing with horizontal axis wind turbine or conventional vertical axis wind turbine because it maintains optimal blade pitch angles according to wind speed, wind direction and rotor rotating speed to produce high electric power at any conditions. It can do self-starting and shows good efficiency at low wind speed and complex wind condition. Optimal blade pitch angle paths are obtained through CFD analysis according to rotor rotating speed and wind speed. The individual rotor blade control system consists of sensors, actuators and microcontroller. To realize the actuating device, servo motors are installed to each rotor blade. Actuating speed and actuating force are calculated to compare with the capacities of servo motor, and some delays of blade pitch angles are corrected experimentally. Performance experiment is carried out by the wind blowing equipment and Labview system, and the rotor rotates from 50 to 100 rpm according to the electric load. From this research, it is concluded that developing new vertical axis wind turbine, Cycloidal Wind Turbine which is adopting individual active blade pitch control method can be a good model for small wind turbine in urban environment.

Comparison of Measured Flapwise Structural Bending Moments on a Teetering Rotor Blade With Results Calculated From the Measured Pressure Distribution

NASA Technical Reports Server (NTRS)

Mayo, Alton P.

1959-01-01

Flapwise bending moments were calculated for a teetering rotor blade using a reasonably rapid theoretical method in which airloads obtained from wind-tunnel tests were employed. The calculated moments agreed reasonably well with those measured with strain gages under the same test conditions. The range of the tests included one hovering and two forward-flight conditions. The rotor speed for the test was very near blade resonance, and difficult-to-calculate resonance effects apparently were responsible for the largest differences between the calculated and measured harmonic components of blade bending moments. These differences, moreover, were largely nullified when the harmonic components were combined to give a comparison of the calculated and measured blade total- moment time histories. The degree of agreement shown is therefore considered adequate to warrant the use of the theoretical method in establishing and applying methods of prediction of rotor-blade fatigue loads. At the same time, the validity of the experimental methods of obtaining both airload and blade stress measurement is also indicated to be adequate for use in establishing improved methods for prediction of rotor-blade fatigue loads during the design stage. The blade stiffnesses and natural frequencies were measured and found to be in close agreement with calculated values; however, for a condition of blade resonance the use of the experimental stiffness values resulted in better agreement between calculated and measured blade stresses.

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

NASA Technical Reports Server (NTRS)

Nallasamy, M.; Podboy, Gary G.

1994-01-01

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

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.

Cooling system with compressor bleed and ambient air for gas turbine engine

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

Marsh, Jan H.; Marra, John J.

A cooling system for a turbine engine for directing cooling fluids from a compressor to a turbine blade cooling fluid supply and from an ambient air source to the turbine blade cooling fluid supply to supply cooling fluids to one or more airfoils of a rotor assembly is disclosed. The cooling system may include a compressor bleed conduit extending from a compressor to the turbine blade cooling fluid supply that provides cooling fluid to at least one turbine blade. The compressor bleed conduit may include an upstream section and a downstream section whereby the upstream section exhausts compressed bleed airmore » through an outlet into the downstream section through which ambient air passes. The outlet of the upstream section may be generally aligned with a flow of ambient air flowing in the downstream section. As such, the compressed air increases the flow of ambient air to the turbine blade cooling fluid supply.« less

Pre-design study for a modern four-bladed rotor for the Rotor System Research Aircraft (RSRA). [integrating the YAH-64 main rotor

NASA Technical Reports Server (NTRS)

Hughes, C. W.; Logan, A. H.

1981-01-01

Various candidate rotor systems were compared in an effort to select a modern four-bladed rotor for the RSRA. The YAH-64 rotor system was chosen as the candidate rotor system for further development for the RSRA. The process used to select the rotor system, studies conducted to mate the rotor with the RSRA and provide parametric variability, and the development plan which would be used to implement these studies are presented. Drawings are included.

Stability of large horizontal-axis axisymmetric wind turbines. Ph.D. Thesis - Delaware Univ.

NASA Technical Reports Server (NTRS)

Hirschbein, M. S.; Young, M. I.

1980-01-01

The stability of large horizontal axis, axi-symmetric, power producing wind turbines was examined. The analytical model used included the dynamic coupling of the rotor, tower and power generating system. The aerodynamic loading was derived from blade element theory. Each rotor blade was permitted tow principal elastic bending degrees of freedom, one degree of freedom in torsion and controlled pitch as a rigid body. The rotor hub was mounted in a rigid nacelle which may yaw freely or in a controlled manner. The tower can bend in two principal directions and may twist. Also, the rotor speed can vary and may induce perturbation reactions within the power generating equipment. Stability was determined by the eigenvalues of a set of linearized constant coefficient differential equations. All results presented are based on a 3 bladed, 300 ft. diameter, 2.5 megawatt wind turbine. Some of the parameters varied were; wind speed, rotor speed structural stiffness and damping, the effective stiffness and damping of the power generating system and the principal bending directions of the rotor blades. Unstable or weakly stable behavior can be caused by aerodynamic forces due to motion of the rotor blades and tower in the plane of rotation or by mechanical coupling between the rotor system and the tower.

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

NASA Technical Reports Server (NTRS)

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

1974-01-01

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

DESIGN OF TWO-DIMENSIONAL SUPERSONIC TURBINE ROTOR BLADES WITH BOUNDARY-LAYER CORRECTION

NASA Technical Reports Server (NTRS)

Goldman, L. J.

1994-01-01

A computer program has been developed for the design of supersonic rotor blades where losses are accounted for by correcting the ideal blade geometry for boundary layer displacement thickness. The ideal blade passage is designed by the method of characteristics and is based on establishing vortex flow within the passage. Boundary-layer parameters (displacement and momentum thicknesses) are calculated for the ideal passage, and the final blade geometry is obtained by adding the displacement thicknesses to the ideal nozzle coordinates. The boundary-layer parameters are also used to calculate the aftermixing conditions downstream of the rotor blades assuming the flow mixes to a uniform state. The computer program input consists essentially of the rotor inlet and outlet Mach numbers, upper- and lower-surface Mach numbers, inlet flow angle, specific heat ratio, and total flow conditions. The program gas properties are set up for air. Additional gases require changes to be made to the program. The computer output consists of the corrected rotor blade coordinates, the principal boundary-layer parameters, and the aftermixing conditions. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 7094. This program was developed in 1971.

Integrated aerodynamic/dynamic optimization of helicopter rotor blades

NASA Technical Reports Server (NTRS)

Chattopadhyay, Aditi; Walsh, Joanne L.; Riley, Michael F.

1989-01-01

An integrated aerodynamic/dynamic optimization procedure is used to minimize blade weight and 4 per rev vertical hub shear for a rotor blade in forward flight. The coupling of aerodynamics and dynamics is accomplished through the inclusion of airloads which vary with the design variables during the optimization process. Both single and multiple objective functions are used in the optimization formulation. The Global Criteria Approach is used to formulate the multiple objective optimization and results are compared with those obtained by using single objective function formulations. Constraints are imposed on natural frequencies, autorotational inertia, and centrifugal stress. The program CAMRAD is used for the blade aerodynamic and dynamic analyses, and the program CONMIN is used for the optimization. Since the spanwise and the azimuthal variations of loading are responsible for most rotor vibration and noise, the vertical airload distributions on the blade, before and after optimization, are compared. The total power required by the rotor to produce the same amount of thrust for a given area is also calculated before and after optimization. Results indicate that integrated optimization can significantly reduce the blade weight, the hub shear and the amplitude of the vertical airload distributions on the blade and the total power required by the rotor.

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.

Effect of blade planform variation on the forward-flight performance of small-scale rotors

NASA Technical Reports Server (NTRS)

Noonan, Kevin W.; Althoff, Susan L.; Samak, Dhananjay K.; Green, Michael D.

1992-01-01

An investigation was conducted in the Glenn L. Martin Wind Tunnel to determine the effect of blade planform variation on the forward-flight performance of four small-scale rotors. The rotors were 5.417 ft in diameter and differed only in blade planform geometry. The four planforms were: (1) rectangular; (2) 3:1 linear taper starting at 94 percent radius; (3) 3:1 linear taper starting at 75 percent radius; and (4) 3:1 linear taper starting at 50 percent radius. Each planform had a thrust-weighted solidity of 0.098. The investigation included forward-flight simulation at advance ratios from 0.14 to 0.43 for a range of rotor lift and drag coefficients. Among the four rotors, the rectangular rotor required the highest torque for the entire range of rotor drag coefficients attained at advanced ratios greater than 0.14 for rotor lift coefficients C sub L from 0.004 to 0.007. Among the rotors with tapered blades and for C sub L = 0.004 to 0.007, either the 75 percent tapered rotor or the 50 percent tapered rotor required the least amount of torque for the full range of rotor drag coefficients attained at each advance ratio. The performance of the 94 percent tapered rotor was generally between that of the rectangular rotor and the 75 and 50 percent tapered rotors at each advance ratio for this range of rotor lift coefficients.

Vibration and loads in hingeless rotors. Volume 1: Theoretical analyses

NASA Technical Reports Server (NTRS)

Watts, G. A.; London, R. J.

1972-01-01

Analytic methods are developed for calculating blade loads and shaft-transmitted vibratory forces in stiff bladed hingeless rotors operating at advance ratios from mu = .3 to mu = 2.0. Calculated shaft harmonic moments compared well with experimental values when the blade first flap frequency was in the region of two-per-revolution harmonic excitation. Calculated blade bending moment azimuthal distributions due to changes in cyclic pitch agreed well with experiment at radial stations near the blade root at values of the ratio of first flap frequency to rotor rotation rate from 1.5 to 5.0. At stations near the blade tip good agreement was only obtained at the higher values of frequency ratio.

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

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

NASA Astrophysics Data System (ADS)

Lee, Daniel H.

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

A rotor-mounted digital instrumentation system for helicopter blade flight research measurements

NASA Technical Reports Server (NTRS)

Knight, V. H., Jr.; Haywood, W. S., Jr.; Williams, M. L.

1978-01-01

A rotor mounted flight instrumentation system developed for helicopter rotor blade research is described. The system utilizes high speed digital techniques to acquire research data from miniature pressure transducers on advanced rotor airfoils which are flight tested on an AH-1G helicopter. The system employs microelectronic pulse code modulation (PCM) multiplexer digitizer stations located remotely on the blade and in a hub mounted metal canister. As many as 25 sensors can be remotely digitized by a 2.5 mm thick electronics package mounted on the blade near the tip to reduce blade wiring. The electronics contained in the canister digitizes up to 16 sensors, formats these data with serial PCM data from the remote stations, and transmits the data from the canister which is above the plane of the rotor. Data are transmitted over an RF link to the ground for real time monitoring and to the helicopter fuselage for tape recording. The complete system is powered by batteries located in the canister and requires no slip rings on the rotor shaft.

Aeroelastic analysis for helicopter rotors with blade appended pendulum vibration absorbers. Mathematical derivations and program user's manual

NASA Technical Reports Server (NTRS)

Bielawa, R. L.

1982-01-01

Mathematical development is presented for the expanded capabilities of the United Technologies Research Center (UTRC) G400 Rotor Aeroelastic Analysis. This expanded analysis, G400PA, simulates the dynamics of teetered rotors, blade pendulum vibration absorbers and the higher harmonic excitations resulting from prescribed vibratory hub motions and higher harmonic blade pitch control. Formulations are also presented for calculating the rotor impedance matrix appropriate to these higher harmonic blade excitations. This impedance matrix and the associated vibratory hub loads are intended as the rotor blade characteristics elements for use in the Simplified Coupled Rotor/Fuselage Vibration Analysis (SIMVIB). Sections are included presenting updates to the development of the original G400 theory, and material appropriate to the user of the G400PA computer program. This material includes: (1) a general descriptionof the tructuring of the G400PA FORTRAN coding, (2) a detaild description of the required input data and other useful information for successfully running the program, and (3) a detailed description of the output results.

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.

Flight Investigation of Effects of Transition, Landing Approaches, Partial-Power Vertical Descents, and Droop-Stop Pounding on the Bending and Torsional Moments Encountered by a Helicopter Rotor Blade

NASA Technical Reports Server (NTRS)

Ludi, LeRoy H.

1959-01-01

Flight tests have been conducted with a single-rotor helicopter, one blade of which was equipped at 14 percent and 40 percent of the blade radius with strain gages calibrated to measure moments rather than stresses, to determine the effects of transition, landing approaches, and partial-power vertical descents on the rotor-blade bending and torsional moments. In addition, ground tests were conducted to determine the effects of static droop-stop pounding on the rotor-blade moments. The results indicate that partial-power vertical descents and landing approaches produce rotor-blade moments that are higher than the moments encountered in any other flight condition investigated to date with this equipment. Decelerating through the transition region in level flight was found to result in higher vibratory moments than accelerating through this region. Deliberately induced static droop-stop pounding produced flapwise bending moments at the 14-percent-radius station which were as high as the moments experienced in landing approaches and partial-power vertical descents.

Piezoelectric actuation of helicopter rotor blades

NASA Astrophysics Data System (ADS)

Lieven, Nicholas A. J.

2001-07-01

The work presented in this paper is concerned with the application of embedded piezo-electric actuators in model helicopter rotor blades. The paper outlines techniques to define the optimal location of actuators to excite particular modes of vibration whilst the blade is rotating. Using composite blades the distribution of strain energy is defined using a Finite Element model with imposed rotor-dynamic and aerodynamics loads. The loads are specified through strip theory to determine the position of maximum bending moment and thus the optimal location of the embedded actuators. The effectiveness of the technique is demonstrated on a 1/4 scale fixed cyclic pitch rotor head. Measurement of the blade displacement is achieved by using strain gauges. In addition a redundant piezo-electric actuator is used to measure the blades' response characteristics. The addition of piezo-electric devices in this application has been shown to exhibit adverse aeroelastic effects, such as counter mass balancing and increased drag. Methods to minimise these effects are suggested. The outcome of the paper is a method for defining the location and orientation of piezo-electric devices in rotor-dynamic applications.

Results of the flight noise measurement program using a standard and modified SH-3A helicopter

NASA Technical Reports Server (NTRS)

Pegg, R. J.; Henderson, H. R.; Hilton, D. A.

1973-01-01

A field noise measurement program has been conducted using both a standard SH-3A helicopter and an SH-3A helicopter modified to reduce external noise levels. Modifications included reducing rotor speed, increasing the number of rotor blades, modifying the blade-tip shapes, and acoustically treating the engine air intakes and exhaust. The purpose of this study was to document the noise characteristics recorded on the ground of each helicopter during flyby, hover, landing, and take-off operations. Based on an analysis of the measured results, the average of the overhead, overall, ontrack noise levels was approximately 4 db lower for the modified helicopter than for the standard helicopter. The improved in-flight noise characteristics, and associated small footprint areas and time durations, were judged to be mainly due to tail-rotor noise reductions. The noise reductions were obtained at the expense of required power increases at airspeeds greater than 70 knots for the modified helicopter.

Component Performance Investigation of J71 Experimental Turbine I : Over-all Performance with 97-percent-design Stator Areas

NASA Technical Reports Server (NTRS)

Schum, Harold J; Davison, Elmer H

1956-01-01

The over-all component performance characteristics of a J71 experimental three-stage turbine with 97 percent design stator areas were determined over a range of speed and pressure ratio at inlet-air conditions of approximately 35 inches of mercury absolute and 700 degrees R. The turbine break internal efficiency at design operating conditions was 0.877; the maximum efficiency of 0.886 occurred at a pressure ratio of 4.0 at 120 percent of design equivalent rotor speed. In general, the turbine yielded a wide range of efficient operation, permitting flexibility in the choice of different modes of engine operation. Limiting blade loading of the third rotor was approached but not obtained over the range of conditions investigated herein. At the design operating point, the turbine equivalent weight flow was approximately 105 percent of design. Choking of the third-rotor blades occurred at design speed and an over-all pressure ratio of 4.2.

Documentation of Helicopter Aeroelastic Stability Analysis Computer Program (HASTA)

DTIC Science & Technology

1977-12-01

of the blade phasing assumption for which all blades of the rotor are identical and equally spaced azimuthally allows the size of the T. matrices...to be significantly reduced by the removal of the submatrices associated with blades other than the first blade. With the use of this assumption ...different program representational options such as the type of rotor system, the type of blades, and the use of the blade phasing assumption , the

Efficiency of operation of wind turbine rotors optimized by the Glauert and Betz methods

NASA Astrophysics Data System (ADS)

Okulov, V. L.; Mikkelsen, R.; Litvinov, I. V.; Naumov, I. V.

2015-11-01

The models of two types of rotors with blades constructed using different optimization methods are compared experimentally. In the first case, the Glauert optimization by the pulsed method is used, which is applied independently for each individual blade cross section. This method remains the main approach in designing rotors of various duties. The construction of the other rotor is based on the Betz idea about optimization of rotors by determining a special distribution of circulation over the blade, which ensures the helical structure of the wake behind the rotor. It is established for the first time as a result of direct experimental comparison that the rotor constructed using the Betz method makes it possible to extract more kinetic energy from the homogeneous incoming flow.

Tests of Full-Scale Helicopter Rotors at High Advancing Tip Mach Numbers and Advance Ratios

NASA Technical Reports Server (NTRS)

Biggers, James C.; McCloud, John L., III; Stroub, Robert H.

2015-01-01

As a continuation of the studies of reference 1, three full-scale helicopter rotors have been tested in the Ames Research Center 40- by SO-foot wind tunnel. All three of them were two-bladed, teetering rotors. One of the rotors incorporated the NACA 0012 airfoil section over the entire length of the blade. This rotor was tested at advance ratios up to 1.05. Both of the other rotors were tapered in thickness and incorporated leading-edge camber over the outer 20 percent of the blade radius. The larger of these rotors was tested at advancing tip Mach numbers up to 1.02. Data were obtained for a wide range of lift and propulsive force, and are presented without discussion.

KSC-2012-5577

NASA Image and Video Library

2012-09-20

CAPE CANAVERAL, Fla. - Astronauts Mike Fossum and Cady Coleman, both in blue flight suits, look over the model capsule fit with rotor blades ahead of tests inside the Vehicle Assembly Building at NASA's Kennedy Space Center in Florida. NASA's Johnson Space Center Aerospace Engineer Jeff Hagen, right, fields questions about the project. The design would give a capsule the stability and control of a helicopter, but would not be powered. Instead, the wind passing over the rotors as the capsule descends would make the blades turn, a process called auto-rotation. The intent is to give real spacecraft a soft landing with enough control that they could touch down anywhere in the world, whether it be a runway or parking lot. In other words, wherever a helicopter could land, a spacecraft could land, too. Photo credit: NASA/Kim Shiflett

In-flight measurements of propeller blade deformation on a VUT100 cobra aeroplane using a co-rotating camera system

NASA Astrophysics Data System (ADS)

Boden, F.; Stasicki, B.; Szypuła, M.; Ružička, P.; Tvrdik, Z.; Ludwikowski, K.

2016-07-01

Knowledge of propeller or rotor blade behaviour under real operating conditions is crucial for optimizing the performance of a propeller or rotor system. A team of researchers, technicians and engineers from Avia Propeller, DLR, EVEKTOR and HARDsoft developed a rotating stereo camera system dedicated to in-flight blade deformation measurements. The whole system, co-rotating with the propeller at its full speed and hence exposed to high centrifugal forces and strong vibration, had been successfully tested on an EVEKTOR VUT 100 COBRA aeroplane in Kunovice (CZ) within the project AIM2—advanced in-flight measurement techniques funded by the European Commission (contract no. 266107). This paper will describe the work, starting from drawing the first sketch of the system up to performing the successful flight test. Apart from a description of the measurement hardware and the applied IPCT method, the paper will give some impressions of the flight test activities and discuss the results obtained from the measurements.

Small Business Innovations (Helicopters)

NASA Technical Reports Server (NTRS)

1992-01-01

The amount of engine power required for a helicopter to hover is an important, but difficult, consideration in helicopter design. The EHPIC program model produces converged, freely distorted wake geometries that generate accurate analysis of wake-induced downwash, allowing good predictions of rotor thrust and power requirements. Continuum Dynamics, Inc., the Small Business Innovation Research (SBIR) company that developed EHPIC, also produces RotorCRAFT, a program for analysis of aerodynamic loading of helicopter blades in forward flight. Both helicopter codes have been licensed to commercial manufacturers.

Prediction of Relaminarization Effects on Turbine Blade Heat Transfer

NASA Technical Reports Server (NTRS)

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

2001-01-01

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

A summary of wind tunnel research on tilt rotors from hover to cruise flight

NASA Technical Reports Server (NTRS)

Poisson-Quinton, PH.; Cook, W. L.

1972-01-01

An experimental research program has been conducted on a series of tilt rotors designed for a range of blade twist in various wind tunnel facilities. The objective was to obtain precise results on the influence of blade twist and aeroelasticity on tilt rotor performance, from hover to high speed cruise Mach number of about 0.7. global forces on the rotor, local loads and blade torsional deflection measurements were compared with theoretical predictions inside a large Reynolds-Mach envelope. Testing techniques developed during the program are described.

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.

Manufacturing Techniques for Application of Erosion Resistant Coatings to Turbine Engine Compressor Components.

DTIC Science & Technology

means of increasing the life of aircraft gas turbine compressor rotor blades and stator vanes . Two proprietary erosion resistant coating systems... engine tests as the two most promising systems for doubling compressor airfoil lives. An Air Force Sponsored program to evaluate the applicability of...Helicopter engine erosion has become a severe problem in S. E. Asia because of extensive operations in sand and dust. Hard coatings offer a potential

Measurements to Understand the Flow Mechanisms Contributing to Tandem Rotor Outwash

DTIC Science & Technology

2015-05-23

Directorate —AFDD Aviation & Missile Research, Development & Engineering Center Research, Development & Engineering Command Ames Research Center, Moffett...pilot visibility issues in brownout, dust entrain- ment into engine inlets, blade erosion, and increased air- craft maintenance. Though almost 50 years...Diameter Taylor , 1950 (Ref. 17) S, C, Ta 20 in & 45 in Fradenburgh, 1958 (Ref. 18) S 24 in Bolanovich & Marks, 1959 (Ref. 19) S 75 ft Bryan, 1960 (Ref

Variable diameter wind turbine rotor blades

DOEpatents

Jamieson, Peter McKeich; Hornzee-Jones, Chris; Moroz, Emilian M.; Blakemore, Ralph W.

2005-12-06

A system and method for changing wind turbine rotor diameters to meet changing wind speeds and control system loads is disclosed. The rotor blades on the wind turbine are able to adjust length by extensions nested within or containing the base blade. The blades can have more than one extension in a variety of configurations. A cable winching system, a hydraulic system, a pneumatic system, inflatable or elastic extensions, and a spring-loaded jack knife deployment are some of the methods of adjustment. The extension is also protected from lightning by a grounding system.

Experimental Investigation of Air-Cooled Turbine Blades in Turbojet Engine. 7: Rotor-Blade Fabrication Procedures

NASA Technical Reports Server (NTRS)

Long, Roger A.; Esgar, Jack B.

1951-01-01

An experimental investigation was conducted to determine the cooling effectiveness of a wide variety of air-cooled turbine-blade configurations. The blades, which were tested in the turbine of a - commercial turbojet engine that was modified for this investigation by replacing two of the original blades with air-cooled blades located diametrically opposite each other, are untwisted, have no aerodynamic taper, and have essentially the same external profile. The cooling-passage configuration is different for each blade, however. The fabrication procedures were varied and often unique. The blades were fabricated using methods most suitable for obtaining a small number of blades for use in the cooling investigations and therefore not all the fabrication procedures would be directly applicable to production processes, although some of the ideas and steps might be useful. Blade shells were obtained by both casting and forming. The cast shells were either welded to the blade base or cast integrally with the base. The formed shells were attached to the base by a brazing and two welding methods. Additional surface area was supplied in the coolant passages by the addition of fins or tubes that were S-brazed. to the shell. A number of blades with special leading- and trailing-edge designs that provided added cooling to these areas were fabricated. The cooling effectiveness and purposes of the various blade configurations are discussed briefly.

Deflection Shape Reconstructions of a Rotating Five-blade Helicopter Rotor from TLDV Measurements

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

Fioretti, A.; Castellini, P.; Tomasini, E. P.

2010-05-28

Helicopters are aircraft machines which are subjected to high level of vibrations, mainly due to spinning rotors. These are made of two or more blades attached by hinges to a central hub, which can make the dynamic behaviour difficult to study. However, they share some common dynamic properties with the ones expected in bladed discs, thereby the analytical modelling of rotors can be performed using some assumptions as the ones adopted for the bladed discs. This paper presents results of a vibrations study performed on a scaled helicopter rotor model which was rotating at a fix rotational speed and excitedmore » by an air jet. A simplified analytical model of that rotor was also produced to help the identifications of the vibration patterns measured using a single point tracking-SLDV measurement method.« less

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

Stability of elastic bending and torsion of uniform cantilever rotor blades in hover with variable structural coupling

NASA Technical Reports Server (NTRS)

Hodges, D. H., Roberta.

1976-01-01

The stability of elastic flap bending, lead-lag bending, and torsion of uniform, untwisted, cantilever rotor blades without chordwise offsets between the elastic, mass, tension, and areodynamic center axes is investigated for the hovering flight condition. The equations of motion are obtained by simplifying the general, nonlinear, partial differential equations of motion of an elastic rotating cantilever blade. The equations are adapted for a linearized stability analysis in the hovering flight condition by prescribing aerodynamic forces, applying Galerkin's method, and linearizing the resulting ordinary differential equations about the equilibrium operating condition. The aerodynamic forces are obtained from strip theory based on a quasi-steady approximation of two-dimensional unsteady airfoil theory. Six coupled mode shapes, calculated from free vibration about the equilibrium operating condition, are used in the linearized stability analysis. The study emphasizes the effects of two types of structural coupling that strongly influence the stability of hingeless rotor blades. The first structural coupling is the linear coupling between flap and lead-lag bending of the rotor blade. The second structural coupling is a nonlinear coupling between flap bending, lead-lag bending, and torsion deflections. Results are obtained for a wide variety of hingeless rotor configurations and operating conditions in order to provide a reasonably complete picture of hingeless rotor blade stability characteristics.

Effects of perforation number of blade on aerodynamic performance of dual-rotor small axial flow fans

NASA Astrophysics Data System (ADS)

Hu, Yongjun; Wang, Yanping; Li, Guoqi; Jin, Yingzi; Setoguchi, Toshiaki; Kim, Heuy Dong

2015-04-01

Compared with single rotor small axial flow fans, dual-rotor small axial flow fans is better regarding the static characteristics. But the aerodynamic noise of dual-rotor small axial flow fans is worse than that of single rotor small axial flow fans. In order to improve aerodynamic noise of dual-rotor small axial flow fans, the pre-stage blades with different perforation numbers are designed in this research. The RANS equations and the standard k-ɛ turbulence model as well as the FW-H noise model are used to simulate the flow field within the fan. Then, the aerodynamic performance of the fans with different perforation number is compared and analyzed. The results show that: (1) Compared to the prototype fan, the noise of fans with perforation blades is reduced. Additionally, the noise of the fans decreases with the increase of the number of perforations. (2) The vorticity value in the trailing edge of the pre-stage blades of perforated fans is reduced. It is found that the vorticity value in the trailing edge of the pre-stage blades decreases with the increase of the number of perforations. (3) Compared to the prototype fan, the total pressure rising and efficiency of the fans with perforation blades drop slightly.

A flight investigation of performance and loads for a helicopter with NLR-1T main-rotor blade sections

NASA Technical Reports Server (NTRS)

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

1979-01-01

Data on performance and rotor loads for a teetering-rotor, AH-1G helicopter flown with a main rotor that had the NLR-1T airfoil as the blade-section contour are presented. The test envelope included hover, forward-flight speed sweeps from 35 to 85 m/sec, and collective-fixed maneuvers at about 0.25 tip-speed ratio. The data set for each test point described vehicle flight state, control positions, rotor loads, power requirements, and blade motions. Rotor loads are reviewed primarily in terms of peak-to-peak and harmonic content. Lower frequency components predominated for most loads and generally increased with increased airspeed, but not necessarily with increased maneuver load factor.

CF6 jet engine performance improvement: New fan

NASA Technical Reports Server (NTRS)

Fasching, W. A.

1980-01-01

As part of the NASA sponsored engine component improvement program, and fan package was developed to reduce fuel consumption in current CF6 turbofan aircraft engine. The new fan package consist of an improved fan blade, reduced fan tip clearance due to a fan case stiffener, and a smooth fan casing tip shroud. CF6 engine performance and acoustic tests demonstrated the predicted 1.8% improvement in cruise sfc without an increase in engine noise. Power management thrust/fan speed characteristics were defined. Mechanical and structural integrity was demonstrated in model fan rotor photoelastic stress tests, full-size fan blade bench fatigue tests, and CF6 engine bird ingestion, crosswind, and cyclic endurance tests. The fan was certified in the CF6-500c2/E2 engines and is in commerical service on the Boeing 747-200, Douglas DC-10-30, and Atrbus industrie A300B aircraft.

Theory, Computation and Experiment on Criticality and Stability of Vortices Separating from Edges

DTIC Science & Technology

2016-08-15

aerospace engineering research. These include dynamic stall in wind turbines and helicopter rotors, and flapping-wing vehicle (micro-air vehicle) design...and Robinson, M., “Blade Three-Dimensional Dynamic Stall Response to Wind Turbine Operating Condition,” Journal of Solar Energy Engineering , Vol...Snapshots of TEV shedding in vortex ring representation. . . . . . . . . . . . . . . . 57 7.3 Schematic description of separated tip flow model

The validation and application of a rotor acoustic prediction computer program

NASA Technical Reports Server (NTRS)

Gallman, Judith M.

1990-01-01

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

Morphing Downwind-Aligned Rotor Concept Based on a 13-MW Wind Turbine

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

Ichter, Brian; Steele, Adam; Loth, Eric

To alleviate the mass-scaling issues associated with conventional upwind rotors of extreme-scale wind turbines (>/=10 MW), a morphing downwind-aligned rotor (MoDaR) concept is proposed herein. The concept employs a downwind rotor with blades whose elements are stiff (no intentional flexibility) but with hub-joints that can be unlocked to allow for moment-free downwind alignment. Aligning the combination of gravitational, centrifugal and thrust forces along the blade path reduces downwind cantilever loads, resulting in primarily tensile loading. For control simplicity, the blade curvature can be fixed with a single morphing degree of freedom using a near-hub joint for coning angle: 22 degreesmore » at rated conditions. The conventional baseline was set as the 13.2-MW Sandia 100-m all glass blade in a three-bladed upwind configuration. To quantify potential mass savings, a downwind load-aligning, two-bladed rotor was designed. Because of the reduced number of blades, the MoDaR concept had a favorable 33% mass reduction. The blade reduction and coning led to a reduction in rated power, but morphing increased energy capture at lower speeds such that both the MoDaR and conventional rotors have the same average power: 5.4 MW. A finite element analysis showed that quasi-steady structural stresses could be reduced, over a range of operating wind speeds and azimuthal angles, despite the increases in loading per blade. However, the concept feasibility requires additional investigation of the mass, cost and complexity of the morphing hinge, the impact of unsteady aeroelastic influence because of turbulence and off-design conditions, along with system-level Levelized Cost of Energy analysis.« less

Far-field noise and internal modes from a ducted propeller at simulated aircraft takeoff conditions

NASA Astrophysics Data System (ADS)

Woodward, Richard P.; Bock, Lawrence A.; Heidelberg, Laurence J.; Hall, David G.

1992-01-01

The ducted propeller offers structural and acoustic benefits typical of conventional turbofan engines while retaining much of the aeroacoustic benefits of the unducted propeller. A model Advanced Ducted Propeller (ADP) was tested in the NASA Lewis Low-Speed Anechoic Wind Tunnel at a simulated takeoff velocity of Mach 0.2. The ADP model was designed and manufactured by the Pratt and Whitney Division of United Technologies. The 16-blade rotor ADP was tested with 22- and 40-vane stators to achieve cut-on and cut-off criterion with respect to propagation of the fundamental rotor-stator interaction tone. Additional test parameters included three inlet lengths, three nozzle sizes, two spinner configurations, and two rotor rub strip configurations. The model was tested over a range of rotor blade setting angles and propeller axis angles-of-attack. Acoustic data were taken with a sideline translating microphone probe and with a unique inlet microphone probe which identified inlet rotating acoustic modes. The beneficial acoustic effects of cut-off were clearly demonstrated. A 5 dB fundamental tone reduction was associated with the long inlet and 40-vane sector, which may relate to inlet propeller axis angle-of-attack at rotor speeds of at least 96 percent design.

Far-field noise and internal modes from a ducted propeller at simulated aircraft takeoff conditions

NASA Astrophysics Data System (ADS)

Woodward, Richard P.; Bock, Lawrence A.; Heidelberg, Laurence J.; Hall, David G.

The ducted propeller offers structural and acoustic benefits typical of conventional turbofan engines while retaining much of the aeroacoustic benefits of the unducted propeller. A model Advanced Ducted Propeller (ADP) was tested in the NASA Lewis Low-Speed Anechoic Wind Tunnel at a simulated takeoff velocity of Mach 0.2. The ADP model was designed and manufactured by the Pratt and Whitney Division of United Technologies. The 16-blade rotor ADP was tested with 22- and 40-vane stators to achieve cut-on and cut-off criterion with respect to propagation of the fundamental rotor-stator interaction tone. Additional test parameters included three inlet lengths, three nozzle sizes, two spinner configurations, and two rotor rub strip configurations. The model was tested over a range of rotor blade setting angles and propeller axis angles-of-attack. Acoustic data were taken with a sideline translating microphone probe and with a unique inlet microphone probe which identified inlet rotating acoustic modes. The beneficial acoustic effects of cut-off were clearly demonstrated. A 5 dB fundamental tone reduction was associated with the long inlet and 40-vane sector, which may relate to inlet duct geometry. The fundamental tone level was essentially unaffected by propeller axis angle-of-attack at rotor speeds of at least 96 percent design.

Far-field noise and internal modes from a ducted propeller at simulated aircraft takeoff conditions

NASA Technical Reports Server (NTRS)

Woodward, Richard P.; Bock, Lawrence A.; Heidelberg, Laurence J.; Hall, David G.

1992-01-01

The ducted propeller offers structural and acoustic benefits typical of conventional turbofan engines while retaining much of the aeroacoustic benefits of the unducted propeller. A model Advanced Ducted Propeller (ADP) was tested in the NASA Lewis Low-Speed Anechoic Wind Tunnel at a simulated takeoff velocity of Mach 0.2. The ADP model was designed and manufactured by the Pratt and Whitney Division of United Technologies. The 16-blade rotor ADP was tested with 22- and 40-vane stators to achieve cut-on and cut-off criterion with respect to propagation of the fundamental rotor-stator interaction tone. Additional test parameters included three inlet lengths, three nozzle sizes, two spinner configurations, and two rotor rub strip configurations. The model was tested over a range of rotor blade setting angles and propeller axis angles-of-attack. Acoustic data were taken with a sideline translating microphone probe and with a unique inlet microphone probe which identified inlet rotating acoustic modes. The beneficial acoustic effects of cut-off were clearly demonstrated. A 5 dB fundamental tone reduction was associated with the long inlet and 40-vane sector, which may relate to inlet duct geometry. The fundamental tone level was essentially unaffected by propeller axis angle-of-attack at rotor speeds of at least 96 percent design.

A new dynamic model of rotor-blade systems

NASA Astrophysics Data System (ADS)

Ma, Hui; Lu, Yang; Wu, Zhiyuan; Tai, Xingyu; Li, Hui; Wen, Bangchun

2015-11-01

A new dynamic model of rotor-blade systems is developed in this paper considering the lateral and torsional deformations of the shaft, gyroscopic effects of the rotor which consists of shaft and disk, and the centrifugal stiffening, spin softening and Coriolis force of the blades. In this model, the rotating flexible blades are represented by Timoshenko beams. The shaft and rigid disk are described by multiple lumped mass points (LMPs), and these points are connected by massless springs which have both lateral and torsional stiffness. LMPs are represented by the corresponding masses and mass moments of inertia in lateral and rotational directions, where each point has five degrees of freedom (dofs) excluding axial dof. Equations of motion of the rotor-blade system are derived using Hamilton's principle in conjunction with the assumed modes method to describe blade deformation. The proposed model is compared with both finite element (FE) model and real experiments. The proposed model is first validated by comparing the model natural frequencies and vibration responses with those obtained from an FE model. A further verification of the model is then performed by comparing the model natural frequencies at zero rotational speed with those obtained from experimental studies. The results shown a good agreement between the model predicted system characteristics and those obtained from the FE model and experimental tests. Moreover, the following interesting phenomena have been revealed from the new model based analysis: The torsional natural frequency of the system decreases with the increase of rotational speed, and the frequency veering phenomenon has been observed at high rotational speed; The complicated coupling modes, such as the blade-blade coupling mode (BB), the coupling mode between the rotor lateral vibration and blade bending (RBL), and the coupling mode between the rotor torsional vibration and blade bending (RBT), have also been observed when the number of blades increases.

Concepts for a theoretical and experimental study of lifting rotor random loads and vibrations (further experiments with progressing/regressing rotor flapping modes), Phase 7-C

NASA Technical Reports Server (NTRS)

Hohenemser, K. H.; Crews, S. T.

1973-01-01

The experiments with progressing/regressing forced rotor flapping modes have been extended in several directions and the data processing method has been considerably refined. The 16 inch hingeless 2-bladed rotor model was equipped with a new set of high precision blades which removed previously encountered tracking difficulties at high advance ratio, so that tests up to .8 rotor advance ratio could be conducted. In addition to data with 1.20 blade natural flapping frequency data at 1.10 flapping frequency were obtained. Outside the wind tunnel, tests with a ground plate located at different distances below the rotor were conducted while recording the dynamic downflow at a station .2R below the rotor plane with a hot wire anemometer.

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.

Assessment of Geometry and In-Flow Effects on Contra-Rotating Open Rotor Broadband Noise Predictions

NASA Technical Reports Server (NTRS)

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

2015-01-01

Application of previously formulated semi-analytical models for the prediction of broadband noise due to turbulent rotor wake interactions and rotor blade trailing edges is performed on the historical baseline F31/A31 contra-rotating open rotor configuration. Simplified two-dimensional blade element analysis is performed on cambered NACA 4-digit airfoil profiles, which are meant to serve as substitutes for the actual rotor blade sectional geometries. Rotor in-flow effects such as induced axial and tangential velocities are incorporated into the noise prediction models based on supporting computational fluid dynamics (CFD) results and simplified in-flow velocity models. Emphasis is placed on the development of simplified rotor in-flow models for the purpose of performing accurate noise predictions independent of CFD information. The broadband predictions are found to compare favorably with experimental acoustic results.

Advanced Technology Blade testing on the XV-15 Tilt Rotor Research Aircraft

NASA Technical Reports Server (NTRS)

Wellman, Brent

1992-01-01

The XV-15 Tilt Rotor Research Aircraft has just completed the first series of flight tests with the Advanced Technology Blade (ATB) rotor system. The ATB are designed specifically for flight research and provide the ability to alter blade sweep and tip shape. A number of problems were encountered from first installation through envelope expansion to airplane mode flight that required innovative solutions to establish a suitable flight envelope. Prior to operation, the blade retention hardware had to be requalified to a higher rated centrifugal load, because the blade weight was higher than expected. Early flights in the helicopter mode revealed unacceptably high vibratory control system loads which required a temporary modification of the rotor controls to achieve higher speed flight and conversion to airplane mode. The airspeed in airplane mode was limited, however, because of large static control loads. Furthermore, analyses based on refined ATB blade mass and inertia properties indicated a previously unknown high-speed blade mode instability, also requiring airplane-mode maximum airspeed to be restricted. Most recently, a structural failure of an ATB cuff (root fairing) assembly retention structure required a redesign of the assembly. All problems have been addressed and satisfactory solutions have been found to allow continued productive flight research of the emerging tilt rotor concept.

Simultaneous Boundary-Layer Transition, Tip Vortex, and Blade Deformation Measurements of a Rotor in Hover

NASA Technical Reports Server (NTRS)

Heineck, James; Schairer, Edward; Ramasamy, Manikandan; Roozeboom, Nettie

2016-01-01

This paper describes simultaneous optical measurements of a sub-scale helicopter rotor in the U.S. Army Hover Chamber at NASA Ames Research Center. The measurements included thermal imaging of the rotor blades to detect boundary layer transition; retro-reflective background-oriented schlieren (RBOS) to visualize vortices; and stereo photogrammetry to measure displacements of the rotor blades, to compute spatial coordinates of the vortices from the RBOS data, and to map the thermal imaging data to a three-dimensional surface grid. The test also included an exploratory effort to measure flow near the rotor tip by tomographic particle image velocimetry (tomo PIV)an effort that yielded valuable experience but little data. The thermal imaging was accomplished using an image-derotation method that allowed long integration times without image blur. By mapping the thermal image data to a surface grid it was possible to accurately locate transition in spatial coordinates along the length of the rotor blade.

The Aerodynamic Performance of an Over-the-Rotor Liner With Circumferential Grooves on a High Bypass Ratio Turbofan Rotor

NASA Technical Reports Server (NTRS)

Bozak, Richard F.; Hughes, Christopher E.; Buckley, James

2013-01-01

While liners have been utilized throughout turbofan ducts to attenuate fan noise, additional attenuation is obtainable by placing an acoustic liner over-the-rotor. Previous experiments have shown significant fan performance losses when acoustic liners are installed over-the-rotor. The fan blades induce an oscillating flow in the acoustic liners which results in a performance loss near the blade tip. An over-the-rotor liner was designed with circumferential grooves between the fan blade tips and the acoustic liner to reduce the oscillating flow in the acoustic liner. An experiment was conducted in the W-8 Single-Stage Axial Compressor Facility at NASA Glenn Research Center on a 1.5 pressure ratio fan to evaluate the impact of this over-the-rotor treatment design on fan aerodynamic performance. The addition of a circumferentially grooved over-the-rotor design between the fan blades and the acoustic liner reduced the performance loss, in terms of fan adiabatic efficiency, to less than 1 percent which is within the repeatability of this experiment.

The Aerodynamic Performance of an Over-The-Rotor Liner with Circumferential Grooves on a High Bypass Ratio Turbofan Rotor

NASA Technical Reports Server (NTRS)

Bozak, Rick; Hughes, Christopher; Buckley, James

2013-01-01

While liners have been utilized throughout turbofan ducts to attenuate fan noise, additional attenuation is obtainable by placing an acoustic liner over-the-rotor. Previous experiments have shown significant fan performance losses when acoustic liners are installed over-the-rotor. The fan blades induce an oscillating flow in the acoustic liners which results in a performance loss near the blade tip. An over-the-rotor liner was designed with circumferential grooves between the fan blade tips and the acoustic liner to reduce the oscillating flow in the acoustic liner. An experiment was conducted in the W-8 Single-Stage Axial Compressor Facility at NASA Glenn Research Center on a 1.5 pressure ratio fan to evaluate the impact of this over-the-rotor treatment design on fan aerodynamic performance. The addition of a circumferentially grooved over-the-rotor design between the fan blades and the acoustic liner reduced the performance loss, in terms of fan adiabatic efficiency, to less than 1% which is within the repeatability of this experiment.

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.

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

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

Signal Separation of Helicopter Radar Returns Using Wavelet-Based Sparse Signal Optimisation

DTIC Science & Technology

2016-10-01

RR–0436 ABSTRACT A novel wavelet-based sparse signal representation technique is used to separate the main and tail rotor blade components of a...helicopter from the composite radar returns. The received signal consists of returns from the rotating main and tail rotor blades , the helicopter body...component signal com- prising of returns from the main body, the main and tail rotor hubs and blades . Temporal and Doppler characteristics of these

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

NASA Technical Reports Server (NTRS)

Kittleson, John K.; Yu, Yung H.

1987-01-01

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

High-power piezo drive amplifier for large stack and PFC applications

NASA Astrophysics Data System (ADS)

Clingman, Dan J.; Gamble, Mike

2001-08-01

This paper describes the continuing development of Boeing High Power Piezo Drive Amplifiers. Described is the development and testing of a 1500 Vpp, 8 amp switching amplifier. This amplifier is used to drive a piezo stack driven rotor blade trailing edge flap on a full size helicopter. Also discuss is a switching amplifier designed to drive a Piezo Fiber Composite (PFC) active twist rotor blade. This amplifier was designed to drive the PFC material at 2000 Vpp and 0.5 amps. These amplifiers recycle reactive energy, allowing for a power and weight efficient amplifier design. This work was done in conjunction with the DARPA sponsored Phase II Smart Rotor Blade program and the NASA Langley Research Center sponsored Active Twist Rotor (ATR) blade program.

An aerodynamic analysis of the autogiro rotor with a comparison between calculated and experimental results

NASA Technical Reports Server (NTRS)

Wheatley, John B

1935-01-01

This report presents an extension of the autogiro theory of Glauert and Lock in which the influence of a pitch varying with the blade radius is evaluated and methods of approximating the effect of blade tip losses and the influence of reversed velocities on the retreating blades are developed. A comparison of calculated and experimental results showed that most of the rotor characteristics could be calculated with reasonable accuracy, and that the type of induced flow assumed has a secondary effect upon the net rotor forces, although the flapping motion is influenced appreciably. An approximate evaluation of the effect of parasite drag on the rotor blades established the importance of including this factor in the analysis.

Modelling and attenuation feasibility of the aeroelastic response of active helicopter rotor systems during the engagement/disengagement phase of maritime operation

NASA Astrophysics Data System (ADS)

Khouli, F.

An aeroelastic phenomenon, known as blade sailing, encountered during maritime operation of helicopters is identified as being a factor that limits the tactical flexibility of helicopter operation in some sea conditions. The hazards associated with this phenomenon and its complexity, owing to the number of factors contributing to its occurrence, led previous investigators to conclude that advanced and validated simulation tools are best suited to investigate it. A research gap is identified in terms of scaled experimental investigation of this phenomenon and practical engineering solutions to alleviate its negative impact on maritime helicopter operation. The feasibility of a proposed strategy to alleviate it required addressing a gap in modelling thin-walled composite active beams/rotor blades. The modelling is performed by extending a mathematically-consistent and asymptotic reduction strategy of the 3-D elastic problem to account for embedded active materials. The derived active cross-sectional theory is validated using 2-D finite element results for closed and open cross-sections. The geometrically-exact intrinsic formulation of active maritime rotor systems is demonstrated to yield compact and symbolic governing equations. The intrinsic feature is shown to allow a classical and proven solution scheme to be successfully applied to obtain time history solutions. A Froude-scaled experimental rotor was designed, built, and tested in a scaled ship airwake environment and representative ship motion. Based on experimental and simulations data, conclusions are drawn regarding the influence of the maritime operation environment and the rotor operation parameters on the blade sailing phenomenon. The experimental data is also used to successfully validate the developed simulation tools. The feasibility of an open-loop control strategy based on the integral active twist concept to counter blade sailing is established in a Mach-scaled maritime operation environment. Recommendations are proposed to improve the strategy and further establish its validity in a full-scale maritime operation environment.

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

NASA Technical Reports Server (NTRS)

Hanson, D. B.

1975-01-01

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

Fabrication of complex structures or assemblies by Hot Isostatic Pressure (HIP) welding

NASA Technical Reports Server (NTRS)

Ashurst, A. N.; Goldstein, M.; Ryan, M. J.; Lessmann, G. G.; Bryant, W. A.

1974-01-01

HIP welding is effective method for fabricating complex structures or assemblies such as alternator rotors, regeneratively-cooled rocket-motor thrust chambers, and jet engine turbine blades. It can be applied to fabrication of many assemblies which require that component parts be welded together along complex interfaces.

Estimation of blade airloads from rotor blade bending moments

NASA Technical Reports Server (NTRS)

Bousman, William G.

1987-01-01

A method is developed to estimate the blade normal airloads by using measured flap bending moments; that is, the rotor blade is used as a force balance. The blade's rotation is calculated in vacuum modes and the airloads are then expressed as an algebraic sum of the mode shapes, modal amplitudes, mass distribution, and frequency properties. The modal amplitudes are identified from the blade bending moments using the Strain Pattern Analysis Method. The application of the method is examined using simulated flap bending moment data that have been calculated for measured airloads for a full-scale rotor in a wind tunnel. The estimated airloads are compared with the wind tunnel measurements. The effects of the number of measurements, the number of modes, and errors in the measurements and the blade properties are examined, and the method is shown to be robust.

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.

Analysis and Testing of a Composite Fuselage Shield for Open Rotor Engine Blade-Out Protection

NASA Technical Reports Server (NTRS)

Pereira, J. Michael; Emmerling, William; Seng, Silvia; Frankenberger, Charles; Ruggeri, Charles R.; Revilock, Duane M.; Carney, Kelly S.

2016-01-01

The Federal Aviation Administration is working with the European Aviation Safety Agency to determine the certification base for proposed new engines that would not have a containment structure on large commercial aircraft. Equivalent safety to the current fleet is desired by the regulators, which means that loss of a single fan blade will not cause hazard to the Aircraft. The NASA Glenn Research Center and The Naval Air Warfare Center (NAWC), China Lake, collaborated with the FAA Aircraft Catastrophic Failure Prevention Program to design and test lightweight composite shields for protection of the aircraft passengers and critical systems from a released blade that could impact the fuselage. LS-DYNA® was used to predict the thickness of the composite shield required to prevent blade penetration. In the test, two composite blades were pyrotechnically released from a running engine, each impacting a composite shield with a different thickness. The thinner shield was penetrated by the blade and the thicker shield prevented penetration. This was consistent with pre-test LS-DYNA predictions. This paper documents the analysis conducted to predict the required thickness of a composite shield, the live fire test from the full scale rig at NAWC China Lake and describes the damage to the shields as well as instrumentation results.

Tilt rotor hover aeroacoustics

NASA Technical Reports Server (NTRS)

Coffen, Charles David

1992-01-01

The methodology, results, and conclusions of a study of tilt rotor hover aeroacoustics and aerodynamics are presented. Flow visualization and hot wire velocity measurement were performed on a 1/12-scale model of the XV-15 Tilt Rotor Aircraft in hover. The wing and fuselage below the rotor cause a complex recirculating flow. Results indicate the physical dimensions and details of the flow including the relative unsteadiness and turbulence characteristics of the flow. Discrete frequency harmonic thickness and the loading noise mechanism were predicted using WOPWOP for the standard metal blades and the Advanced Technology Blades. The recirculating flow created by the wing below the rotor is a primary sound mechanism for a hovering tilt rotor. The effects of dynamic blade response should be included for fountain flow conditions which produce impulsive blade loading. Broadband noise mechanisms were studied using Amiet's method with azimuthally varying turbulence characteristics derived from the measurements. The recirculating fountain flow with high turbulence levels in the recirculating zone is the dominant source of broadband noise for a hovering rotor. It is shown that tilt rotor hover aeroacoustic noise mechanisms are now understood. Noise predictions can be made based on reasonably accurate aerodynamic models developed here.

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.

Active-Twist Rotor Control Applications for UAVs

NASA Technical Reports Server (NTRS)

Wilbur, Matthew L.; Wilkie, W. Keats

2004-01-01

The current state-of-the-art in active-twist rotor control is discussed using representative examples from analytical and experimental studies, and the application to rotary-wing UAVs is considered. Topics include vibration and noise reduction, rotor performance improvement, active blade tracking, stability augmentation, and rotor blade de-icing. A review of the current status of piezoelectric fiber composite actuator technology, the class of piezoelectric actuators implemented in active-twist rotor systems, is included.

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

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

Large Wind Energy Converter: Growian 3 MW

NASA Technical Reports Server (NTRS)

Feustel, J. E.; Helm, S.; Koerber, F.

1980-01-01

The final report on the projected application of larger-scale wind turbine on the northern German coast is summarized. The designs of the tower, machinery housing, rotor, and rotor blades are described accompanied various construction materials are examined. Rotor blade adjustment devices auxiliary and accessory equipment are examined.

Performance Evaluation of Reduced-Chord Rotor Blading as Applied to J73 Two-Stage Turbine

NASA Technical Reports Server (NTRS)

Schurn, Harold J.

1957-01-01

The multistage turbine from the J73 turbojet engine has previously been investigated with standard and with reduced-chord rotor blading in order to determine the individual performance characteristics of each configuration over a range of over-all pressure ratio and speed. Because both turbine configurations exhibited peak efficiencies of over 90 percent, and because both units had relatively wide efficient operating ranges, it was considered of interest to determine the performance of the first stage of the turbine as a separate component. Accordingly, the standard-bladed multistage turbine was modified by removing the second-stage rotor disk and stator and altering the flow passage so that the first stage of the unit could be operated independently. The modified single-stage turbine was then operated over a range of stage pressure ratio and speed. The single-stage turbine operated at a peak brake internal efficiency of over 90 percent at an over-all stage pressure ratio of 1.4 and at 90 percent of design equivalent speed. Furthermore, the unit operated at high efficiencies over a relatively wide operating range. When the single-stage results were compared with the multistage results at the design operating point, it was found that the first stage produced approximately half the total multistage-turbine work output.

HHC study in the DNW to reduce BVI noise - An analysis

NASA Technical Reports Server (NTRS)

Brooks, Thomas F.; Booth, Earl R., Jr.; Boyd, D. D., Jr.; Splettstoesser, Wolf R.; Schultz, Klaus -J.; Kube, Roland; Niesl, Georg H.; Streby, Olivier

1991-01-01

The noise of an aeroelastically scaled helicopter rotor has been studied in the German-Dutch wind tunnel in order to assess the utility of higher-harmonic control (HHC) in reducing blade-vortex interaction (BVI) noise. Acoustic data are presented for 3/rev, 4/rev, and 5/rev HHC, as applied to a typical landing approach rotor operating condition; noise reduction of up to 6 dB were found for advancing-blade BVI noise radiating upstream of the rotor, as well as for retreating blade BVI noise radiating below and downstream of the rotor.

A Note about Self-Induced Velocity Generated by a Lifting-Line Wing or Rotor Blade

NASA Technical Reports Server (NTRS)

Harris, Franklin D.

2006-01-01

This report presents an elementary analysis of the induced velocity created by a field of vortices that reside in the wake of a rotor blade. Progress achieved by other researchers in the last 70 years is briefly reviewed. The present work is presented in four stages of complexity that carry a lifting-line representation of a fixed wing into a single-blade rotor. The analysis leads to the conclusion that the lifting rotor's spiraling vortex wake structure has very high induced power when compared to the ideal wing. For an advanced ratio of one-half, induced power is on the order of 10 times that of the wing when the comparison is made at wingspan equal to rotor diameter and wing and rotor having equal lift.

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

NASA Astrophysics Data System (ADS)

Carolus, Thomas

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

Cyclogiro windmill

DOEpatents

Brulle, R.V.

1981-09-03

A cyclogiro windmill has a rotor provided with blades shaped in the configuration of symmetrical airfoils and actuators to pivot the blades about axes parallel to the axis of rotation for the rotor. The actuator for each blade constantly changes the rock angle for the blade, that is its angle with respect to a reference on the rotor, and this modulation is such that the blade in making a revolution around the axis of rotation for the rotor undergoes an interval of static operation wherein its angle of attack is for the most part constant and less than the static stall angle, a short interval where the blade flips to position in which its opposite surface is presented toward the free wind, a short interval of dynamic operation wherein the angle of attack exceeds the static stal angle, another interval of static operation at an angle of attack of essentially the same magnitude as before, another interval of blade flip, and another interval of dynamic operation. During the intervals of dynamic operation, the blades experience a significant increase in lift force without a corresponding increase in drag, so that a high lift-to-drag ratio develops. The blades during dynamic operation further develop strong vortices which are directed outwardly at the sides of the windmill stream tube, and this increases the width of the stream tube, causing a greater mass of air to flow through the rotor. The short intervals of operation under dynamic conditions enable the blades to extract more energy from the free wind than would be possible if the blade operated solely under static conditions, and this in turn renders the windmill more useful in moderate velocity winds as well as high velocity winds.

Cyclogiro windmill

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

Brulle, R.V.

1981-09-03

A cyclogiro windmill has a rotor provided with blades shaped in the configuration of symmetrical airfoils and actuators to pivot the blades about axes parallel to the axis of rotation for the rotor. The actuator for each blade constantly changes the rock angle for the blade, that is its angle with respect to a reference on the rotor, and this modulation is such that the blade in making a revolution around the axis of rotation for the rotor undergoes an interval of static operation wherein its angle of attack is for the most part constant and less than the staticmore » stall angle, a short interval where the blade flips to position in which its opposite surface is presented toward the free wind, a short interval of dynamic operation wherein the angle of attack exceeds the static stal angle, another interval of static operation at an angle of attack of essentially the same magnitude as before, another interval of blade flip, and another interval of dynamic operation. During the intervals of dynamic operation, the blades experience a significant increase in lift force without a corresponding increase in drag, so that a high lift-to-drag ratio develops. The blades during dynamic operation further develop strong vortices which are directed outwardly at the sides of the windmill stream tube, and this increases the width of the stream tube, causing a greater mass of air to flow through the rotor. The short intervals of operation under dynamic conditions enable the blades to extract more energy from the free wind than would be possible if the blade operated solely under static conditions, and this in turn renders the windmill more useful in moderate velocity winds as well as high velocity winds.« less

Analyzing Aeroelastic Stability of a Tilt-Rotor Aircraft

NASA Technical Reports Server (NTRS)

Kvaternil, Raymond G.

2006-01-01

Proprotor Aeroelastic Stability Analysis, now at version 4.5 (PASTA 4.5), is a FORTRAN computer program for analyzing the aeroelastic stability of a tiltrotor aircraft in the airplane mode of flight. The program employs a 10-degree- of-freedom (DOF), discrete-coordinate, linear mathematical model of a rotor with three or more blades and its drive system coupled to a 10-DOF modal model of an airframe. The user can select which DOFs are included in the analysis. Quasi-steady strip-theory aerodynamics is employed for the aerodynamic loads on the blades, a quasi-steady representation is employed for the aerodynamic loads acting on the vibrational modes of the airframe, and a stability-derivative approach is used for the aerodynamics associated with the rigid-body DOFs of the airframe. Blade parameters that vary with the blade collective pitch can be obtained by interpolation from a user-defined table. Stability is determined by examining the eigenvalues that are obtained by solving the coupled equations of motions as a matrix eigenvalue problem. Notwithstanding the relative simplicity of its mathematical foundation, PASTA 4.5 and its predecessors have played key roles in a number of engineering investigations over the years.

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

NASA Astrophysics Data System (ADS)

Smith, Natalie Rochelle

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

Determination of the number of Vertical Axis Wind Turbine blades based on power spectrum

NASA Astrophysics Data System (ADS)

Fedak, Waldemar; Anweiler, Stanisław; Gancarski, Wojciech; Ulbrich, Roman

2017-10-01

Technology of wind exploitation has been applied widely all over the world and has already reached the level in which manufacturers want to maximize the yield with the minimum investment outlays. The main objective of this paper is the determination of the optimal number of blades in the Cup-Bladed Vertical Axis Wind Turbine. Optimizing the size of the Vertical Axis Wind Turbine allows the reduction of costs. The maximum power of the rotor is selected as the performance target. The optimum number of Vertical Axis Wind Turbine blades evaluation is based on analysis of a single blade simulation and its superposition for the whole rotor. The simulation of working blade was done in MatLab environment. Power spectrum graphs were prepared and compared throughout superposition of individual blades in the Vertical Axis Wind Turbine rotor. The major result of this research is the Vertical Axis Wind Turbine power characteristic. On the basis of the analysis of the power spectra, optimum number of the blades was specified for the analysed rotor. Power spectrum analysis of wind turbine enabled the specification of the optimal number of blades, and can be used regarding investment outlays and power output of the Vertical Axis Wind Turbine.

Flowfield Characteristics on a Retreating Rotor Blade

DTIC Science & Technology

2015-12-03

dimensional airfoil aerodynamics. This project used stereo particle image velocimetry on a 2-bladed rotor at advance ratios of 0.7, 0.85 and 1.0...ABSTRACT 2. REPORT TYPE 17. LIMITATION OF ABSTRACT 15. NUMBER OF PAGES 5d. PROJECT NUMBER 5e. TASK NUMBER 5f. WORK UNIT NUMBER 5c. PROGRAM...attempted to make yaw corrections to 2- dimensional airfoil aerodynamics. This project used stereo particle image velocimetry on a 2-bladed rotor at advance

Numerical Investigations of the Influence of Unsteady Vane Trailing Edge Shock Wave on Film Cooling Effectiveness of Rotor Blade Leading Edge

NASA Astrophysics Data System (ADS)

Wang, Yufeng; Cai, Le; Wang, Songtao; Zhou, Xun

2018-04-01

Unsteady numerical simulations of a high-load transonic turbine stage have been carried out to study the influences of vane trailing edge outer-extending shockwave on rotor blade leading edge film cooling performance. The turbine stage used in this paper is composed of a vane section and a rotor one which are both near the root section of a transonic high-load turbine stage. The Mach number is 0.94 at vane outlet, and the relative Mach number is above 1.10 at rotor outlet. Various positions and oblique angles of film cooling holes were investigated in this research. Results show that the cooling efficiency on the blade surface of rotor near leading edge is significantly affected by vane trailing edge outer-extending shockwave in some cases. In the cases that film holes are close to leading edge, cooling performance suffers more from the sweeping vane trailing edge outer-extending shockwave. In addition, coolant flow ejected from oblique film holes is harder to separate from the blade surface of rotor, and can cover more blade area even under the effects of sweeping vane trailing edge shockwave. As a result, oblique film holes can provide better film cooling performance than vertical film holes do near the leading edge on turbine blade which is swept by shockwaves.

Performance of Axial-Flow Supersonic Compressor of the XJ55-FF-1 Turbojet Engine. IV - Analysis of Compressor Operation over a Range of Equivalent Tip Speeds from 801 to 1614 Feet Per Second

NASA Technical Reports Server (NTRS)

Graham, Robert C.; Hartmann, Melvin J.

1949-01-01

An investigation was conducted to determine the performance characteristics of the axial-flow supersonic compressor of the XJ55-FF-1 turbojet engine. An analysis of the performance of the rotor was made based on detailed flow measurements behind the rotor. The compressor apparently did not obtain the design normal-shock configuration in this investigation. A large redistribution of mass occurred toward the root of the rotor over the entire speed range; this condition was so acute at design speed that the tip sections were completely inoperative. The passage pressure recovery at maximum pressure ratio at 1614 feet per second varied from a maximum of 0.81 near the root to 0.53 near the tip, which indicated very poor efficiency of the flow process through the rotor. The results, however, indicated that the desired supersonic operation may be obtained by decreasing the effective contraction ratio of the rotor blade passage.

Recent advances in integrated multidisciplinary optimization of rotorcraft

NASA Technical Reports Server (NTRS)

Adelman, Howard M.; Walsh, Joanne L.; Pritchard, Jocelyn I.

1992-01-01

A joint activity involving NASA and Army researchers at NASA LaRC to develop optimization procedures to improve the rotor blade design process by integrating appropriate disciplines and accounting for all of the important interactions among the disciplines is described. 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 are closely coupled while acoustics and airframe dynamics are decoupled and are accounted for as effective constraints on the design for the first three disciplines. In phase 2, acoustics is integrated with the first three disciplines. Finally, in phase 3, airframe dynamics is integrated with the other four disciplines. Representative results from work performed to date are described. These include optimal placement of tuning masses for reduction of blade vibratory shear forces, integrated aerodynamic/dynamic optimization, and integrated aerodynamic/dynamic/structural optimization. Examples of validating procedures are described.

Recent advances in multidisciplinary optimization of rotorcraft

NASA Technical Reports Server (NTRS)

Adelman, Howard M.; Walsh, Joanne L.; Pritchard, Jocelyn I.

1992-01-01

A joint activity involving NASA and Army researchers at NASA LaRC to develop optimization procedures to improve the rotor blade design process by integrating appropriate disciplines and accounting for all of the important interactions among the disciplines is described. 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 are closely coupled while acoustics and airframe dynamics are decoupled and are accounted for as effective constraints on the design for the first three disciplines. In phase 2, acoustics is integrated with the first three disciplines. Finally, in phase 3, airframe dynamics is integrated with the other four disciplines. Representative results from work performed to date are described. These include optimal placement of tuning masses for reduction of blade vibratory shear forces, integrated aerodynamic/dynamic optimization, and integrated aerodynamic/dynamic/structural optimization. Examples of validating procedures are described.

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.

Influence of the conservative rotor loads on the near wake of a wind turbine

NASA Astrophysics Data System (ADS)

Herráez, I.; Micallef, D.; van Kuik, G. A. M.

2017-05-01

The presence of conservative forces on rotor blades is neglected in the blade element theory and all the numerical methods derived from it (like e.g. the blade element momentum theory and the actuator line technique). This might seem a reasonable simplification of the real flow of rotor blades, since conservative loads, by definition, do not contribute to the power conversion. However, conservative loads originating from the chordwise bound vorticity might affect the tip vortex trajectory, as we discussed in a previous work. In that work we also hypothesized that this effect, in turn, could influence the wake induction and correspondingly the rotor performance. In the current work we extend a standard actuator line model in order to account for the conservative loads at the blade tip. This allows to isolate the influence of conservative forces from other effects. The comparison of numerical results with and without conservative loads enables to confirm qualitatively their relevance for the near wake and the rotor performance. However, an accurate quantitative assessment of the effect still remains out of reach due to the inherent uncertainty of the numerical model.

Rotary-wing aerodynamics. Volume 1: Basic theories of rotor aerodynamics with application to helicopters. [momentum, vortices, and potential theory

NASA Technical Reports Server (NTRS)

Stepniewski, W. Z.

1979-01-01

The concept of rotary-wing aircraft in general is defined. The energy effectiveness of helicopters is compared with that of other static thrust generators in hover, as well as with various air and ground vehicles in forward translation. The most important aspects of rotor-blade dynamics and rotor control are reviewed. The simple physicomathematical model of the rotor offered by the momentum theory is introduced and its usefulness and limitations are assessed. The combined blade-element and momentum theory approach, which provides greater accuracy in performance predictions, is described as well as the vortex theory which models a rotor blade by means of a vortex filament or vorticity surface. The application of the velocity and acceleration potential theory to the determination of flow fields around three dimensional, non-rotating bodies as well as to rotor aerodynamic problems is described. Airfoil sections suitable for rotors are also considered.

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.

Performance of Swashplateless Ultralight Helicopter Rotor with Trailing-edge Flaps for Primary Flight Control

NASA Technical Reports Server (NTRS)

Shen, Jin-Wei; Chopra, Inderjit

2003-01-01

The objective of present study is to evaluate the rotor performance, trailing-edge deflections and actuation requirement of a helicopter rotor with trailing-edge flap system for primary flight control. The swashplateless design is implemented by modifying a two-bladed teetering rotor of an production ultralight helicopter through the use of plain flaps on the blades, and by replacing the pitch link to fixed system control system assembly with a root spring. A comprehensive rotorcraft analysis based on UMARC is carried out to obtain the results for both the swashplateless and a conventional baseline rotor configuration. The predictions show swashplateless configuration achieve superior performance than the conventional rotor attributed from reduction of parasite drag by eliminating swashplate mechanic system. It is indicated that optimal selection of blade pitch index angle, flap location, length, and chord ratio reduces flap deflections and actuation requirements, however, has virtually no effect on rotor performance.

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.

Helicopters for the future

NASA Technical Reports Server (NTRS)

Ward, J. F.

1984-01-01

Technology needed to provide the basis for creating a widening rotary wing market include: well defined and proven design; reductions in noise, vibration, and fuel consumption; improvement of flying and ride quality; better safety; reliability; maintainability; and productivity. Unsteady transonic flow, yawed flow, dynamic stall, and blade vortex interaction are some of the problems faced by scientists and engineers in the helicopter industry with rotorcraft technology seen as an important development for future advanced high speed vehicle configurations. Such aircraft as the Boeing Vertol medium lift Model 360 composite aircraft, the Sikorsky Advancing Blade Concept (ABC) aircraft, the Bell Textron XV-15 Tilt Rotor Aircraft, and the X-wing rotor aircraft are discussed in detail. Even though rotorcraft technology has become an integral part of the military scene, the potential market for its civil applications has not been fully developed.

Investigating Dynamics of Eccentricity in Turbomachines

NASA Technical Reports Server (NTRS)

Baun, Daniel

2010-01-01

A methodology (and hardware and software to implement the methodology) has been developed as a means of investigating coupling between certain rotordynamic and hydrodynamic phenomena in turbomachines. Originally, the methodology was intended for application in an investigation of coupled rotordynamic and hydrodynamic effects postulated to have caused high synchronous vibration in the space shuttle s high-pressure oxygen turbopump (HPOTP). The methodology can also be applied in investigating (for the purpose of developing means of suppressing) undesired hydrodynamic rotor/stator interactions in turbomachines in general. The methodology and the types of phenomena that can be investigated by use of the methodology are best summarized by citing the original application as an example. In that application, in consideration of the high synchronous vibration in the space-shuttle main engine (SSME) HPOTP, it was determined to be necessary to perform tests to investigate the influence of inducer eccentricity and/or synchronous whirl motion on inducer hydrodynamic forces under prescribed flow and cavitation conditions. It was believed that manufacturing tolerances of the turbopump resulted in some induced runout of the pump rotor. Such runout, if oriented with an inducer blade, would cause that blade to run with tip clearance smaller than the tip clearances of the other inducer blades. It was hypothesized that the resulting hydraulic asymmetry, coupled with alternating blade cavitation, could give rise to the observed high synchronous vibration. In tests performed to investigate this hypothesis, prescribed rotor whirl motions have been imposed on a 1/3-scale water-rig version of the SSME LPOTP inducer (which is also a 4-biased inducer having similar cavitation dynamics as the HPOTP) in a magnetic-bearing test facility. The particular magnetic-bearing test facility, through active vibration control, affords a capability to impose, on the rotor, whirl orbits having shapes and whirl rates prescribed by the user, and to simultaneously measure the resulting hydrodynamic forces generated by the impeller. Active control also made it possible to modulate the inducer-blade running tip clearance and consequently effect alternating blade cavitation. The measured hydraulic forces have been compared and correlated with shroud dynamic-pressure measurements.

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.

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.

An experimental evaluation of the performance deficit of an aircraft engine starter turbine

NASA Technical Reports Server (NTRS)

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

1980-01-01

An experimental investigation was made to determine the reasons for the low aerodynamic performance of a 13.5 centimeter tip diameter aircraft engine starter turbine. The investigation consisted of an evaluation of both the stator and the stage. An approximate ten percent improvement in turbine efficiency was obtained when the honeycomb shroud over the rotor blade tips was filled to obtain a solid shroud surface.

Endwall Treatment and Method for Gas Turbine

NASA Technical Reports Server (NTRS)

Hathaway, Michael D. (Inventor); Strazisar, Anthony J. (Inventor); Suder, Kenneth L. (Inventor)

2006-01-01

An endwall treatment for a gas turbine engine having at least one rotor blade extending from a rotatable hub and a casing circumferentially surrounding the rotor and the hub, the endwall treatment including, an inlet formed in an endwall of the gas turbine engine adapted to ingest fluid from a region of a higher-pressure fluid, an outlet formed in the endwall and located in a region of lower pressure than the inlet, wherein the inlet and the outlet are in a fluid communication with each other, the outlet being adapted to inject the fluid from the inlet in the region of lower pressure, and wherein the outlet is at least partially circumferentially offset relative to the inlet.

Aerodynamic loads on a Darrieus rotor blade

NASA Astrophysics Data System (ADS)

Wilson, R. E.; McKie, W. R.; Lissaman, P. B. S.; James, M.

1983-03-01

A method is presented for the free vortex analysis of a Darrieus rotor blade in nonsteady motion, which employs the circle theorem to map the moving rotor airfoil into the circle plane and models the wake generated in terms of point vortices. Nascent vortex strength and position are taken from the Kutta condition, so that the nascent vortex has the same strength as a vortex sheet of uniform strength. Pressure integration over the plate and wake vortex impulse methods yields the same numerical results. The numerical results presented for a one-bladed Darrieus rotor at a tip/speed ratio of three, and two different chord sizes, indicate that the moment on the blade can be adequately approximated by quasi-steady relationships, although the accurate determination of local velocity and circulation are still required.

Modeling and Design of a Full-Scale Rotor Blade with Embedded Piezocomposite Actuators

NASA Astrophysics Data System (ADS)

Kovalovs, A.; Barkanov, E.; Ruchevskis, S.; Wesolowski, M.

2017-05-01

An optimization methodology for the design of a full-scale rotor blade with an active twist in order to enhance its ability to reduce vibrations and noise is presented. It is based on a 3D finite-element model, the planning of experiments, and the response surface technique to obtain high piezoelectric actuation forces and displacements with a minimum actuator weight and energy applied. To investigate an active twist of the helicopter rotor blade, a structural static analysis using a 3D finite-element model was carried out. Optimum results were obtained at two possible applications of macrofiber composite actuators. The torsion angle found from the finite-element simulation of helicopter rotor blades was successfully validated by its experimental values, which confirmed the modeling accuracy.

Results of turbojet engine operation tests using a 50-50 mixture of JP-4 and tributyl borate as the fuel

NASA Technical Reports Server (NTRS)

Schafer, Louis J , Jr; Stepka, Francis S

1957-01-01

An experimental investigation was conducted on a centrifugal-type turbojet engine using a 50-50 mixture of tributyl borate and JP-4 as the fuel to determine the magnitude and the location of the boric oxide deposits in the engine as well as the effect of these deposits on the engine performance. Large deposits of boric acid formed in the combustor walls and on the turbine rotor and stator blades. The deposits had no effect on the engine thrust.

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

NASA Technical Reports Server (NTRS)

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

1977-01-01

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

System Identification of Mistuned Bladed Disks from Traveling Wave Response Measurements

NASA Technical Reports Server (NTRS)

Feiner, D. M.; Griffin, J. H.; Jones, K. W.; Kenyon, J. A.; Mehmed, O.; Kurkov, A. P.

2003-01-01

A new approach to modal analysis is presented. By applying this technique to bladed disk system identification methods, one can determine the mistuning in a rotor based on its response to a traveling wave excitation. This allows system identification to be performed under rotating conditions, and thus expands the applicability of existing mistuning identification techniques from integrally bladed rotors to conventional bladed disks.

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.

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.

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

DTIC Science & Technology

2012-05-01

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

Induced strain actuation of composite beams and rotor blades with embedded piezoceramic elements

NASA Astrophysics Data System (ADS)

Chen, Peter C.; Chopra, Inderjit

1994-05-01

The objective of this research is to develop a dynamically scaled (Froude scale) helicopter rotor blade with embedded piezoceramic elements as sensors and actuators to control blade vibrations. A 6-ft-diameter two-bladed bearingless rotor model was built, where each blade is embedded with banks of piezoelectric actuators at +/- 45-degree angles with respect to the beam axis on the top and bottom surfaces. A twist distribution along the blade span is achieved through in-phase excitation of the top and bottom actuators at equal potentials, while a bending distribution is achieved through out-of-phase excitation. In order to fix design variables and to optimize blade performance, a uniform strain beam theory is formulated to analytically predict the static bending and torsional response of composite rectangular beams with embedded piezoelectric actuators. Parameters such as bond thicknesses, actuator skew angle, and actuator spacing are investigated by experiments and then validated by theory. The static bending and torsional response of the rotor blades is experimentally measured and correlated with theory. Dynamic torsional and bending responses are experimentally determined for frequencies from 2-120 HZ to assess the viability of a vibration reduction system based on piezoactuation of blade twist. Although the magnitudes of blade twist attained in this experiment were small, it is expected that future models can be built with improved performance.

Methods for Expanding Rotary Wing Aircraft Health and Usage Monitoring Systems to the Rotating Frame through Real-time Rotor Blade Kinematics Estimation

NASA Astrophysics Data System (ADS)

Allred, Charles Jefferson

Since the advent of Health and Usage Monitoring Systems (HUMS) in the early 1990's, there has been a steady decrease in the number of component failure related helicopter accidents. Additionally, measurable cost benefits due to improved maintenance practices based on HUMS data has led to a desire to expand HUMS from its traditional area of helicopter drive train monitoring. One of the areas of greatest interest for this expansion of HUMS is monitoring of the helicopter rotor head loads. Studies of rotor head load and blade motions have primarily focused on wind tunnel testing with technology which would not be applicable for production helicopter HUMS deployment, or measuring bending along the blade, rather than where it is attached to the rotor head and the location through which all the helicopter loads pass. This dissertation details research into finding methods for real time methods of estimating rotor blade motion which could be applied across helicopter fleets as an expansion of current HUMS technology. First, there is a brief exploration of supporting technologies which will be crucial in enabling the expansion of HUMS from the fuselage of helicopters to the rotor head: wireless data transmission and energy harvesting. A brief overview of the commercially available low power wireless technology selected for this research is presented. The development of a relatively high-powered energy harvester specific to the motion of helicopter rotor blades is presented and two different prototypes of the device are shown. Following the overview of supporting technologies, two novel methods of monitoring rotor blade motion in real time are developed. The first method employs linear displacement sensors embedded in the elastomer layers of a high-capacity laminate bearing of the type commonly used in fully articulated rotors throughout the helicopter industry. The configuration of these displacement sensors allows modeling of the sensing system as a robotic parallel mechanism, similar to a Stewart Platform. A calibration method for this device is developed and the improved orientation estimation results are shown. The second method is not specific to the fully articulated rotor head mounting geometry of the first method. Rather, it utilizes micro-electromechanical (MEMS) accelerometers and gyroscopes configured to measure the centrifugal acceleration and rotation rate induced through rotor head rotation differentially. By measuring these quantities differentially, other accelerations from the fuselage reference frame are removed from the measurement, resulting in acceleration and rate quantities that are impacted only by the angle of the sensors relative to the plane of rotation. By mounting these sensors strategically and symmetrically about the rotor blade root center of rotation, the orientation of the rotor blade can be estimated in real time.

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

NASA Technical Reports Server (NTRS)

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

1973-01-01

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

Overall and blade element performance of a 1.20-pressure-ratio fan stage with rotor blades reset -5 deg

NASA Technical Reports Server (NTRS)

Lewis, G. W., Jr.; Osborn, W. M.; Moore, R. D.

1976-01-01

A 51-cm-diam model of a fan stage for a short haul aircraft was tested in a single stage-compressor research facility. The rotor blades were set 5 deg toward the axial direction (opened) from design setting angle. Surveys of the air flow conditions ahead of the rotor, between the rotor and stator, and behind the stator were made over the stable operating range of the stage. At the design speed of 213.3 m/sec and a weight flow of 31.5 kg/sec, the stage pressure ratio and efficiency were 1.195 and 0.88, respectively. The design speed rotor peak efficiency of 0.91 occurred at the same flow rate.

Overall and blade element performance of a 1.20 pressure ratio fan stage with rotor blades reset -7 deg

NASA Technical Reports Server (NTRS)

Lewis, G. W., Jr.; Kovich, G.

1976-01-01

A 51-cm-diam model of a fan stage for short haul aircraft was tested in a single stage compressor research facility. The rotor blades were set 7 deg toward the axial direction (opened) from the design setting angle. Surveys of the air flow conditions ahead of the rotor, between the rotor and stator, and behind the stator were made over the stable operating range of the stage. At the design speed and a weight flow of 30.9 kg/sec, the stage pressure ratio and efficiency were 1.205 and 0.85, respectively. The design speed rotor peak efficiency of 0.90 occurred at a flow rate of 32.5 kg/sec.

Materials and structural aspects of advanced gas-turbine helicopter engines

NASA Technical Reports Server (NTRS)

Freche, J. C.; Acurio, J.

1979-01-01

Advances in materials, coatings, turbine cooling technology, structural and design concepts, and component-life prediction of helicopter gas-turbine-engine components are presented. Stationary parts including the inlet particle separator, the front frame, rotor tip seals, vanes and combustors and rotating components - compressor blades, disks, and turbine blades - are discussed. Advanced composite materials are considered for the front frame and compressor blades, prealloyed powder superalloys will increase strength and reduce costs of disks, the oxide dispersion strengthened alloys will have 100C higher use temperature in combustors and vanes than conventional superalloys, ceramics will provide the highest use temperature of 1400C for stator vanes and 1370C for turbine blades, and directionally solidified eutectics will afford up to 50C temperature advantage at turbine blade operating conditions. Coatings for surface protection at higher surface temperatures and design trends in turbine cooling technology are discussed. New analytical methods of life prediction such as strain gage partitioning for high temperature prediction, fatigue life, computerized prediction of oxidation resistance, and advanced techniques for estimating coating life are described.

Effect of Helicopter Blade Dynamics on Blade Aerodynamic and Structural Loads

NASA Technical Reports Server (NTRS)

Heffernan, Ruth M.

1987-01-01

The effect of rotor blade dynamics on aerodynamic and structural loads is examined for a conventional, main- rotor helicopter using both a comprehensive rotorcraft analysis (CAMRAD) and night test data. The impact of blade dynamics on blade section lift-coefficient time histories is studied by comparing predictions from both a rigid blade analysis and an elastic blade analysis with helicopter flight test data. The elastic blade analysis better predicts high-frequency behavior of section lift. In addition, components of the blade angle of attack, such as elastic blade twist, blade nap rate, blade slope velocity, and inflow, are examined as a function of blade mode. Elastic blade motion affects the blade angle of attack by a few tenths of a degree, and up to the sixth rotor harmonic. A similar study of the influence of blade dynamics on bending and torsion moments was also conducted. The modal analysis of the predicted blade structural loads suggested that five elastic bending deg of freedom (four flap and one lag) and three elastic torsion deg of freedom contributed to calculations of the blade structural loads. However, when structural bending load predictions from several elastic blade analyses were compared with flight test data, an elastic blade model consisting of only three elastic bending modes (first and second flap, and first lag), and two elastic torsion modes was found to be sufficient for maximum correlation.

Wave Engine Topping Cycle Assessment

NASA Technical Reports Server (NTRS)

Welch, Gerard E.

1996-01-01

The performance benefits derived by topping a gas turbine engine with a wave engine are assessed. The wave engine is a wave rotor that produces shaft power by exploiting gas dynamic energy exchange and flow turning. The wave engine is added to the baseline turboshaft engine while keeping high-pressure-turbine inlet conditions, compressor pressure ratio, engine mass flow rate, and cooling flow fractions fixed. Related work has focused on topping with pressure-exchangers (i.e., wave rotors that provide pressure gain with zero net shaft power output); however, more energy can be added to a wave-engine-topped cycle leading to greater engine specific-power-enhancement The energy addition occurs at a lower pressure in the wave-engine-topped cycle; thus the specific-fuel-consumption-enhancement effected by ideal wave engine topping is slightly lower than that effected by ideal pressure-exchanger topping. At a component level, however, flow turning affords the wave engine a degree-of-freedom relative to the pressure-exchanger that enables a more efficient match with the baseline engine. In some cases, therefore, the SFC-enhancement by wave engine topping is greater than that by pressure-exchanger topping. An ideal wave-rotor-characteristic is used to identify key wave engine design parameters and to contrast the wave engine and pressure-exchanger topping approaches. An aerodynamic design procedure is described in which wave engine design-point performance levels are computed using a one-dimensional wave rotor model. Wave engines using various wave cycles are considered including two-port cycles with on-rotor combustion (valved-combustors) and reverse-flow and through-flow four-port cycles with heat addition in conventional burners. A through-flow wave cycle design with symmetric blading is used to assess engine performance benefits. The wave-engine-topped turboshaft engine produces 16% more power than does a pressure-exchanger-topped engine under the specified topping constraints. Positive and negative aspects of wave engine topping in gas turbine engines are identified.

Simulations of wind turbine rotor with vortex generators

NASA Astrophysics Data System (ADS)

Troldborg, Niels; Zahle, Frederik; Sørensen, Niels N.

2016-09-01

This work presents simulations of the DTU 10MW wind turbine rotor equipped with vortex generators (VGs) on the inner part of the blades. The objective is to study the influence of different VG configurations on rotor performance and in particular to investigate the radial dependence of VGs, i.e. how VGs at one section of the blade may affect the aerodynamic characteristics at other radial positions. Furthermore, the performance of different sections on the blade is compared to their corresponding performance in 2D flow.

Development of Improved Design and 3D Printing Manufacture of Cross-Flow Fan Rotor

DTIC Science & Technology

2016-06-01

the design study, each solver run was monitored. Plotting the value of the mass flows, as well as the torque on the rotor blades , allowed a simple...DISTRIBUTION CODE A 13. ABSTRACT (maximum 200 words) This study determined the optimum blade stagger angle for a cross-flow fan rotor and evaluated the...parametric study determined optimum blade stagger angle using thrust, power, and thrust-to-power ratio as desired output variables. A MarkForged Mark One 3D

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.

Parametric Blade Study Test Report Rotor Configuration. Number 4

DTIC Science & Technology

1988-11-01

Figure 2. The rotor shaft is mounted on an oil-damped roller bearing at the forward location and a ball bearing at the aft location; radial runout does...thermodynamic properties. 22 d. Corrections were made to measured compressor temperatures and pressures, facility flowrate, and rotor wheel speed to...1152 .Z660 .1024 STRM- BLADE BLADE WHEEL LINE SECT. LEAN SPEED NUMBER ANGLE ANGLE 1 -55.15 7.32 1497.9 2 -53.85 8.09 1434.7 3 -52.96 7.11 1372.1 4

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.