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At least 595 records · Page 33
All sourcesNASA NTRSDOE OSTI
NASA NTRS1983

Helicopter blade tips

Methods of improving helicopter performance and vibration level by proper shaping of helicopter blade tips are considered. The principle involved consists of reducing the extent of the supersonic zone above the advancing tip and of the turbulent interaction. For stationary and advancing flight, the influence of the rotor and the problems posed by blade tips are reviewed. The theoretical methods of dealing with the two types of flight are briefly stated, and the experimental apparatus is described, including model triple and quadruple rotors. Different blade tip shapes are shown and briefly discussed. The theoretical results include an advancing speed of 309 km/H and a blade tip rotational speed of 215 m/s. The experimental values are advancing speed of 302 km/h and blade tip Mach number 0.86 for both types of rotors.

Lyothier, R.↗
NASA NTRS1983

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

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

Fontana, R. R.↗
NASA NTRS1984

Stagger angle dependence of inertial and elastic coupling in bladed disks

Conditions which necessitate the inclusion of disk and shaft flexibility in the analysis of blade response in rotating blade-disk-shaft systems are derived in terms of nondimensional parameters. A simple semianalytical Rayleigh-Ritz model is derived in which the disk possesses all six rigid body degrees of freedom, which are elastically constrained by the shaft. Inertial coupling by the rigid body motion of the disk on a flexible shaft and out-of-plane elastic coupling due to disk flexure are included. Frequency ratios and mass ratios, which depend on the stagger angle, are determined for three typical rotors: a first stage high-pressure core compressor, a high bypass ratio fan, and an advanced turboprop. The stagger angle controls the degree of coupling in the blade-disk system. In the blade-disk-shaft system, the stagger angle determines whether blade-disk motion couples principally to the out-of-plane or in-plane motion of the disk on the shaft. The Ritz analysis shows excellent agreement with experimental results.

Crawley, E. F.↗
NASA NTRS1984

Flap-lag damping of an elastic rotor blade with torsion and dynamic inflow in hover from symbolically generated equations

A program written in FORTRAN IV and coded into subroutines for subsequent numerical study is used to derive symbolically the governing equations of motion of the elastic rotor blade and the dynamic inflow equations. By rearranging the coefficients of the equations, the blade and dynamic inflow equations are converted into equations in a multiblade coordinated system. The final multiblade equations can accommodate any number of elastic blade modes. It is noted that with a given data set, the entire process, from the derivation of the equations to the numerical calculations, is computerized and requires only limited user interface. To demonstrate the usefulness of the approach, damping data in hover with and without dynamic inflow effects are presented for various rotor blade models, including flap-lag, flap-lag-torsion, quasi-static torsion, and rigid blade lag-flap.

Reddy, T. S. R.↗
NASA NTRS1986

Photo-Optical Blade-Vibration-Data Acquisition System

Techniques developed for measuring blade vibrations using optical sensors fixed in casing around bladed turbine-engine stage. Measurement system coordinates and assembles inputs of many optical probes in order to monitor all of blades on spinning stage. Several points on every blade monitored in order to develop quantitative measure of overall bladed-stage response.

Kiraly, L. J.↗
NASA NTRS1985

Vacuum plasma coatings for turbine blades

Turbine blades, vacuum plasma spray coated with NiCrAlY, CoCrAlY or NiCrAlY/Cr2O3, were evaluated and rated superior to standard space shuttle main engine (SSME) coated blades. Ratings were based primarily on 25 thermal cycles in the MSFC Burner Rig Tester, cycling between 1700 F (gaseous H2) and -423 F (liquid H2). These tests showed no spalling on blades with improved vacuum plasma coatings, while standard blades spalled. Thermal barrier coatings of ZrO2, while superior to standard coatings, lacked the overall performance desired. Fatigue and tensile specimens, machined from MAR-M-246(Hf) test bars identical to the blades were vacuum plasma spray coated, diffusion bond treated, and tested to qualify the vacuum plasma spray process for flight hardware testing and application. While NiCrAlY/Cr2O3 offers significant improvement over standard coatings in durability and thermal protection, studies continue with an objective to develop coatings offering even greater improvements.

Holmes, R. R.↗
NASA NTRS1986

Individual-blade-control research in the MIT VTOL Technology Laboratory 1977-1985

A new, advanced system for active control of helicopters and its application to the solution of rotor aerodynamic and aeroelastic problems is described. Each blade is individually controlled in the rotating frame over a wide range of frequencies. Application of the system to gust alleviation, attitude stabilization, vibration alleviation, blade lag damping augmentation, stall flutter suppression, blade flapping stabilization, stall alleviation, and performance enhancement is outlined. The effectiveness of the system in achieving most of these applications is demonstrated by experimental results from wind tunnel tests of a model helicopter rotor with individual blade control. The feasibility of achieving many or all of the applications of individual blade control using the conventional helicopter swash plate is demonstrated, and the necessary control laws are presented.

Ham, N. D.↗
NASA NTRS1985

The effect of limiting aerodynamic and structural coupling in models of mistuned bladed disk vibration

A model has been developed for studying the effect of mistuning on bladed disk vibration which has the unique feature that the extent of aerodynamic and structural interaction which it simulates can be readily varied from full coupling of all blades on the disk to coupling of each blade with only its nearest neighbors. Simulations utilizing the resulting algorithm show that limited coupling models may be used to predict the statistical distribution of blade amplitudes that characterizes the mistuning effect, which in turn determines stage durability. This approach is used to study the effect of changing various system parameters on amplitude scatter. Gas density, the number of blades on the disk, disk stiffness, and the engine order of the excitation are considered. The results are used to draw some conclusions about how to improve laboratory tests and component design.

Basu, P.↗
NASA NTRS1986

Surface pressure measurement during low speed two-dimensional blade-vortex interaction

Unsteady blade surface pressures were measured during two-dimensional blade-vortex interaction. A periodically generated vortex filament interacted with a blade model instrumented over the first 20 percent of the blade surface. Analysis of data confirm the leading edge nature of BVI and show the amplitude of the interaction to be a strong function of blade-to-vortex spacing. Lift perturbation time histories were computed and are used to examine details of the encounter process.

Booth, E. R., Jr.↗
NASA NTRS1984

STAEBL: Structural tailoring of engine blades, phase 2

The Structural Tailoring of Engine Blades (STAEBL) program was initiated at NASA Lewis Research Center in 1980 to introduce optimal structural tailoring into the design process for aircraft gas turbine engine blades. The standard procedure for blade design is highly iterative with the engineer directly providing most of the decisions that control the design process. The goal of the STAEBL program has been to develop an automated approach to generate structurally optimal blade designs. The program has evolved as a three-phase effort with the developmental work being performed contractually by Pratt & Whitney Aircraft. Phase 1 was intended as a proof of concept in which two fan blades were structurally tailored to meet a full set of structural design constraints while minimizing DOC+I (direct operating cost plus interest) for a representative aircraft. This phase was successfully completed and was reported in reference 1 and 2. Phase 2 has recently been completed and is the basis for this discussion. During this phase, three tasks were accomplished: (1) a nonproprietary structural tailoring computer code was developed; (2) a dedicated approximate finite-element analysis was developed; and (3) an approximate large-deflection analysis was developed to assess local foreign object damage. Phase 3 is just beginning and is designed to incorporated aerodynamic analyses directly into the structural tailoring system in order to relax current geometric constraints.

Hirschbein, M. S.↗
NASA NTRS1986

Application of higher harmonic blade feathering on the OH-6A helicopter for vibration reduction

The design, implementation, and flight test results of higher harmonic blade feathering for vibration reduction on the OH-6A helicopter are described. The higher harmonic control (HHC) system superimposes fourth harmonic inputs upon the stationary swashplate. These inputs are transformed into 3P, 4P and 5P blade feathering angles. This results in modified blade loads and reduced fuselage vibrations. The primary elements of this adaptive vibration suppression system are: (1) acceleration transducers sensing the vibratory response of the fuselage; (2) a higher harmonic blade pitch actuator system; (3) a flightworthy microcomputer, incorporating the algorithm for reducing vibrations, and (4) a signal conditioning system, interfacing between the sensors, the microcomputer and the HHC actuators. The program consisted of three distinct phases. First, the HHC system was designed and implemented on the MDHC OH-6A helicopter. Then, the open loop, or manual controlled, flight tests were performed, and finally, the closed loop adaptive control system was tested. In 1983, one portion of the closed loop testing was performed, and in 1984, additional closed loop tests were conducted with improved software. With the HHC system engaged, the 4P pilot seat vibration levels were significantly lower than the baseline ON-6A levels. Moreover, the system did not adversely affect blade loads or helicopter performance. In conclusion, this successful proof of concept project demonstrated HHC to be a viable vibration suppression mechanism.

Straub, F. K.↗
NASA NTRS1986

Oxide-dispersion-strengthened turbine blades, volume 1

The objective of Project 4 was to develop a high-temperature, uncooled gas turbine blade using MA6000 alloy. The program objectives were achieved. Production scale up of the MA6000 alloy was achieved with a fair degree of tolerance to nonoptimum processing. The blade manufacturing process was also optimized. The mechanical, environmental, and physical property evaluations of MA6000 were conducted. The ultimate tensile strength, to about 704 C (130 F), is higher than DS MAR-M 247 but with a corresponding lower tensile elongation. Also, above 982 C (180 F) MA6000 tensile strength does not decrease as rapidly as MAR-M 247 because the ODS mechanism still remains active. Based on oxidation resistance and diffusional stability considerations, NiCrAlY coatings are recommended. CoCrAly coating should be applied on top of a thin NiCrAlY coating. Vibration tests, whirlpit tests, and a high-rotor-rig test were conducted to ensure successful completion of the engine test of the MA6000 TFE731 high pressure turbine blades. The results of these tests were acceptable. In production quantities, the cost of the Project 4 MA6000 blade is estimated to be about twice that of a cast DS MAR-M 247 blade.

Millan, P. P., Jr.↗
NASA NTRS1987

SSME blade damper technology

Before 1975 turbine blade damper designs were based on experience and very simple mathematical models. Failure of the dampers to perform as expected showed the need to gain a better understanding of the physical mechanism of friction dampers. Over the last 10 years research on friction dampers for aeronautical propulsion systems has resulted in methods to optimize damper designs. The first-stage turbine blades on the Space Shuttle Main Engine (SSME) high-pressure oxygen pump have experienced cracking problems due to excessive vibration. A solution is to incorporate a well-designed friction dampers to attenuate blade vibration. The subject study, a cooperative effort between NASA Lewis and Carnegie-Mellon University, represents an application of recently developed friction damper technology to the SSME high-pressure oxygen turbopump. The major emphasis was the contractor's design known as the two-piece damper. Damping occurs at the frictional interface between the top half of the damper and the underside of the platforms of the adjacent blades. The lower half of the damper is an air seal to retard airflow in the volume between blade necks.

Kielb, Robert E.↗
NASA NTRS1987

The spectral characteristics of rotor blade-vortex interaction noise - Experimental and mathematical results

The BVI impulsive content of a rotor acoustic signal is shown to appear in the mid-frequency range of the power spectrum, between the fifth and thirtieth harmonics of the blade passage frequency, concentrated at the harmonics of the blade passage frequency. These harmonics exhibit a humped or scalloped shape in this mid-frequency spectral region. Increased energy at the harmonics of the shaft frequency appears when the BVI impulsive content demonstrates unsteadiness and blade-to-blade differences in the time domain. A mathematical model of a generalized BVI acoustic signal and its power spectrum shows that the power spectrum is scalloped and filtered by a comb function. The spectrum amplitude is defined by the impulse amplitude and emission time. The scalloping of the spectrum is related to the emission time of the impulse itself, and the spacing of the comb function is related to the repetition time (period) of the impulse. The decay rate of the spectral humps is governed by the inverse of frequency squared. The mathematical model validates the characteristics observed in the data and verify that these characteristics are due to blade-vortex interaction activity.

Martin, Ruth M.↗
NASA NTRS1986

Prediction of blade stresses due to gust loading

An analysis is developed for investigating the response of a rotor-fuselage system in a three-dimensional gust field wherein the gust velocity conponents can have arbitrary variation in space and time. Each rotor blade undergoes flap bending, lag bending and torsional deflections. The blades are divided into beam elements and each element consists of fifteen nodal degrees of freedom. Quasi-steady strip theory is used to obtain the aerodynamic loads. Unsteady aerodynamic effects are introduced through dynamic inflow modeling. Dynamic stall and reverse flow effects are also included. The fuselage is allowed five degrees of freedom: vertical, longitudinal, lateral, pitch and roll motions. The gust response equations are linearized about the vehicle trim state and the blade steady-state deflected position, and then solved by time integration. The blade bending moments, which determine blade stresses, are evaluated using the force summation technique. Systematic studies are made to identify the importance of several parameters including dynamic stall, forward speed, lag stiffness, gust profile, gust penetration rate and gust velocity direction.

Bir, Gunjit↗
NASA NTRS1987

Dynamic characteristics of the national full-scale aerodynamic complex drive fan blades

An extensive experimental test program was conducted to define the dynamic characteristics of the newly installed fan blades of the NASA Ames 40- by 80-/80- by 120-Foot Wind Tunnel. The tests were intended to substantiate the blade design and manufacturing process, verify installation integrity, and guarantee safe fan-drive operation. The program included a shake test of each fan blade after installation and before operation as well as the monitoring of blade strain levels during initial operation. Structural characteristics, as determined by the shake test, include the modal frequency and damping values for elastic bending and torsion modes between 5 and 130 Hz. Operational testing defined the maximum and oscillatory blade-bending strains throughout the drive-system operational envelope. Results are presented for the shake test and initial operation of the 40- by 80-Foot Wind Tunnel circuit.

Peterson, Randall L.↗
NASA NTRS1986

Helicopter individual-blade-control research at MIT 1977-1985

A new advanced system for active control of helicopters and its application to the solution of rotor aerodynamic and aeroelastic problems is described. Each blade is individually controlled in the rotating frame over a wide range of frequencies. Application of the system to gust alleviation, attitude stabilization, vibration alleviation, blade lag damping augmentation, stall flutter suppression, blade flapping stabilization, stall alleviation, and performance enhancement is outlined. The effectiveness of the system in achieving most of these applications is demonstrated by experimental results from wind tunnel tests of a model helicopter rotor with individual blade control. The feasibility of achieving many or all of the applications of individual blade control using the conventional helicopter swash plate is demonstrated, and the necessary control laws are presented.

Ham, Norman D.↗
NASA NTRS1988

Unsteady heat transfer in turbine blade ducts: Focus on combustor sources

Thermal waves generated by either turbine rotor blades cutting through nonuniform combustor temperature fields or unsteady burning could lead to thermal fatigue cracking in the blades. To determine the magnitude of the thermal oscillation in blades with complex shapes and material compositions, a finite element Galerkin formulation has been developed to study combustor generated thermal wave propagation in a model two-dimensional duct with a uniform plug flow profile. The reflection and transmission of the thermal waves at the entrance and exit boundaries are determined by coupling the finite element solutions at the entrance and exit to the eigenfunctions of an infinitely long adiabatic duct. Example solutions are presented. In general, thermal wave propagation from an air passage into a metallic blade wall is small and not a problem. However, if a thermal barrier coating is applied to a metallic surface under conditions of a high heat transfer, a good impedance match is obtained and a significant portion of the thermal wave can pass into the blade material.

Baumeister, Kenneth J.↗