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  • 1
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    Core compressor exit stage study. Volume 3: Data and performance report for screening test configurations (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Rear stage blading designs that have lower losses in their endwall boundary layer regions were developed. Test data and performance results for rotor B, stator B, and stator C - blading designs that offer promise of reducing endwall losses relative to the baseline are given. A low speed research compressor was the principal investigative tool. The tests were conducted using four identical stages of blading so that the test data would be obtained in a true multistage environment.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-159499 , R80AEG313-VOL-3
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 2
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    Core compressor exit stage study. Volume 2: Data and performance report for the baseline configuration (1980)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The objective of the program is to develop rear stage blading designs that have lower losses in their endwall boundary layer regions. The overall technical approach in this efficiency improvement program utilized General Electric's Low Speed Research Compressor as the principal investigative tool. Tests were conducted in two ways: using four identical stages of blading so that test data would be obtained in a true multistage environment and using a single stage of blading so that comparison with the multistage test results could be made.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-159498 , R80AEG312-VOL-2
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Study of casing treatment stall margin improvement phenomena (1974)
    In:  CASI
    Details
    Publication Date: 2019-06-27
    Description: The results of a program of experimental and analytical research in casing treatments over axial compressor rotor blade tips are presented. Circumferential groove, axial-skewed slot, and blade angle slot treatments were tested. These yielded, for reduction in stalling flow and loss in peak efficiency, 5.8% and 0 points, 15.3% and 2.0 points, and 15.0% and 1.2 points, respectively. These values are consistent with other experience. The favorable stalling flow situations correlated well with observations of higher-then-normal surface pressures on the rotor blade pressure surfaces in the tip region, and with increased maximum diffusions on the suction surfaces. Annular wall pressure gradients, especially in the 50-75% chord region, are also increased and blade surface pressure loadings are shifted toward the trailing edge for treated configurations. Rotor blade wakes may be somewhat thinner in the presence of good treatments, particularly under operating conditions close to the baseline stall.
    Keywords: PROPULSION SYSTEMS
    Type: NASA-CR-134552 , R73AEG326
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Core compressor exit stage study, volume 6 (1981)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: Rear stage blading designs that have lower losses in their endwall boundary layer regions were studied. A baseline Stage A was designed as a low-speed model of stage 7 of a 10-stage compressor. Candidate rotors and stators were designed which have the potential of reducing endwall losses relative to the baseline. Rotor B uses a type of meanline in the tip region that unloads the leading edge and loads the trailing edge relative to the baseline rotor A designs. Rotor C incorporates a more skewed (hub strong) radial distribution of total pressure and smoother distribution of static pressure on the rotor tip than those of rotor B. Candidate stator B embodies twist gradients in the endwall region. Stator C embodies airfoil sections near the endwalls that have reduced trailing edge loading relative to stator A. The baseline and candidate bladings were tested using four identical stages to produce a true multistage environment. Single-stage tests were also conducted. The test data were analyzed and performances were compared. Several of the candidate configurations showed a performance improvement relative to the baseline.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-165553 , NAS 1.26:165554 , GE-R81AEG288
    Format: application/pdf
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  • 5
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    A study of casing treatment stall margin improvement phenomena (1975)
    In:  Other Sources
    Details
    Publication Date: 2019-07-13
    Description: The results of a program of experimental and analytical research in casing treatments over axial compressor rotor blade tips are presented. Circumferential groove, axial-skewed slot and blade angle slot treatments were tested at low speeds. With the circumferential groove treatment the stalling flow was reduced 5.8% at negligible efficiency sacrifice. The axial-skewed slot treatment improved the stalling flow by 15.3%; 1.8 points in peak efficiency were sacrificed. The blade angle slot treatment improved the stalling flow by 15.0%; 1.4 points in peak efficiency were sacrificed. The favorable stalling flow situations correlated well with observations of higher-than-normal surface pressures on the rotor blade pressure surfaces in the tip region, and with increased maximum diffusions on the suction surfaces. Annulus wall pressure gradients, especially in the 50 to 75% chord region, are also increased and blade surface pressure loadings are shifted toward the trailing edge for treated configurations.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: ASME PAPER 75-GT-60 , American Society of Mechanical Engineers, Gas Turbine Conference and Products Show; Mar 02, 1975 - Mar 06, 1975; Houston, TX
    Format: text
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  • 6
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    Stator hub treatment study (1974)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The results of an experimental research program to investigate the potential of improving compressor stall margin by the application of hub treatment are presented. Extensive tuft probing showed that the two-stage, 0.5 radius ratio compressor selected for the test was indeed hub critical. Circumferential groove and baffled wide blade angle slot hub treatments under the stators were tested. Performance measurements were made with total and static pressure probes, wall static pressure taps, flow angle measuring instrumentation and hot film anemometers. Stator hub treatment was not found to be effective in improving compressor stall margin by delaying the point of onset of rotating stall or in modifying compressor performance for any of the configurations tested. Extensive regions of separated flow were observed on the suction surface of the stators near the hub. However, the treatment did not delay the point where flow separation in the stator hub region becomes apparent.
    Keywords: MECHANICAL ENGINEERING
    Type: NASA-CR-134729 , R74AEG410
    Format: application/pdf
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  • 7
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    Preliminary design study of advanced multistage axial flow core compressors (1977)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: A preliminary design study was conducted to identify an advanced core compressor for use in new high-bypass-ratio turbofan engines to be introduced into commercial service in the 1980's. An evaluation of anticipated compressor and related component 1985 state-of-the-art technology was conducted. A parametric screening study covering a large number of compressor designs was conducted to determine the influence of the major compressor design features on efficiency, weight, cost, blade life, aircraft direct operating cost, and fuel usage. The trends observed in the parametric screening study were used to develop three high-efficiency, high-economic-payoff compressor designs. These three compressors were studied in greater detail to better evaluate their aerodynamic and mechanical feasibility.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-135133 , R77AEG222
    Format: application/pdf
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  • 8
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    Core compressor exit stage study. Volume 5: Design and performance report for the Rotor C/Stator B configuration (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: Rear stage blading designs that have lower losses in their endwall boundary layer regions were developed. The design of rotor-C and the performance results for rotor-C running with stator B are described. A low speed research compressor is utilized as the principal investigative tool. Four identical stages of blading are used to obtained data in a true multistage environment.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-165358 , R81AEG287-VOL-5
    Format: application/pdf
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  • 9
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    Core compressor exit stage study. Volume 4: Data and performance report for the best stage configuration (1981)
    In:  CASI
    Details
    Publication Date: 2019-06-28
    Description: The core compressor exit stage study program develops rear stage blading designs that have lower losses in their endwall boundary layer regions. The test data and performance results for the best stage configuration consisting of Rotor-B running with Stator-B are described. The technical approach in this efficiency improvement program utilizes a low speed research compressor. Tests were conducted in two ways: (1) to use four identical stages of blading to obtain test data in a true multistage environment and (2) to use a single stage of blading to compare with the multistage test results. The effects of increased rotor tip clearances and circumferential groove casing treatment are evaluated.
    Keywords: AIRCRAFT PROPULSION AND POWER
    Type: NASA-CR-165357 , R80AEG314-VOL-4
    Format: application/pdf
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  • 10
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    Loss reduction in axial-flow compressors through low-speed model testing (1984)
    In:  Other Sources
    Details
    Publication Date: 2019-06-28
    Description: A systematic procedure for reducing losses in axial-flow compressors is presented. In this procedure, a large, low-speed, aerodynamic model of a high-speed core compressor is designed and fabricated based on aerodynamic similarity principles. This model is then tested at low speed where high-loss regions associated with three-dimensional endwall boundary layers flow separation, leakage, and secondary flows can be located, detailed measurements made, and loss mechanisms determined with much greater accuracy and much lower cost and risk than is possible in small, high-speed compressors. Design modifications are made by using custom-tailored airfoils and vector diagrams, airfoil endbends, and modified wall geometries in the high-loss regions. The design improvements resulting in reduced loss or increased stall margin are then scaled to high speed. This paper describes the procedure and presents experimental results to show that in some cases endwall loss has been reduced by as much as 10 percent, flow separation has been reduced or eliminated, and stall margin has been substantially improved by using these techniques.
    Keywords: AERODYNAMICS
    Type: ASME PAPER 84-GT-184
    Format: text
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