Results for “Impeller” — Page 11

NASA NTRS1999

Rotary blood pump

A blood pump that comprises a pump housing having a blood flow path therethrough, a blood inlet, and a blood outlet; a stator mounted to the pump housing, the stator having a stator field winding for producing a stator magnetic field; a flow straightener located within the pump housing, and comprising a flow straightener hub and at least one flow straightener blade attached to the flow straightener hub; a rotor mounted within the pump housing for rotation in response to the stator magnetic field, the rotor comprising an inducer and an impeller; the inducer being located downstream of the flow straightener, and comprising an inducer hub and at least one inducer blade attached to the inducer hub; the impeller being located downstream of the inducer, and comprising an impeller hub and at least one impeller blade attached to the impeller hub; and preferably also comprising a diffuser downstream of the impeller, the diffuser comprising a diffuser hub and at least one diffuser blade. Blood flow stagnation and clot formation within the pump are minimized by, among other things, providing the inducer hub with a diameter greater than the diameter of the flow straightener hub; by optimizing the axial spacing between the flow straightener hub and the inducer hub, and between the impeller hub and the diffuser hub; by optimizing the inlet angle of the diffuser blades; and by providing fillets or curved transitions between the upstream end of the inducer hub and the shaft mounted therein, and between the impeller hub and the shaft mounted therein.

Benkowski, Robert J.↗

NASA NTRS2014

High Efficiency Centrifugal Compressor for Rotorcraft Applications

The report "High Efficiency Centrifugal Compressor for Rotorcraft Applications" documents the work conducted at UTRC under the NRA Contract NNC08CB03C, with cost share 2/3 NASA, and 1/3 UTRC, that has been extended to 4.5 years. The purpose of this effort was to identify key technical barriers to advancing the state-of-the-art of small centrifugal compressor stages; to delineate the measurements required to provide insight into the flow physics of the technical barriers; to design, fabricate, install, and test a state-of-the-art research compressor that is representative of the rear stage of an axial-centrifugal aero-engine; and to acquire detailed aerodynamic performance and research quality data to clarify flow physics and to establish detailed data sets for future application. The design activity centered on meeting the goal set outlined in the NASA solicitation-the design target was to increase efficiency at higher work factor, while also reducing the maximum diameter of the stage. To fit within the existing Small Engine Components Test Facility at NASA Glenn Research Center (GRC) and to facilitate component re-use, certain key design parameters were fixed by UTRC, including impeller tip diameter, impeller rotational speed, and impeller inlet hub and shroud radii. This report describes the design effort of the High Efficiency Centrifugal Compressor stage (HECC) and delineation of measurements, fabrication of the compressor, and the initial tests that were performed. A new High-Efficiency Centrifugal Compressor stage with a very challenging reduction in radius ratio was successfully designed, fabricated and installed at GRC. The testing was successful, with no mechanical problems and the running clearances were achieved without impeller rubs. Overall, measured pressure ratio of 4.68, work factor of 0.81, and at design exit corrected flow rate of 3 lbm/s met the target requirements. Polytropic efficiency of 85.5 percent and stall margin of 7.5 percent were measured at design flow rate and speed. The measured efficiency and stall margin were lower than pre-test CFD predictions by 2.4 percentage points (pt) and 4.5 pt, respectively. Initial impressions from the experimental data indicated that the loss in the efficiency and stall margin can be attributed to a design shortfall in the impeller. However, detailed investigation of experimental data and post-test CFD simulations of higher fidelity than pre-test CFD, and in particular the unsteady CFD simulations and the assessment with a wider range of turbulence models, have indicated that the loss in efficiency is most likely due to the impact of unfavorable unsteady impeller/diffuser interactions induced by diffuser vanes, an impeller/diffuser corrected flow-rate mismatch (and associated incidence levels), and, potentially, flow separation in the radial-to-axial bend. An experimental program with a vaneless diffuser is recommended to evaluate this observation. A subsequent redesign of the diffuser (and the radial-to-axial bend) is also recommended. The diffuser needs to be redesigned to eliminate the mismatching of the impeller and the diffuser, targeting a slightly higher flow capacity. Furthermore, diffuser vanes need to be adjusted to align the incidence angles, to optimize the splitter vane location (both radially and circumferentially), and to minimize the unsteady interactions with the impeller. The radial-to-axial bend needs to be redesigned to eliminate, or at least minimize, the flow separation at the inner wall, and its impact on the flow in the diffuser upstream. Lessons were also learned in terms of CFD methodology and the importance of unsteady CFD simulations for centrifugal compressors was highlighted. Inconsistencies in the implementation of a widely used two-equation turbulence model were identified and corrections are recommended. It was also observed that unsteady simulations for centrifugal compressors require significantly longer integration times than what is current practice in industry.

Gorazd Medic↗

NASA NTRS2014

High Efficiency Centrifugal Compressor for Rotorcraft Applications

The report "High Efficiency Centrifugal Compressor for Rotorcraft Applications" documents the work conducted at UTRC under the NRA Contract NNC08CB03C, with cost share 2/3 NASA, and 1/3 UTRC, that has been extended to 4.5 years. The purpose of this effort was to identify key technical barriers to advancing the state-of-the-art of small centrifugal compressor stages; to delineate the measurements required to provide insight into the flow physics of the technical barriers; to design, fabricate, install, and test a state-of-the-art research compressor that is representative of the rear stage of an axial-centrifugal aero-engine; and to acquire detailed aerodynamic performance and research quality data to clarify flow physics and to establish detailed data sets for future application. The design activity centered on meeting the goal set outlined in the NASA solicitation-the design target was to increase efficiency at higher work factor, while also reducing the maximum diameter of the stage. To fit within the existing Small Engine Components Test Facility at NASA Glenn Research Center (GRC) and to facilitate component re-use, certain key design parameters were fixed by UTRC, including impeller tip diameter, impeller rotational speed, and impeller inlet hub and shroud radii. This report describes the design effort of the High Efficiency Centrifugal Compressor stage (HECC) and delineation of measurements, fabrication of the compressor, and the initial tests that were performed. A new High-Efficiency Centrifugal Compressor stage with a very challenging reduction in radius ratio was successfully designed, fabricated and installed at GRC. The testing was successful, with no mechanical problems and the running clearances were achieved without impeller rubs. Overall, measured pressure ratio of 4.68, work factor of 0.81, and at design exit corrected flow rate of 3 lbm/s met the target requirements. Polytropic efficiency of 85.5 percent and stall margin of 7.5 percent were measured at design flow rate and speed. The measured efficiency and stall margin were lower than pre-test CFD predictions by 2.4 percentage points (pt) and 4.5 pt, respectively. Initial impressions from the experimental data indicated that the loss in the efficiency and stall margin can be attributed to a design shortfall in the impeller. However, detailed investigation of experimental data and post-test CFD simulations of higher fidelity than pre-test CFD, and in particular the unsteady CFD simulations and the assessment with a wider range of turbulence models, have indicated that the loss in efficiency is most likely due to the impact of unfavorable unsteady impeller/diffuser interactions induced by diffuser vanes, an impeller/diffuser corrected flow-rate mismatch (and associated incidence levels), and, potentially, flow separation in the radial-to-axial bend. An experimental program with a vaneless diffuser is recommended to evaluate this observation. A subsequent redesign of the diffuser (and the radial-to-axial bend) is also recommended. The diffuser needs to be redesigned to eliminate the mismatching of the impeller and the diffuser, targeting a slightly higher flow capacity. Furthermore, diffuser vanes need to be adjusted to align the incidence angles, to optimize the splitter vane location (both radially and circumferentially), and to minimize the unsteady interactions with the impeller. The radial-to-axial bend needs to be redesigned to eliminate, or at least minimize, the flow separation at the inner wall, and its impact on the flow in the diffuser upstream. Lessons were also learned in terms of CFD methodology and the importance of unsteady CFD simulations for centrifugal compressors was highlighted. Inconsistencies in the implementation of a widely used two-equation turbulence model were identified and corrections are recommended. It was also observed that unsteady simulations for centrifugal compressors require significantly longer integration times than what is current practice in industry.

Gorazd Medic↗

NASA NTRS2017

High Efficiency Centrifugal Compressor for Rotorcraft Applications

The report "High Efficiency Centrifugal Compressor for Rotorcraft Applications" documents the work conducted at UTRC under the NRA Contract NNC08CB03C, with cost share 2/3 NASA, and 1/3 UTRC, that has been extended to 4.5 years. The purpose of this effort was to identify key technical barriers to advancing the state-of-the-art of small centrifugal compressor stages; to delineate the measurements required to provide insight into the flow physics of the technical barriers; to design, fabricate, install, and test a state-of-the-art research compressor that is representative of the rear stage of an axial-centrifugal aero-engine; and to acquire detailed aerodynamic performance and research quality data to clarify flow physics and to establish detailed data sets for future application. The design activity centered on meeting the goal set outlined in the NASA solicitation-the design target was to increase efficiency at higher work factor, while also reducing the maximum diameter of the stage. To fit within the existing Small Engine Components Test Facility at NASA Glenn Research Center (GRC) and to facilitate component re-use, certain key design parameters were fixed by UTRC, including impeller tip diameter, impeller rotational speed, and impeller inlet hub and shroud radii. This report describes the design effort of the High Efficiency Centrifugal Compressor stage (HECC) and delineation of measurements, fabrication of the compressor, and the initial tests that were performed. A new High-Efficiency Centrifugal Compressor stage with a very challenging reduction in radius ratio was successfully designed, fabricated and installed at GRC. The testing was successful, with no mechanical problems and the running clearances were achieved without impeller rubs. Overall, measured pressure ratio of 4.68, work factor of 0.81, and at design exit corrected flow rate of 3 lbm/s met the target requirements. Polytropic efficiency of 85.5 percent and stall margin of 7.5 percent were measured at design flow rate and speed. The measured efficiency and stall margin were lower than pre-test CFD predictions by 2.4 percentage points (pt) and 4.5 pt, respectively. Initial impressions from the experimental data indicated that the loss in the efficiency and stall margin can be attributed to a design shortfall in the impeller. However, detailed investigation of experimental data and post-test CFD simulations of higher fidelity than pre-test CFD, and in particular the unsteady CFD simulations and the assessment with a wider range of turbulence models, have indicated that the loss in efficiency is most likely due to the impact of unfavorable unsteady impeller/diffuser interactions induced by diffuser vanes, an impeller/diffuser corrected flow-rate mismatch (and associated incidence levels), and, potentially, flow separation in the radial-to-axial bend. An experimental program with a vaneless diffuser is recommended to evaluate this observation. A subsequent redesign of the diffuser (and the radial-to-axial bend) is also recommended. The diffuser needs to be redesigned to eliminate the mismatching of the impeller and the diffuser, targeting a slightly higher flow capacity. Furthermore, diffuser vanes need to be adjusted to align the incidence angles, to optimize the splitter vane location (both radially and circumferentially), and to minimize the unsteady interactions with the impeller. The radial-to-axial bend needs to be redesigned to eliminate, or at least minimize, the flow separation at the inner wall, and its impact on the flow in the diffuser upstream. Lessons were also learned in terms of CFD methodology and the importance of unsteady CFD simulations for centrifugal compressors was highlighted. Inconsistencies in the implementation of a widely used two-equation turbulence model were identified and corrections are recommended. It was also observed that unsteady simulations for centrifugal compressors require significantly longer integration times than what is current practice in industry.

Centrifugal compressor↗

NASA NTRSDate not supplied

NASA HECC Geometry and Performance Review Part 3: A Numerical and Experimental Investigation of Tip Clearance Effects on the Vaneless Diffuser Configuration

Tip clearance effects in centrifugal compressors have been extensively investigated to understand the losses associated with the flow in the impeller tip clearance gap between the rotating blades and the stationary shroud. In Part 3 of this multipart investigation, experimental data and validated numerical simulations from Parts 1 and 2 were used to analyze the effects of the size of the tip clearance gap on the High Efficiency Centrifugal Compressor performance and aerodynamics. Four tip clearance gaps ranging from 0.012-in to 0.030-in (2.0% to 4.9% of the impeller exit blade height) were considered at both design and off-design operating conditions. The total pressure ratio and efficiency of the stage are found to decay linearly with increasing tip gaps. The sensitivity of the impeller performance to the tip gap was found to vary with the rotational speed of the compressor. Spanwise surveys of flow angle, total pressure, and total temperature collected at the impeller exit at design speed are used to validate numerical simulations at each experimental tip gap condition. Numerical simulations show that increased turbulence kinetic energy near the shroud at larger tip gaps leads to spanwise mixing of high entropy fluid near the impeller trailing edge which decreases the useful work input by the impeller. The data presented have been made available to the public in the HECC Data Archive located at https://storage.googleapis.com/hecc-data/NASA-HECC-Data-Archive.zip.

centrifugal compressor↗

NASA NTRSDate not supplied

NASA HECC Geometry and Performance Review Part 3: A Numerical and Experimental Investigation of Tip Clearance Effects on the Vaneless Diffuser Configuration

Tip clearance effects in centrifugal compressors have been extensively investigated to understand the losses associated with the flow in the impeller tip clearance gap between the rotating blades and the stationary shroud. In Part 3 of this multipart investigation, experimental data and validated numerical simulations from Parts 1 and 2 were used to analyze the effects of the size of the tip clearance gap on the High Efficiency Centrifugal Compressor performance and aerodynamics. Four tip clearance gaps ranging from 0.012-in to 0.030-in (2.0% to 4.9% of the impeller exit blade height) were considered at both design and off-design operating conditions. The total pressure ratio and efficiency of the stage are found to decay linearly with increasing tip gaps. The sensitivity of the impeller performance to the tip gap was found to vary with the rotational speed of the compressor. Spanwise surveys of flow angle, total pressure, and total temperature collected at the impeller exit at design speed are used to validate numerical simulations at each experimental tip gap condition. Numerical simulations show that increased turbulence kinetic energy near the shroud at larger tip gaps leads to spanwise mixing of high entropy fluid near the impeller trailing edge which decreases the useful work input by the impeller. The data presented have been made available to the public in the HECC Data Archive located at https://storage.googleapis.com/hecc-data/NASA-HECC-Data-Archive.zip.

centrifugal compressor↗

NASA NTRS1988

Turbomachinery rotor forces

The fluid-induced forces, both steady and unsteady, acting upon an impeller of a centrifugal pump, and impeller blade-diffuser vane interaction in centrifugal pumps with vaned radial diffusers were evaluated experimentally and theoretically. Knowledge of the steady and unsteady forces, and the associated rotordynamic coefficients are required to effectively model the rotor dynamics of the High Pressure Fuel Turbopump (HPFTP) of the Space Shuttle Main Engine (SSME). These forces and rotordynamic coefficients were investigated using different impellers in combination with volutes and vaned diffusers, and axial inducers. These rotor forces are global. Local forces and pressures are also important in impeller-diffuser interaction, for they may cause cavitation damage and even vane failures. Thus, in a separate investigation, impeller wake, and impeller blade and diffuser vane pressure measurements were made. The nature of the rotordynamic forces is discussed, the experimental facility is described, and the measurements of unsteady forces and pressure are reported together with a brief and incomplete attempt to calculate these flows.

Arndt, Norbert↗

NASA NTRS1993

Bulk-Flow Analysis, part A

The bulk-flow analysis results for this contract are incorporated in the following publications: 'Fluid-Structure Interaction Forces at Pump-Impeller Shroud Surfaces for Axial Vibration Analysis'; 'Centrifugal Acceleration Modes for Incompressible Fluid in the Leakage Annulus Between a Shrouded Pump Impeller and Its Housing'; 'Influence of Impeller Shroud Forces on Pump Rotordynamics'; 'Pressure Oscillation in the Leakage Annulus Between a Shrouded Impeller and Its Housing Due to Impeller-Discharge-Pressure Disturbances'; and 'Compressibility Effects on Rotor Forces in the Leakage Path Between a Shrouded Pump Impeller and Its Housing'. These publications are summarized and included in this final report. Computational Fluid Mechanics (CFD) results developed by Dr. Erian Baskharone are reported separately.

Childs, Dara W.↗

NASA NTRS1993

Incompressible Navier-Stokes Calculations in Pump Flows

Flow through pump components, such as the SSME-HPFTP Impeller and an advanced rocket pump impeller, is efficiently simulated by solving the incompressible Navier-Stokes equations. The solution method is based on the pseudo compressibility approach and uses an implicit-upwind differencing scheme together with the Gauss-Seidel line relaxation method. The equations are solved in steadily rotating reference frames and the centrifugal force and the Coriolis force are added to the equation of motion. Current computations use one-equation Baldwin-Barth turbulence model which is derived from a simplified form of the standard k-epsilon model equations. The resulting computer code is applied to the flow analysis inside an 11-inch SSME High Pressure Fuel Turbopump impeller, and an advanced rocket pump impeller. Numerical results of SSME-HPFTP impeller flow are compared with experimental measurements. In the advanced pump impeller, the effects of exit and shroud cavities are investigated. Flow analyses at design conditions will be presented.

Kiris, Cetin↗

NASA NTRS1948

Performance Investigations of a Large Centrifugal Compressor from an Experimental Turbojet Engine

An investigation was conducted on a large centrifugal compressor from an experimental turbojet engine to determine the performance of the compressor and to obtain fundamental information on the aerodynamic problems associated with large centrifugal-type compressors. The results of the research conducted on the compressor indicated that the compressor would not meet the desired engine-design air-flow requirements (78 lb/sec) because of an air-flow restriction in the vaned collector (diffuser). Revision of the vaned collector resulted in an increased air-flow capacity over the speed range and showed improved matching of the impeller and diffuser components. At maximum flow, the original compressor utilized approximately 90 percent of the available geometric throat area at the vaned-collector inlet and the revised compressor utilized approximately 94 percent, regardless of impeller speed. The ratio of the maximum weight flows of the revised and original compressors were less than the ratio of effective critical throat areas of the two compressors because of the large pressure losses in the impeller near the impeller inelt and the difference increased with an increase in impeller speed. In order to further increase the pressure ratio and maximum weight flow of the compressor, the impeller must be modified to eliminate the pressure losses therein.

Ginsburg, Ambrose↗

NASA NTRS1948

Performance of Compressor of XJ-41-V Turbojet Engine VI - Analysis of Compressor Flow Choking: Analysis of Compressor Flow Choking - 6

An extended analysis was made of the previously reported performance investigation of the original compressor from the XJ-41-v turbojet engine and a similar compressor revised a to obtain a 33-percent increase in the geometric passage area at the vaned-collector entrance. This analysis was based on the concept of the vaned-collector entrance as the throat section of a nozzle. Because of nonuniform air distribution at the vaned-collector entrance, approximately 90 percent of the available flow area was utilized in the original compressor and 94percent in the revised com$ressor. The increase in maximum weight flow obtained with the revised compressor was disproportionate to the increased effective critical throat area because. the air density at the revised vaned-collector entrance for maximum flow was lower than that obtained in the original compressor. This reduction in density resulted from the large pressure losses near the impeller inlet of the revised compressor, which is indicative of impending flow choking in the impeller, The.calculated maximum corrected weight-flow capacity of a compressor consisting of the revised vaneless diffuser and vaned collector with a theoretical impeller that combined peak impeller pressure ratio and peak impeller efficiency at the . maximum flow point would be 112 pounds per second for an equivalent impeller speed of 11,500 rpm;

Creagh, John W. R.↗

NASA NTRS1949

Calibration Tests of a Japanese Log Rodmeter

A Japanese log rodmeter of the rotating-vane impeller type, with a co mmutator on the impeller shaft, was calibrated in Langley tank no. 1. The rotational speed of two impellers was determined for forward speeds up to 24 knots at angles of yaw up to ?10 0 . In general, the rota tional speeds of two apparently identical impellers tested in the rod meter decreased with increasing yaw angle, right yaw causing a greate r decrease than left yaw. The difference in calibration between the t wo impellers was approximately the same as that produced by a change in yaw angle from 50 left to 50 right. Evidence of cavitation within the impeller fairing appeared at speeds above 24 knots.

Mottard, Elmo J.↗

NASA NTRSDate not supplied

Experimental and Numerical Investigation of the NASA High Efficiency Centrifugal Compressor Vaned Stage Geometry and Aerodynamic Performance

Since its inception in the early 2010s, the NASA High Efficiency Centrifugal Compressor (HECC) has been enigmatic for the propulsion research community: experimental data and numerical simulations of the stage have generally not aligned in their quantifications of performance metrics. Typically, the fault for these disagreements is assigned to the numerical simulations as the simulations are models, and models are inherently incomplete representations of the experiment. This “incompleteness” may manifest in assumptions regarding roughness or heat transfer, simplifications of the flow path (i.e., neglecting bleed flows), or the oft-scapegoated turbulence model. In the case of HECC, recent work showed unexpected discrepancies between the intended impeller geometry defined in the design report (termed Design-Intent) and the manufactured impeller used in the experimental campaigns (termed As-Manufactured). That work used numerical simulations to establish that the geometric differences between the Design-Intent and As-Manufactured impellers were significant enough to result in drastically different performance predictions for the HECC vaneless diffuser configuration. The Design-Intent impeller simulations over predicted the performance relative to the experiment, whereas the As-Manufactured simulations better represented the experimental data, both in terms of one-dimensional performance metrics and spanwise flow profiles. This effort expands on that work by examining in detail the geometry and aerodynamic performance of the HECC vaned diffuser configuration. Further differences between the Design-Intent and As-Manufactured geometries have been discovered in the vaned diffuser and exit guide vanes, and these differences are documented herein. The summations of the geometric differences for all of the components were used to create two numerical models of HECC vaned diffuser configuration: the Design-Intent simulations which are generated from the original geometry definitions given in the design report and the As-Manufactured simulations which are the best available representation of the manufactured compressor hardware used in the experimental test campaigns. In congruence with the earlier vaneless diffuser work, the numerical predictions of the Design-Intent choked mass flow rate, total pressure ratio, and efficiency were notably greater than that of the As-Manufactured simulations. To increase confidence in the experimental dataset, measurements from a recent test campaign conducted in 2024 are used to validate the original experimental data acquired from 2012 to 2014 with good repeatability overall, especially considering the passage of time and differences in the data acquisition systems between the test campaigns. Both numerical simulations were then extensively evaluated against the experimental data. The As-Manufactured simulations provided better estimates of the stage performance than the Design-Intent cases in terms of most performance metrics. Nonetheless, more detailed results still show opportunities for improvement. Despite a more accurate representation of the physical hardware, characterization of the impeller work input remains a challenge even for rigorously developed numerical models.

vaned diffuser↗

NASA NTRSDate not supplied

Experimental and Numerical Investigation of the NASA High Efficiency Centrifugal Compressor Vaned Stage Geometry and Aerodynamic Performance

Since its inception in the early 2010s, the NASA High Efficiency Centrifugal Compressor (HECC) has been enigmatic for the propulsion research community: experimental data and numerical simulations of the stage have generally not aligned in their quantifications of performance metrics. Typically, the fault for these disagreements is assigned to the numerical simulations as the simulations are models, and models are inherently incomplete representations of the experiment. This “incompleteness” may manifest in assumptions regarding roughness or heat transfer, simplifications of the flow path (i.e., neglecting bleed flows), or the oft-scapegoated turbulence model. In the case of HECC, recent work showed unexpected discrepancies between the intended impeller geometry defined in the design report (termed Design-Intent) and the manufactured impeller used in the experimental campaigns (termed As-Manufactured). That work used numerical simulations to establish that the geometric differences between the Design-Intent and As-Manufactured impellers were significant enough to result in drastically different performance predictions for the HECC vaneless diffuser configuration. The Design-Intent impeller simulations over predicted the performance relative to the experiment, whereas the As-Manufactured simulations better represented the experimental data, both in terms of one-dimensional performance metrics and spanwise flow profiles. This effort expands on that work by examining in detail the geometry and aerodynamic performance of the HECC vaned diffuser configuration. Further differences between the Design-Intent and As-Manufactured geometries have been discovered in the vaned diffuser and exit guide vanes, and these differences are documented herein. The summations of the geometric differences for all of the components were used to create two numerical models of HECC vaned diffuser configuration: the Design-Intent simulations which are generated from the original geometry definitions given in the design report and the As-Manufactured simulations which are the best available representation of the manufactured compressor hardware used in the experimental test campaigns. In congruence with the earlier vaneless diffuser work, the numerical predictions of the Design-Intent choked mass flow rate, total pressure ratio, and efficiency were notably greater than that of the As-Manufactured simulations. To increase confidence in the experimental dataset, measurements from a recent test campaign conducted in 2024 are used to validate the original experimental data acquired from 2012 to 2014 with good repeatability overall, especially considering the passage of time and differences in the data acquisition systems between the test campaigns. Both numerical simulations were then extensively evaluated against the experimental data. The As-Manufactured simulations provided better estimates of the stage performance than the Design-Intent cases in terms of most performance metrics. Nonetheless, more detailed results still show opportunities for improvement. Despite a more accurate representation of the physical hardware, characterization of the impeller work input remains a challenge even for rigorously developed numerical models.

centrifugal compressor↗

NASA NTRS1972

Study of blade clearance effects on centrifugal pumps

A program of analysis, design, fabrication, and testing has been conducted to develop and experimentally verify analytical models to predict the effects of impeller blade clearance on centrifugal pumps. The effect of tip clearance on pump efficiency, and the relationship between the head coefficient and torque loss with tip clearance was established. Analysis were performed to determine the cost variation in design, manufacture, and test that would occur between unshrouded and shrouded impellers. An impeller, representative of typical rocket engine impellers, was modified by removing its front shroud to permit variation of its blade clearances. It was tested in water with special instrumentation to provide measurements of blade surface pressures during operation. Pump performance data were obtained from tests at various impeller tip clearances. Blade pressure data were obtained at the nominal tip clearance. Comparisons of predicted and measured data are given.

Hoshide, R. K.↗

NASA NTRS1985

Small centrifugal pumps for low thrust rockets

This paper presents the results of a combined analytical and experimental investigation of low specific speed pumps for potential use as components of propellant feed systems for low thrust rocket engines. Shrouded impellers and open face impellers were tested in volute type and vaned diffuser type pumps. Full- and partial-emission diffusers and full- and partial-admission impellers were tested. Axial and radial loads, head and efficiency versus flow, and cavitation tests were conducted. Predicted performance of two pumps are compared when pumping water and liquid hydrogen. Detailed pressure loss and parasitic power values are presented for two pump configurations. Partial-emission diffusers were found to permit use of larger impeller and diffuser passages with a minimal performance penalty. Normal manufacturing tolerances were found to result in substantial power requirement variation with only a small pressure rise change. Impeller wear ring leakage was found to reduce pump pressure rise to an increasing degree as the pump flowrate was decreased.

Gulbrandsen, N. C.↗

NASA NTRS1989

Turbomachinery shaft insert

An insert is presented for a hollow rotatable shaft on the end of which an impeller is mounted, the insert having a first cylindrical portion receivable within the end of the shaft. The insert includes a radially extending portion adjacent the end of the shaft having an annular wall including radially inner and outer surfaces for engaging tightly similar surfaces in a recess formed in the impeller to provide inner and outer pilot surfaces. The insert also includes a tubular extension which is received within a bore in the impeller, the tubular extension having spaced longitudinally extending slits for permitting the extension together with radially extending hook-like tongues thereon to be received within radial slots in the bore of the impeller to lock the impeller to the insert.

Vonpragenau, George L.↗

NASA NTRS1992

A status of the activities of the NASA/MSFC pump stage technology team

The Consortium for Computational Fluid Dynamics (CFD) Application in Propulsion Technology was established to aid the transfer of CFD related advancements among academia, government agencies, and industry. The specific goals of the Consortium are to develop CFD methodologies necessary to solve propulsion problems, to validate these methodologies, and to apply these methodologies in the design process. To accomplish these goals, a team of experts in various related fields was formed, a schedule of activities necessary to meet the goals was generated, and funding for the activities was obtained from NASA. During the past year (Mar. 1991 - Mar. 1992) the team's activities have focused on preliminary code validation and on the design of an advanced impeller. Six codes were used to calculate the flow in a Rocketdyne 0.3 flow coefficient inducer, and the results were compared to L2F data available for the inducer. This activity identified shortcomings in the experimental data sets and in the analytical solutions which must be surmounted in any future team activity. The design of the advanced impeller relied heavily on CFD results to obtain an optimized geometry. The optimized geometry was analyzed using four different codes, at design and off-design conditions. Activities for the next year include the optimization of a tandem blade impeller design, benchmark of CFD codes for diffuser and volute flows, the collection of L2F data for 'state-of-the-art' impeller and inducer, and the verification of the advanced pump team impeller design in a water rig.

Garcia, R.↗