Method of Determining Centrifugal-flow-compressor Performance with Water Injection
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Hamrick, Joseph T↗
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Altitude-performance and Reynolds number investigation of centrifugal-flow-compressor turbojet engine
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Wilsted, H D↗
Miniature high speed compressor having embedded permanent magnet motor
A high speed centrifugal compressor for compressing fluids includes a permanent magnet synchronous motor (PMSM) having a hollow shaft, the being supported on its ends by ball bearing supports. A permanent magnet core is embedded inside the shaft. A stator with a winding is located radially outward of the shaft. The PMSM includes a rotor including at least one impeller secured to the shaft or integrated with the shaft as a single piece. The rotor is a high rigidity rotor providing a bending mode speed of at least 100,000 RPM which advantageously permits implementation of relatively low-cost ball bearing supports.
Zhou, Lei↗
NASA HECC Geometry and Performance Review Part 1: Validation of a Computational Model for the Vaneless Diffuser Configuration with As-Manufactured Impeller Geometry
An investigation of the NASA High Efficiency Centrifugal Compressor (HECC) vaneless diffuser configuration was performed. This multipart investigation presents an extensive validation of the computational model against experimental data. The validated model was then used to explore the effects geometric differences between the as-manufactured and design-intent impellers. Lastly, the validated model was used for comparison against an experimental tip clearance study. Part I of the investigation focused on describing the HECC vaneless computational model and its validation against the experimental data published in GT2023-103128. To generate the model, a verification of the manufactured impeller was performed and was confirmed to match the impeller CAD. Section data from the impeller CAD, referred to as the As-Manufactured geometry, were used for development of the computational model. Steady RANS simulations utilizing the Wilcox K-Omega 98 turbulence model and a multiblock structured mesh were used to conduct a mesh convergence study. The resultant mesh was then used for simulations at both design-speed and off-design speed conditions. The simulations compared well to data in both 1D and 2D terms for the major performance metrics of efficiency, total pressure rise, and total temperature rise. Based on the observed results the model was deemed validated and was used in the latter two parts of the multipart investigation.
centrifugal compressor↗
NASA HECC Geometry and Performance Review Part 2: Geometric Differences Between the as-Manufactured and Design-Intent Impeller Geometry and Their Effects on the Vaneless Diffuser Configuration Performance
An investigation of the NASA High Efficiency Centrifugal Compressor (HECC) vaneless diffuser configuration was performed. This multipart investigation focused on validating the computational model against experimental data. The validated model was then used to explore the effects of observed geometric differences between the as-manufactured and design-intent impellers. Lastly, the validated model was used for comparison against an experimental tip clearance study. Details regarding each of the parts are detailed below: This second part of a multipart investigation of the NASA HECC vaneless diffuser configuration focused on the comparison of the as-manufactured HECC vaneless computational model validated in Part I against a new model that used the same grid and solver settings but replaced the impeller section data with the design-intent sections from NASA/CR-2014-218114/Rev1. The 1D performance values showed that the as-manufactured impeller underperformed compared to the design-intent for all four speedlines that were simulated. Further investigation of radial profiles, loading and contours showed consistent underperformance of the as-manufactured. An investigation into the source of the difference led to the discovery of several geometric differences in the two impellers of magnitudes that could not be explained outside of manufacturing. The differences that were observed can be summarized into four categories: fillets, TE exit radius differences, splitter LE differences and main blade thickness differences. The investigation then shifted to attempting to isolate the geometric differences as best as possible and to understand their contribution to the observed underperformance. The investigation showed that out of the four geometric differences that were explored, the differences in the splitter blade had the largest effect.
centrifugal compressor↗
NASA HECC Geometry and Performance Review Part 2: Geometric Differences Between the As-Manufactured and Design-Intent Impeller Geometry and their Effects on the Vaneless Diffuser Configuration Performance
An investigation of the NASA High Efficiency Centrifugal Compressor (HECC) vaneless diffuser configuration was performed. This multipart investigation focused on validating the computational model against experimental data. The validated model was then used to explore the effects of observed geometric differences between the As-Manufactured and Design-Intent impellers. Lastly, the validated model was used for comparison against an experimental tip clearance study. Part II of the investigation of the NASA HECC vaneless diffuser configuration focused on understanding the differences in geometry and performance between the As-Manufactured impeller and the Design-Intent impeller. This was achieved through comparison of the As-Manufactured HECC vaneless computational model, developed from the solid model of the impeller and validated in Part I, against a new model developed from the Design-Intent blade sections published in NASA/CR-2014-218114/Rev1. The 1D performance values showed that the As-Manufactured impeller underperformed compared to the Design-Intent for all four speedlines that were simulated. Further investigation of profiles, loading and contours showed consistent underperformance of the As-Manufactured relative to the Design-Intent. An investigation into the source of the performance differences led to the discovery of several significant variations in the impeller geometries. The differences that were observed can be summarized into four categories: fillets, trailing-edge exit radii variation, dissimilarity in the splitter leading-edge geometry and main blade thickness differences. The investigation then shifted its effort to quantify the performance effects caused by geometric differences. The investigation showed that out of the four geometric differences that were explored, the differences in the splitter blade had the largest impact. These differences most notably changed flow physics near the splitter leading edge, especially near the tip of the impeller.
centrifugal compressor↗
Internal hysteresis experienced on a high pressure syn gas compressor
A vibration instability phenomenon experienced in operating high pressure syn gas centrifugal compressors in two ammonia plants is described. The compressors were monitored by orbit and spectrum analysis for changes from baseline readings. It is found that internal hysteresis was the major destabilizing force; however, the problem was further complicated by seal lockup at the suction end of the compressor. A coupling lockup problem and a coupling fit problem, which frettage of the shaft, are also considered as contributors to the self excited vibrations.
Zeidan, F. Y.↗
A compressor designed for the energy research and development agency automotive gas turbine program
A centrifugal compressor was designed for a gas turbine powered automobile as part of the Energy Research and Development Agency program to demonstrate emissions characteristics that meet 1978 standards with fuel economy and acceleration which are competitive with conventionally powered vehicles. A backswept impeller was designed for the compressor in order to attain the efficiency goal range required for the objectives of this program. Details of the design and method of flow analysis of the compressor are presented.
Galvas, M. R.↗
Small, high pressure ratio compressor: Aerodynamic and mechanical design
The Small, High-Pressure-Ratio Compressor Program was directed toward the analysis, design, and fabrication of a centrifugal compressor providing a 6:1 pressure ratio and an airflow rate of 2.0 pounds per second. The program consists of preliminary design, detailed areodynamic design, mechanical design, and mechanical acceptance tests. The preliminary design evaluate radial- and backward-curved blades, tandem bladed impellers, impeller-and diffuser-passage boundary-layer control, and vane, pipe, and multiple-stage diffusers. Based on this evaluation, a configuration was selected for detailed aerodynamic and mechanical design. Mechanical acceptance test was performed to demonstrate that mechanical design objectives of the research package were met.
Bryce, C. A.↗
Performance of Compressor of XJ-41-V Turbojet Engine V-Performance Analysis of Compressor with Revised Vaned Collector over Range of Compressor Speeds from 3600 to 11,500 RPM
An investigation of the XJ-41-V turbojet-engine compressor with a revised vaned collector was conducted to determine the performance of the compressor and to obtain fundamental information on the aerodynamic problems associated with large centrifugal compressors of this type. The original vaned collector was revised by increasing the flow area at the vaned collector entrance. A maximum adiabatic efficiency of 0.81 was obtained et a corrected weight flow of 36.5 pounds per second and a pressure ratio of 1.90. The peak pressure ratio was 3.93 and occurred at an impeller speed of 11,500 rpm at a corrected weight flow of 65.5 pounds per second. Revision of the vaned collector resulted in an increased airflow capacity over the speed range. The design air-flow capacity of 78 pounds per second was very nearly reached at the engine design speed of 11,500 rpm. The compressor air-flow choking point occurred in the vaned collector passage; however, at speeds above 8300 rpm, the air-flow capacity of the impeller was being approached as indicated by large pressure losses in the impeller at maximum air-flow conditions. An increase in compressor air-flow capacity at the higher speeds can possibly be obtained 5y removal of the flow restriction in the impeller, which would result in an increased air density at the vaned collector entrance.
Ginsburg, Ambrose↗
Fluid mechanics, acoustics, and design of turbomachinery, part 1
A conference was conducted to analyze the effects of air flow on turbomachinery design. The subjects discussed are: (1) equations for compressible flow through turbomachines, (2) influence of axial velocity ratio on cascade performance, (3) three dimensional flow in transonic axial compressor blade rows, (4) prediction of turbulent shear layers in turbomachines, and (5) boundary layers in centrifugal compressors.
Lakshminarayana, B.↗
Performance of J-33-A-21 Turbojet-Engine Compressor I - Over-All Performance Characteristics at Equivalent Impeller Speeds from 6000 to 13,400 RPM
The NACA is investigating a series of J-33 turbojet-engine compressors to determine the over-all and component performances and to improve theories of flow through large centrifugal compressors, The production model J-33-A-21 was operated over a range of inlet temperatures from 80 to -40 F and inlet pressures from 14 to 5 inches mercury absolute for equivalent impeller speeds from 6000 to 13,400 rpm. At the equivalent design speed of 11,500 rpm, the compressor had a peak pressure ratio of 3.98 at an equivalent weight flow of 73.4 pounds per second and an adiabatic temperature-rise , efficiency of 0.701. When the compressor speed was reduced from the design speed to 6000 rpm, the adiabatic temperature-rise efficiency increased to 0.747. At the maximum equivalent speed investigated (13,400 rpm), a peak pressure ratio of 5.09 was obtained at an adiabatic temperature-rise efficiency of 0.617 and an equivalent weight flow of 66.O pounds per second. An increase in inlet pressure from 5.5 to 14 inches mercury absolute, with a consequent increase in Reynolds number index, improved the pressure ratio but had no apparent effect on the ratio of temperature rise through the compressor to inlet temperature. The variation of the peak adiabatic temperature-rise efficiency with inlet pressure is in the direction that would be expected from a Reynolds number effect. Decrease in the inlet temperature from 80 to -40 F, with a consequent increase in Reynolds number index, resulted in scatter of the pressure-ratio data and increased values of temperature ratio. The variation of the adiabatic temperature-rise efficiency with inlet temperature is probably the result of heat-transfer effects and scatter in the pressure ratio.
Beede, William L.↗
Investigation of X24C-2 10-Stage Axial-Flow Compressor. III - Surge Characteristics
Compressor operation at low air flows for a given speed is limited by unstable flow conditions, commonly called surge. An investigation of surge in centrifugal compressors (reference 1) showed that the pulsation of pressures and velocities occurred when the slope of the compressor characteristic curve was positive and that the magnitude and frequency, as well as the incidence of surge, depended on the capacity and resistance of the total system. Although the theory presented in reference 1 is applicable to axial-floe compressors, little experimental information is available on the surge characteristics of the individual stages of axial-flow compressors, or on the variation of the surge characteristics with operating conditions. During the investigation to determine the performance of the X24C-2 compressor (references 2 and 3), instrumentation was added to study the surge characteristics and to determine the effect of speed and inlet pressure on the frequency, amplitude, and phase relation of the pressure pulsations behind each stage.
Buckner, Howard A., Jr.↗
Improved method for calculating transonic velocities on blade-to-blade stream surfaces of a turbomachine
A method was developed to improve the accuracy of an existing computer program used to calculate transonic velocities on a blade-to-blade surface of a turbomachine. The method eliminates problems encountered in obtaining solutions with the velocity gradient equation when large gradients in velocity occur through the blade row. With the improved method, results indicate that the transonic solution can be obtained by scaling the velocities obtained at the reduced mass flow rate where all velocities are subsonic thereby eliminating the need for a solution of the velocity gradient equation. Solutions obtained with the scaling method on a two dimensional compressor cascade and an axial turbine stator show good agreement with experimental data. The results obtained for the stationary blade rows and comparison of analytical results obtained with and without the present method suggest that the method will yield an improved solution for centrifugal compressor impellers.
Wood, J. R.↗
Conceptual Design of a Two Spool Compressor for the NASA Large Civil Tilt Rotor Engine
This paper focuses on the conceptual design of a two spool compressor for the NASA Large Civil Tilt Rotor engine, which has a design-point pressure ratio goal of 30:1 and an inlet weight flow of 30.0 lbm/sec. The compressor notional design requirements of pressure ratio and low-pressure compressor (LPC) and high pressure ratio compressor (HPC) work split were based on a previous engine system study to meet the mission requirements of the NASA Subsonic Rotary Wing Projects Large Civil Tilt Rotor vehicle concept. Three mean line compressor design and flow analysis codes were utilized for the conceptual design of a two-spool compressor configuration. This study assesses the technical challenges of design for various compressor configuration options to meet the given engine cycle results. In the process of sizing, the technical challenges of the compressor became apparent as the aerodynamics were taken into consideration. Mechanical constraints were considered in the study such as maximum rotor tip speeds and conceptual sizing of rotor disks and shafts. The rotor clearance-to-span ratio in the last stage of the LPC is 1.5% and in the last stage of the HPC is 2.8%. Four different configurations to meet the HPC requirements were studied, ranging from a single stage centrifugal, two axi-centrifugals, and all axial stages. Challenges of the HPC design include the high temperature (1,560deg R) at the exit which could limit the maximum allowable peripheral tip speed for centrifugals, and is dependent on material selection. The mean line design also resulted in the definition of the flow path geometry of the axial and centrifugal compressor stages, rotor and stator vane angles, velocity components, and flow conditions at the leading and trailing edges of each blade row at the hub, mean and tip. A mean line compressor analysis code was used to estimate the compressor performance maps at off-design speeds and to determine the required variable geometry reset schedules of the inlet guide vane and variable stators that would result in the transonic stages being aerodynamically matched with high efficiency and acceptable stall margins based on user specified maximum levels of rotor diffusion factor and relative velocity ratio.
Veres, Joseph P.↗
A laser velocimeter system for small radial turbomachinery
The paper describes the design and test of a laser Doppler velocimeter (LDV) to map the velocity fields in small high-speed turbomachinery. This instrument is presently being used as a diagnostic tool for the study of the basic fluid dynamics of the inducer, impeller and the diffuser regions of high-pressure-ratio small centrifugal compressors. The LDV instrumentation has been optimized to permit measurement of instantaneous velocities up to approximately 500 m/s within a rotating impeller. It is shown that LDV can make nondisturbing gas velocity measurements and surmount the difficulties found in applying other types of aerodynamic instrumentation, e.g., hot-wire anemometers, to this measurement problem.
Runstadler, P. W., Jr.↗
Advanced radial inflow turbine rotor program: Design and dynamic testing
The advancement of small, cooled, radial inflow turbine technology in the area of operation at higher turbine inlet temperature is discussed. The first step was accomplished by designing, fabricating, and subjecting to limited mechanical testing an advanced gas generator rotating assembly comprising a radial inflow turbine and two-stage centrifugal compressor. The radial inflow turbine and second-stage compressor were designed as an integrally machined monorotor with turbine cooling taking place basically by conduction to the compressor. Design turbine inlet rotor gas temperature, rotational speed, and overall gas generator compressor pressure ratio were 1422 K (2560 R), 71,222 rpm, and 10/1 respectively. Mechanical testing on a fabricated rotating assembly and bearing system covered 1,000 cold start/stop cycles and three spins to 120 percent design speed (85,466 rpm).
Rodgers, C.↗
Calculation of 3-dimensional choking mass flow in turbomachinery with 2-dimensional flow models
An approach is considered for obtaining an approximate flow solution in the case of a cross-sectional flow surface within a guided channel, taking into account a pair of typical turbine blades with three-dimensional orthogonal surfaces across the flow passage, the calculation of the mass flow across the throat in the case of a 2-D passage with curved walls, and the determination of the choking mass flow. It is pointed out that the choking solution for a three-dimensional guided passage in a blade row can be obtained in a very similar manner by satisfying momentum equations for the blade-to-blade and the hub-to-tip direction. A considered example involves the calculation of the choking mass flow for a centrifugal compressor impeller in an automotive application.
Katsanis, T.↗