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  • 1
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    Experimental Methods Applied in a Study of Stall Flutter in an Axial Flow Fan (2004)
    Hindawi
    Details
    Publication Date: 2004-01-01
    Description: Flutter testing is an integral part of aircraft gas turbine engine development. In typical flutter testing blade mounted sensors in the form of strain gages and casing mounted sensors in the form of light probes (NSMS) are used. Casing mounted sensors have the advantage of being non-intrusive and can detect the vibratory response of each rotating blade. Other types of casing mounted sensors can also be used to detect flutter of rotating blades. In this investigation casing mounted high frequency response pressure transducers are used to characterize the part-speed stall flutter response of a single stage unshrouded axial-flow fan. These dynamic pressure transducers are evenly spaced around the circumference at a constant axial location upstream of the fan blade leading edge plane. The pre-recorded experimental data at 70% corrected speed is analyzed for the case where the fan is back-pressured into the stall flutter zone. The experimental data is analyzed using two probe and multi-probe techniques. The analysis techniques for each method are presented. Results from these two analysis methods indicate that flutter occurred at a frequency of 411 Hz with a dominant nodal diameter of 2. The multi-probe analysis technique is a valuable method that can be used to investigate the initiation of flutter in turbomachines.
    Print ISSN: 1070-9622
    Electronic ISSN: 1875-9203
    Topics: Mathematics
    Published by Hindawi
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    PAPER CURRENT
  • 2
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    An Experimental and Computational Investigation of Oscillating Airfoil Unsteady Aerodynamics at Large Mean Incidence (2003)
    In:  CASI
    Details
    Publication Date: 2019-07-10
    Description: A major challenge in the design and development of turbomachine airfoils for gas turbine engines is high cycle fatigue failures due to flutter and aerodynamically induced forced vibrations. In order to predict the aeroelastic response of gas turbine airfoils early in the design phase, accurate unsteady aerodynamic models are required. However, accurate predictions of flutter and forced vibration stress at all operating conditions have remained elusive. The overall objectives of this research program are to develop a transition model suitable for unsteady separated flow and quantify the effects of transition on airfoil steady and unsteady aerodynamics for attached and separated flow using this model. Furthermore, the capability of current state-of-the-art unsteady aerodynamic models to predict the oscillating airfoil response of compressor airfoils over a range of realistic reduced frequencies, Mach numbers, and loading levels will be evaluated through correlation with benchmark data. This comprehensive evaluation will assess the assumptions used in unsteady aerodynamic models. The results of this evaluation can be used to direct improvement of current models and the development of future models. The transition modeling effort will also make strides in improving predictions of steady flow performance of fan and compressor blades at off-design conditions. This report summarizes the progress and results obtained in the first year of this program. These include: installation and verification of the operation of the parallel version of TURBO; the grid generation and initiation of steady flow simulations of the NASA/Pratt&Whitney airfoil at a Mach number of 0.5 and chordal incidence angles of 0 and 10 deg.; and the investigation of the prediction of laminar separation bubbles on a NACA 0012 airfoil.
    Keywords: Aircraft Propulsion and Power
    Type: NASA/CR-2003-212199 , NAS 1.26:212199 , E-13802
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 3
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    Oscillating cascade aerodynamics at large mean incidence (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The aerodynamics of a cascade of airfoils oscillating in torsion about the midchord is investigated experimentally at a large mean incidence angle and, for reference, at a low mean incidence angle. The airfoil section is representative of a modern, low aspect ratio, fan blade tip section. Time-dependent airfoil surface pressure measurements were made for reduced frequencies of up to 1.2 for out-of-phase oscillations at a Mach number of 0.5 and chordal incidence angles of 0 deg and 10 deg; the Reynolds number was 0.9 x l0(exp 6). For the 10 deg chordal incidence angle, a separation bubble formed at the leading edge of the suction surface. The separated flow field was found to have a dramatic effect on the chordwise distribution of the unsteady pressure. In this region, substantial deviations from the attached flow data were found with the deviations becoming less apparent in the aft region of the airfoil for all reduced frequencies. In particular, near the leading edge the separated flow had a strong destabilizing influence while the attached flow had a strong stabilizing influence.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107247 , NAS 1.15:107247 , E-10300 , Gas Turbine and Aeroengine Congress; Jun 10, 1996 - Jun 13, 1996; Brimingham; United Kingdom
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 4
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    Unsteady Aerodynamic Response of a Linear Cascade of Airfoils in Separated Flow (2004)
    In:  CASI
    Details
    Publication Date: 2019-07-11
    Description: The overall objective of this research program was to investigate methods to modify the leading edge separation region, which could lead to an improvement in aeroelastic stability of advanced airfoil designs. The airfoil section used is representative of current low aspect ratio fan blade tip sections. The experimental potion of this study investigated separated zone boundary layer from removal through suction slots. Suction applied to a cavity in the vicinity of the separation onset point was found to be the most effective location. The computational study looked into the influence of front camber on flutter stability. To assess the influence of the change in airfoil shape on stability the work-per-cycle was evaluated for torsion mode oscillations. It was shown that the front camberline shape can be an important factor for stabilizing the predicted work-per-cycle and reducing the predicted extent of the separation zone. In addition, data analysis procedures are discussed for reducing data acquired in experiments that involve periodic unsteady data. This work was conducted in support of experiments being conducted in the NASA Glenn Research Center Transonic Flutter Cascade. The spectral block averaging method is presented. This method is shown to be able to account for variations in airfoil oscillation frequency that can occur in experiments that force oscillate the airfoils to simulate flutter.
    Keywords: Aircraft Stability and Control
    Format: application/pdf
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    NASA TECHNICAL REPORTS
  • 5
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    Experimental investigation of unsteady flows at large incidence angles in a linear oscillating cascade (1996)
    In:  CASI
    Details
    Publication Date: 2019-07-13
    Description: The aerodynamics of a cascade of airfoils oscillating in torsion about the midchord is investigated experimentally at a large mean incidence angle and, for reference, at a low mean incidence angle. The airfoil section is representative of a modern, low aspect ratio, fan blade tip section. Time-dependent airfoil surface pressure measurements were made for reduced frequencies up to 0.8 for out-of-phase oscillations at Mach numbers up to 0.8 and chordal incidence angles of 0 deg and 10 deg. For the 10 deg chordal incidence angle, a separation bubble formed at the leading edge of the suction surface. The separated flow field was found to have a dramatic effect on the chordwise distribution of the unsteady pressure. In this region, substantial deviations from the attached flow data were found with the deviations becoming less apparent in the aft region of the airfoil for all reduced frequencies. In particular, near the leading edge the separated flow had a strong destabilizing influence while the attached flow had a strong stabilizing influence.
    Keywords: Aircraft Propulsion and Power
    Type: NASA-TM-107283 , NAS 1.15:107283 , AIAA Paper 96-2823 , E-10358 , Joint Propulsion Conference; Jul 01, 1996 - Jul 03, 1996; Lake Buena Vista, FL; United States
    Format: application/pdf
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    NASA TECHNICAL REPORTS
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