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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1979

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

Compressor rotor strain gage data from an engine test conducted with an inlet screen distortion were reduced and analyzed. The data were compared to results 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 effect of the distortion was found to be most prominent in the first three compressor stages, with the rotor speed establishing the period of the complex wave and, consequently, the frequencies of all the higher engine order excitations. At certain speeds it was observed that the complex pressure wave had the frequency content to excite a number of modes simultaneously, although the overall magnitudes were small and well within allowable stress limits.

Lubomski, J. F.↗
NASA NTRS1979

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

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.

Lubomski, J. F.↗
NASA NTRS1978

Perturbation solutions for transonic flow on the blade-to-blade surface of compressor blade rows

A preliminary investigation was conducted to establish the theoretical basis of perturbation techniques with the objective of minimizing computational requirements associated with parametric studies of transonic flows in turbomachines. The theoretical analysis involved the development of perturbation methods for determining first order changes in the flow solution due to variations of one or more geometrical or flow parameters. The formulation is primarily directed toward transonic flows on the blade to blade surface of a single blade row compressor. Two different perturbation approaches were identified and studied. Applications and results of these methods for various perturbations are presented for selected two dimensional transonic cascade flows to illustrate the advantages and disadvantages of each technique. Additionally, it was found that, for flows with shock waves, proper account of shock displacement was crucial.

Stahara, S. S.↗
NASA NTRS1974

Lifting surface theory of axial compressor blade rows. I - Subsonic compressor. II - Transonic compressor

A linearized three-dimensional lifting surface theory is considered, taking into account relations to axisymmetric 'through-flow' or actuator disk theories, general results for flow within the blade row, relations to two-dimensional cascade theory, and finite blade and wake effects. Numerical results are presented, giving attention to the problem of wakeless flow and of wake flow. Expressions for wake functions are given in an appendix. The lifting line velocity potential employed by Namba (1972) is used as a starting point in the discussion of the lifting surface theory. A study of subsonic conditions is supplemented by an investigation of a transonic compresso, taking into account the general three-dimensional solution, induced velocities, and the acoustic pressure field.

Okurounmu, O.↗
NASA NTRS1956

Low-speed cascade investigation of loaded leading-edge compressor blades

Six percent thick NACA 63-series compressor-blade sections having a loaded leading-edge A4K6 mean line have been investigated systematically in a two-dimensional porous-wall cascade over a range of Reynolds numbers from 160,000 to 385,000. Blades cambered to have isolated-airfoil lift coefficients of 0.6, 1.2, 1.8, and 2.4 were tested over the usable angle-of-attack range at inlet-air angles of 30 degrees, 45 degrees, and 60 degrees and solidities of 1.0 and 1.5. A comparison with data of NACA RM L51G31, shows that the angle-of-attack operating range is 2 degrees to 4 degrees less than the range for the uniformly loaded section; however, the wake losses near design angle of attack are slightly lower than those for the uniformly loaded section. Except for highly cambered blades at high inlet angles, the 63-(C s oA4K6)06 compressor-blade sections are capable of more efficient operation for moderate-speed subsonic compressors at design angle of attack than are the 65-(C s oa10)10 or the 65-(c s oA2I8b)10 compressor-blade sections. In contrast to the other sections, the loaded leading-edge sections are capable of operating efficiently at the lower Reynolds numbers.

Emery, James C↗
NASA NTRS1982

The use of optimization techniques to design controlled diffusion compressor blading

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

Sanger, N. L.↗
NASA NTRS1987

Shot peening for Ti-6Al-4V alloy compressor blades

A text program was conducted to determine the effects of certain shot-peening parameters on the fatigue life of the Ti-6Al-4V alloys as well as the effect of a demarcation line on a test specimen. This demarcation line, caused by an abrupt change from untreated surface to shot-peened surface, was thought to have caused the failure of several blades in a multistage compressor at the NASA Lewis Research Center. The demarcation line had no detrimental effect upon bending fatigue specimens tested at room temperature. Procedures for shot peening Ti-6Al-4V compressor blades are recommended for future applications.

Carek, Gerald A.↗
NASA NTRS1990

Stacking optimization of compressor blades of gas turbine engines

A procedure is presented to obtain optimal designs of axial compressor blades with structural design constraints. Coefficients of the polynomials defining the circumferential tilting angles and the axial leaning distances of the airfoil cross sections from the initial design geometry are used as design variables. The compressor blades are modeled by 20-node solid elements. An efficient finite element method is developed for modal analysis and sensitivity analysis with respect to the design variables. Based on this information, a sequential linear programming method is applied to calculate the required change of geometry for the desired structural design constraints.

Cheu, Tsu-Chien↗
NASA NTRS1988

Vortex shedding in high-speed compressor blade wakes

The wakes of highly loaded compressor blades are generally considered to be turbulent flows. Recent work has suggested that the blade wakes are dominated by a vortex streetlike structure. The experimental evidence supporting the wake vortex structure is reviewed. This structure is shown to redistribute thermal energy within the flowfield. The effect of the wake structure on conventional aerodynamic measurements of compressor performance is noted. A two-dimensional, time-accurate, viscous numerical simulation of the flow exhibits both vortex shedding in the wake and a lower-frequency flow instability that modulates the shedding. The numerical results are shown to agree quite well with the measurement from transonic compressor rotors.

Epstein, A. H.↗
NASA NTRS1957

Comparison of NACA 65-series compressor-blade pressure distributions and performance in a rotor and in cascade

An investigation has been conducted to compare the performance of NACA 65-series compressor blades in two-dimensional cascade with that in an axial flow compressor. Blade pressure distributions were obtained by the use of a mercury-seal pressure-transfer device. The comparison indicated that cascade data accurately predicted the turning angle and blade pressure distribution obtained in the compressor at design conditions.

CASCADES↗
NASA NTRS1957

Identification of Foreign Objects Damaging Compressor Blades in Turbojet Engines

Damage to the compressor blades of turbojet engines due to ingestion of foreign objects is a growing problem, the solution of which has been made more difficult by the large percentage of damaging materials that have remained unknown. A rapid emission spectroscopic method was devised to identify the chemical composition of these foreign objects. Results on laboratory-prepared specimens and blades from damaged engines show that the method can be utilized to determine the nature of the ingested foreign objects.

Spakowski, A E↗
NASA NTRS1987

Vortex shedding in compressor blade wakes

The wakes of highly loaded axial compressor blades were often considered to be turbulent, unstructured flows. Recent work has suggested that the blade wakes are in fact dominated by a vortex street-like structure. The work on the wake structure at MIT is reviewed, the results of a viscous numerical simulation are presented, the blade wake vortices are compared to those shed from a cylinder, and the implications of the wake structure on compressor performance are discussed. In particular, a two-dimensional, time accurate, viscous calculation shows both a periodic wake structure and time variations in the passage shock strength. The numerical calculations are compared to laser anemometer and high frequency response probe data. The effect of the wake structure on the entropy production and apparent adiabatic efficiency of the compressor rotor is discussed.

Epstein, A. H.↗
NASA NTRS1951

Systematic Two-Dimensional Cascade Tests of NACA 65-Series Compressor Blades at Low Speeds

The performance of NACA 65-series compressor blade section in cascade has been investigated systematically in a low-speed cascade tunnel. Porous test-section side walls and for high-pressure-rise conditions, porous flexible end walls were employed to establish conditions closely simulating two-dimensional flow. Blade sections of design lift coefficients from 0 to 2.7 were tested over the usable angle-of-attack range for various combinations of inlet-flow angle. A sufficient number of combinations were tested to permit interpolation and extrapolation of the data to all conditions within the usual range of application. The results of this investigation indicate a continuous variation of blade-section performance as the major cascade parameters, blade camber, inlet angle, and solidity were varied over the test range. Summary curves of the results have been prepared to enable compressor designers to select the proper blade camber and angle of attack when the compressor velocity diagram and desired solidity have been determined.

Herrig, L. Joseph↗
NASA NTRS1973

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

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

Sanger, N. L.↗
NASA NTRS1993

Data On Flow About A Compressor Blade In Cascade

Development and testing of software to compute values of flow-field characteristics greatly enhanced by using recorded experimental data to test results of computations. Report "The Measurement Of Boundary Layers On Compressor Blade In Cascade" is data base providing detailed measurements of boundary layer and wake in flow field about double-circular-arc compressor blades in cascade of such blades.

Deutsch, S.↗