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At least 1,099 records · Page 61
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Reynolds number effect on overall performance of a 10.8-centimeter (4.25-inch) sweptback bladed centrifugal compressor

The performance data were taken at 50,000 rpm, using argon gas. As the Reynolds number was reduced from near design value to 30 percent of design, the maximum efficiency decreased about 1.5 percentage points. Reducing the Reynolds number from 30 percent to approximately 10 percent of design caused the maximum efficiency to decrease another 2.5 percentage points. The variation in loss with Reynolds number is compared with inverse power relation of loss with Reynolds number.

Weigel, C.↗
NASA NTRS1972

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

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.

Brent, J. A.↗
NASA NTRS1972

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors. Part 2: Data and performance for stage A

Stage A, comprised of a conventional rotor and stator, was designed and tested to establish a performance baseline for comparison with the results of subsequent tests planned for two tandem-blade stages. 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 A achieved a maximum adiabatic efficiency of 85.1 percent at a pressure ratio of 1.29. The stage maximum adiabatic efficiency was 78.6 percent at a pressure ratio of 1.27.

Brent, J. A.↗
NASA NTRS1972

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

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.

Brent, J. A.↗
NASA NTRS1971

Unsteady flow through compressor stages

The application of the nonsteady Lax-Wendroff technique to problems with asymptotically periodic solution which offers a potentially powerful method for the investigation of the interaction of rotating and stationary blade rows in turbomachinery is reported. A technique for specifying boundary conditions with phase lag was developed to accomplish this. A complete nonlinear analysis is carried out numerically to determine the entire flow field without recourse to the assumption of small disturbances of linear equations which underlie the previous acoustic theories. The result, obtained for the case of equal number of rotor and stator blades shows that transonic flow can be handled without difficulty. In addition, the program is not limited with regard to blade thickness, camber or loading. Extension of this method to incorporate viscous wakes and to analysis of fully three dimensional configuration is feasible, and would greatly expand its utility in practical applications.

Alzner, E.↗
NASA NTRS1972

High-loading, 1800 ft/sec tip speed transonic compressor fan stage. 1: Aerodynamic and mechanical design

A single stage fan with a tip speed of 1800 ft/sec (548.6m/sec) and hub/tip ratio of 0.5 was designed to produce a pressure ratio of 2.285:1 with an adiabatic efficiency of 84.0%. The design flow per inlet annulus area is 38.7 lbm/sq ft-sec (188.9KG/sqm-sec). Rotor blades have modified multiple-circular-arc and precompression airfoil sections. The stator vanes have multiple-circular-arc airfoil sections.

Morris, A. L.↗
NASA NTRS1973

Digital slope-threshold data compressor

Slope-threshold compression scheme for telemetered video data is efficient, and its principle of operation is as follows: when slope of raw data exceeds threshold decision reference, previous sample is transmitted. All-digital design is more economical than analog system. It exhibits well-defined accuracy, provides unlimited storage time, and is convenient and reliable.

Anderson, T. O.↗
NASA NTRS1973

Transient analysis of energy Transfer Control (ECT) and compressor bleed concepts of remote lift fan control

The transient performance of two concepts for control of vertical takeoff aircraft remote lift fans is analyzed and discussed. Both concepts employ flow transfer between pairs of lift fans located in separate parts of the aircraft in order to obtain attitude control moments for hover and low-speed flight. The results presented are from a digital computer, dynamic analysis of the YJ97/LF460 remote drive turbofan. The transient responses of the two systems are presented for step demands in lift and moment.

Sellers, J. F.↗
NASA NTRS1973

Steady state analysis of Energy Transfer Control (ETC) and compressor bleed concepts of remote lift fan control

Two new concepts of control for VTOL aircraft remote lift fans are analyzed and discussed. Both concepts employ flow transfer among lift fans located in separate parts of the aircraft in order to obtain attitude control moments in hover and low-speed flight. The results presented are from a digital computer analysis of the YJ97/LF460 remote drive turbofan, although qualitative conclusions should apply to any remote fan system. The basic operating principles of the two systems are described, and their steady-state moment generating capabilities are discussed. Lift-recovery strategies for component failures are also analyzed.

Sellers, J. F.↗
NASA NTRS1973

Two-dimensional analytical and experimental performance comparison for a compressor stator section with D-factor of 0.47

Analytically computed flow parameters were compared to measured values for the midspan double-circular-arc section of a stator in subsonic flow. Analytical procedures included calculations for inviscid flow, blade surface boundary layers, and loss coefficients. Comparisons were made at three incidence angles. Methods for prescribing the exit fluid angle in inviscid flow calculations were investigated. Measured loss coefficients were compared to calculated values. Two methods for predicting performance involving iterative use of ideal flow and boundary layer calculations were investigated.

Sanger, N. L.↗
NASA NTRS1974

Estimation of deviation angle for axial-flow compressor blade sections using inviscid-flow solutions

Development of a method of estimating deviation angles by analytical procedures was begun. Solutions for inviscid, irrotational flow in the blade-to-blade plane were obtained with a finite-difference calculation method. Deviation angles for a plane cascade with a rounded trailing edge were estimated by using the inviscid-flow solutions and three trailing-edge hypotheses. The estimated deviation angles were compared with existing experimental data over a range of incidence angles at inlet flow angles of 30 deg and 60 deg. The results indicate that deviation angles can be estimated accurately (within 1 deg) by using one of the three trailing-edge hypotheses, but only when pressure losses are low. A new trailing-edge hypotheses is presented which is suitable (for the cascade considered) for both low- and high-loss operating points.

Miller, M. J.↗
NASA NTRS1974

Application of holography to flow visualization within rotating compressor blade row

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

Wuerker, R. F.↗
NASA NTRS1974

Performance of a transonic compressor rotor with an aspect ratio of 6.5

The overall and blade-element performances and the aerodynamic design parameters of a transonic rotor with an aspect ratio of 6.5, designed to investigate the effects of aspect ratio on range and performance are presented. The rotor was designed for a total-pressure ratio of 1.53, an efficiency of 0.898, and a weight flow of 67.76 pounds per second.

Reid, L.↗
NASA NTRS1974

Deviation angle estimation for axial-flow compressors using inviscid flow solutions

Development of a more satisfactory method of estimating deviation angles was attempted using analytical procedures. Solutions for inviscid, irrotational flow in the blade-to-blade plane were obtained with a finite-difference calculation method. Deviation angles for three cascades with rounded trailing edges were estimated using the inviscid flow solutions and four existing trailing edge hypotheses. The estimated deviation angles were compared with experimental values for a range of incidence angle and inlet flow angle. None of the hypotheses were satisfactory at all incidence angles for all the cascades considered. An alternative hypothesis is presented which is suitable for one of the cascade profiles over a range of incidence and blade setting angle.

Miller, M. J.↗